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NATIONAL BUREAU OF STANDARDS • Lewis M. Branscomb, Director

Gaithersburg, Maryland

September 14-16, 1970

J. C. Richmond, Editor

NBS, Institute for Basic Standards

National Bureau of Standards

Washington, D.C. 20234

The American Society for Testing and Materials

Institute of Environmental Sciences

The American Institute of Aeronautics and Astronautics

and The National Bureau of Standards

Space Simulation

Proceedings of a Conference held at NBS

Sponsored by

t^O / National Bureau of Standards , Special Publication 336 / Nat. Bur. Stand. (U.S.), Spec. Publ. 336, 984 pages (October 1970) CODEN: XNBSA

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NATIONAL BUREAU OF STANDARDS

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Library of Congress Catalog Card No.: 75-6084^3
FOREWORD

The National Bureau of Standards served as a co-sponsor of

the Fifth Space Simulation Symposium, made its facilities available for the meeting and, with this special volume, is making

the conference proceedings available to participants and to the

public.

Space exploration has posed the most challenging technical

difficulties man has ever faced and successfully overcome. Tlie National Bureau of Standards is proud to have played a significant role in meeting both scientific and technological needs of

the space program, particularly in the field of accurate and

meaningful physical measurements. The standards and new methods

being developed to meet space requirements find their way into

the national measurement system so that everyone ultimately

benefits.

NBS is pleased to cooperate with the American Society for

Testing and Materials, Institute of Environmental Sciences, and

the American Institute of Aeronautics and Astronautics, in

advancing the frontiers of knowledge in the important area of space simulation, and to make these proceedings available to other areas of science and technology.

Lewis M. Branscomb, Director

National Bureau of Standards

iii
PREFACE

The series of Space Simulation Conferences of which this

is the fifth, is sponsored jointly by the American Society for

Testing and Materials, the Institute of Environmental Sciences,

and the American Institute of Aeronautics and Astronautics.

Each society in turn assumes the responsibility for organizing

and conducting the conference and publishing the Proceedings.

The reasons for this cooperative action are threefold: first to provide a form for the presentation of papers before a large audience of listeners who are interested and knowledgeable in this important interdisciplinary area of technology;

second to provide a mechanism by which a majority of the papers

in this area can be published in a single volume annually,

rather than being scattered in a multiplicity of technical and

scientific journals; and finally to reduce the number of annual

meetings , and hence the cost in both time and money to the

sponsoring societies and participants. The Technical Committees of the Sponsoring Societies that participate in organizing

and conducting the conferences include ASTM Committee E-21 on Space Simulation, the lES Committees on Solar Simulation and

Vacuum, and the AIAA Technical Committee on Ground Testing and

Simulation

Three committees are involved in the organization and conduct of each conference, a Permanent Policy Committee, a General Committee, and a Program Committee. The committees for

this conference are listed on the following pages.

The Permanent Policy Committee is made up of two representatives from each of the sponsoring societies. The chair- man is elected by the members of the committee. Each sponsoring society designates its own members on the committee, who

serve for an indefinite term. This committee formulates general policy for the organization of the conferences, and sees that they are conducted in conformity with the rules and regulations of the sponsoring societies, so far as that can be done

in a jointly sponsored conference.

The General Committee is responsible for arranging for the

physical facilities required for the meeting, for conducting

the meeting and for publishing the Proceedings. The General

Chairman and most of the members of this committee, as well as the Chairman of the Program Committee , are from the society

that is responsible for organizing and conducting the meeting.

iv
The Program Committee is responsible for arranging the program,

and is made up of representatives of all three sponsoring

societies

The Fifth Space Simulation Conference was held at the

National Bureau of Standards, Gaithershurg , Maryland, and was co-sponsored by the National Bureau of Standards in addition

to the three societies. The American Society for Testing and

Materials was responsible for organizing and conducting the

Conference, with the assistance of the National Bureau of

Standards. The Proceedings were published by the National

Bureau of Standards

Joseph C. Richmond

Editor

V
COMMITTEES FOR FIFTH SPACE SIMULATION CONFEFvENCE

General Committee

General Chairman - Joseph C. Richmond

National Bureau of Standards

Co-Chairman - Rodney W. Johnson

NASA Headquarters

Program Chairman - Eugene N. Borscr

The Aerospace Corporation

Arrangements Chairman - John Bidwell

ASTM Headquarters

Local Arrangements - Robert T, Cook. National Bureau of Standards

Program Committee

Frank J. Koubek

U.S. Naval Ordnance Lab,

William F. Carroll

Jet Propulsion Laboratory

D. William Conner

NASA Langley Research Ctr.

R. E. (Zeke) King

McDonnell Douglas Astronautics

Charles H. Duncan

NASA Goddard Space Flight Ctr. John H, Boeckel

NASA Goddard Space Flight Ctr. Robert S. Shane

National Materials Advisory Boards

National Research Council - National Academy of Sciences

Joint Permanent Policy Committee

John C. Simons - ASTM John D. Campbell - lES

Rodney W. Johnson - ASTM Thomas Miller - lES

Aleck Bond - AIAA

R. E. Covey - AIAA

vi
TABLE OF CONTENTS

Session I, Contamination

Session Organizer - E. N. Borson > The Aerospace Corporation

Session Chairman - J. Dozier

NASA Marshall Space Flight Center

Session Co-Chairman - Kenneth Taylor

NASA Marshall Space Flight Center

Paper

No. Page

1 Introduction, Arthur Reetz, Jr., National Aeronautics and Space Administration* *

2 Molecular Contamination in Environmental Testing

at Goddard Space Flight Center

J. C. Goldsmith and E. R. Nelson, NASA Goddard

Space Flight Ctr 1

3 Contamination Measurement in Space Environment

Simulation Chambers, C. M. Wolff and M. L. Ritter, Brown & Root-Northrop , Inc., SESL, P.O. Box 34416,

Houston, Tx 25

4 Contamination Studies in a Space Simulated Environment, D. L. Enlow, General Electric Co 51

5 External Spacecraft Contamination Modeling and

Countermeasures , T. Baurer, M. H. Bortner and

I. M. Pikus, General Electric Co. and A. M. Cooper,

Grumman Aerospace Corporation 79

6 Particulate Contaminant Measurement by Quartz

Crystal Microbalance , R. L. Chuan, Atlantic Research

Corporation 105

*Not available for publication.

vii
7 Space Measurements of the Contamination of Surfaces by OGO-6 Outgassing and their Cleaning

by Sputtering and Desorption, D. McKeown and

W. E. Corbin, Jr., Faraday Laboratories 113

Session II, Instrumentation

Session Organizer - J. C. Richmond

National Bureau of Standards

Session Chairman - J. J. Triolo

NASA Goddard Space Flight Center

Session Co-Chairman - D. F. Stevison

Air Force Materials Laboratory

Paper

No. Page

8 Simulation of Radiation from RTG Power Sources,

P. A. Newman, Jr., NASA Goddard Space Flight

Center 129

9 In Situ Vacuum Gage Calibration by the Reference

Transfer Method, L. D. Blado and R. H. Lilienkamp,

McDonnell Aircraft Company 141

10 Methods of Determining Residual Gas Composition

from Residual Gas Analyzer Data, R. H. Lilienkamp,

McDonnell Aircraft Company 151

11 Spectral Absolute Reflectance Measurements of CO2 Frosts in the 0.5 to 12.0 Micron Region,

B. E. Wood,. A. M. Smith, B. A. Seiber, ARO, Inc.,

J. A. Roux, University of Tennessee Space

Institute 165

12 NBS Radiometry Standards - Present and Future,

H. J. Kostkowski, National Bureau of Standards . . 185

13 Apparatus for Laboratory Simulation of Ionospheric

Flow Over On-Board Instruments of an AES,

E. N. Evlanov, V. B. Leonas and S. V. Umansky,

USSR Space Sciences Laboratory

USSR Academy of Sciences 189

viii
14 Thermal Conductivity Measurements of a Candidate

Viking Heat-Shield Material After Sterilization,

and During Exposure to Vacuum, and to a Sim.ulated

Martian Atmosphere, L. R. Greenwood, NASA Langley

Research Center, and Roy M. Fleming,

N. C. State Univ 197

Session III, Special Topics

Session Organizer - D. W. Conner

NASA Langley Research Center

Session Chairman - D. W. Conner

Session Co-Chairman - Emile S. J. Wang

McDonnell Douglas Corporation

Paper

No. Page

15 Method for Rock Property Determination in

Ultrahigh Vacuum, E. R. Podnieks and Peter G. Chamberlain, Twin Cities Mining

Research Center, U. S. Bureau of Mines 209

16 Friction Between Solids and Simulated Lunar

Soils in Ultrahigh Vacuum and Its Significance

for the Design of Lunar Roving Vehicles,

Leslie L. Karafiath, Grumman Aerospace Corp. . . . 225

17 Real Time Simulation of Atmospheric Balloon

Environment, A. Paillous and M. Roussel, ONERA - CERT - Department d' Etudes en Technologie

Spatiale 245

18 Analysis of the Simulation of the Solar Wind,

David Zuccaro, Hughes Research Laboratories .... 261

19 Solar Wind Simulator, H. J. King, Hughes

Research Laboratories 279

20 Laboratory Experiment on Solar Wind Interaction

with Geomagnetic Field, E. M. Dubinin,

G. G. Managadze and I. M. Podgorny, Institute for

Space Research, Academy of Sciences, USSR 297

ix
Session IV, Degradation of Materials by Siinulated

Space Environments

Session Organizer - Charles H. Duncan

NASA Goddard Space Flight Center

Session Chairman - Charles H. Duncan

Session Co-Chairman - Malcolm A. Lillywhite

Martin Marietta Corp.

Denver, Colorado

Paper

No . Page

21 Effect of Photodegradation on Organic Coatings,

P. J. Hearst, U.S. Naval Civil Engineering Lab . . 299

22 Relation Between Dose Rate Effects and Laboratory Simulation of Space Radiation Damage,

A. Roiz^s and R. Schuttler, ONERA - CERT - Dept. d' Etudes et Recherches en Technologic

Spatiale 313

23 Some Problems Concerning the Simulation of

Plastic Films Degradation by Ionizing

Radiations, A. Paillous, ONERA - CERT - Dept.

d' Etudes et Recherches en Technologie Spatiale,

Ph. Fayet, Centre National d 'Etudes Spaciales . . . 323

24 Ultraviolet and Charged Particle Degradation

of Aluminum and Silver Coated FEP Teflon Second

Surface Mirrors, F. G. Cunningham, NASA Goddard

Space Flight Center, B. L. Bean and S. G. Park,

Electromechanical Research, Inc 345

25 The Interaction of Low Energy Electrons with

Polymeric Perf luorinated Ethlyene-Propylene

(FEP), J. B. Schutt, NASA Goddard Space Flight

Center 359

26 The Degradation of Alzak by Short Wavelength

Ultraviolet Radiation, M. J. Donohoe,

R. Mcintosh, Jr., and J. H. Henninger,

NASA Goddard Space Flight Center 383

X
27 The Photo-Response at an Alkaline Paint

Interface with Aluminum, W. A. Wappaus,

NASA Goddard Space Flight Center . . . 393

Session V, Predictive Testing I Session Organizer - Robert S. Shane

National Materials Advisory Board - National Academy of Sciences

Session Chairman - James D. McClelland

The Aerospace Corp.

Session Co-Chairman - Robert S. Shane

Paper

No. Page

28 Plans and Status of NMAB ad hoc Committee on Testing for Prediction of Material Performance

in Structures and Components

Robert S. Shane, National Materials Advisory

Board - National Academy of Sciences 403

29 Monitoring Service Testing by Nondestructive

Testing, Robert W. McClung, Oak Ridge National

Laboratory 405

30 Signature Analysis - Non-Intrusive Techniques for

Incipient Failure Identification Application to Bearings and Gears, B. Weichbrodt and K. A. Smith,

Research & Development Center, General

Electric Co 407

31 Measurement Methods for Microcircuits

W. Murray Bullis, National Bureau of Standards . . 449

32 Life Testing Using Continuous Acceleration,

Ernest Rabinowicz, Massachusetts Institute of Technology, and B. Shiralkar, General Electric

Nuclear Energy Div., San Jose, California 457

xi
Session VI, Predictive Testing II Session Organizer - Robert S. Shane

National Materials Advisory Board - National Academy of Sciences

Session Chairman - James D. McClelland

The Aerospace Corporation

Session Co-Chairman - Robert S. Shane

Paper

No . Page

33 Establishment of an Optimum Duration for

Spacecraft Component Thermal Vacuum Tests,

William L. Harvey, Martin Marietta Corp.,

Denver, Colorado 473

34 Equipment to Study High Velocity Dislocations

in Materials, V. R. Parameswaran and J. Weertman,

Northwestern University ..... 487

35 Some Effects of Equalization on Acceleration

Response and Fatigue Life Expectancy,

Robert M. Mains, Washington University,

St. Louis, Mo 509

36 Charged Particle Track Implications for

Accelerated Testing, Terence M. Flanagan and

Victor A. J. van Lint, Gulf General Atomics, Inc. . 525

37 A Simulation to Validate Future Space Missions,

T. P. Foley, A. W. Metz and A. W. Yonda,

Sudbury Engineering Facility, Raytheon Co 535

38 General Purpose Centralized Automatic Testing

for Improvement of Shipboard System Performance

Frank Zupan, U.S. Navy Electronics Lab. Center . . 549

xii
Session VII, Simulation and Effects - Manned Environments

Session Organizer - R. E. King

McDonnell Douglas Astronautics Co. McDonnell Douglas Corp.

Huntington Beach, Calif.

Session Chairman - R. E. King

Session Co-Chairman - A. H. Bay

McDonnell Aircraft Company

McDonnell Douglas Corp.

St. Louis, Missouri

Paper

No. Page

39 Explosive Oxidiations Initiated by Simulated

Meteoroid Penetration into Spacecraft Atmospheres, William H. Carden, ARO, Inc 569

40 The Dynamic Environment Simulator, A Multienvironmental Man-Rated Centrifuge,

D. B. Rogers and M. McCally, M.D., USAF

Aerospace Medical Research Lab, and

K. L. Cappel, Franklin Institute 587

41 Use of the Ben Franklin Submersible as a Space Station Simulator, M. J. Ferguson,

Grumman Aerospace Corp., and C. B. May,

Advanced Systems Office, NASA 599

42 Development of an Open Circuit Cryogenic Life

Support System for Use in Neutral Buoyancy

Space Simulation, Clifton M. McClure III, and

Charles W. Williams, Hayes International Corp. . . 623

43 Advanced Diving Techniques Applicable to Extended Mission Neutral Buoyancy Space Simulation,

J. B. Maclnnis, M.D., Ocean Systems, Inc.,

and C. M. McClure III, Aquaspace 651

xiii
Session VIII, Facilities

Session Organizer - R. E. King

McDonnell Douglas Astronautics Co. McDonnell Douglas Corp.

Huntington Beach, Calif.

Session Chairman - H. Allgauer

Hughes Aircraft Co. Session Co-Chairman - J. Ratliff

Martin Marietta Corp.

Denver, Colorado

Paper

No. Page

44 A Method to Obtain an Ultraclean Environment,

Maurice S. Cridlin and Joseph W. O'Connor,

NASA Goddard Space Flight Center .... 661

45 A Programmable Dynamic Thermal Vacuum System

For Solar Array Component Testing, John W. Fairbanks, NASA Goddard Space Flight Center,

Marshall B. Eck, H. C. Slack & Associates 669

46 The Cryopumping System for the Space Simulation

Chamber, Hiroshi Nakagawa and Ryuji Tsunoda,

Japan Oxygen Co., Ltd., (Nippon Sanso K.K.) .... 683

47 Simulation of Surface - Atmosphere Interaction

Processes for Mars, K. M. Foreman, Grumman

Aerospace Corporation 699

48 Operation of a Large Thermal Vacuum Chamber at Martian Pressure Levels, T. Buna and J. R. Ratliff,

Martin Marietta Corporation 725

49 A Cryopump for Steady-State Testing of Rocket

Engines Under Altitude Conditions, M. J. Triplett

. and W. E. Riggs, ARO, Inc 749

xiv
Session IX, Ablative Reentry Materials

Session Organizer - Frank J. Koubek

U.S. Naval Ordnance Laboratory

Silver Spring, Md.

Session Chairman - Frank J. Koubek

Session Co-Chairman - Robert L. Bierman

General Electric Co. Paper

No. Page

50 Ablative Heat Shields for Planetary Entries - A Technology Review, G. D. Walberg and

E. M. Sullivan, NASA Langley Research Center . . . 751

51 Ablation Phenomena in Supersonic Laminar and

Turbulent Flows, E. M. Winkler, M. T. Madden,

R. L. Humphrey and J. A. Koenig, U.S. Naval

Ordnance Laboratory 813

52 Prediction of Mechanical Erosion of Charring

Ablators, Lauri H. Hillberg, The Boeing Company . . 837

53 Heating in Cracks on Ablative Heat Shields,

E. M. Winkler, R. L. Humphrey, M. T. Madden

and J. A. Koenig, U.S. Naval Ordnance Laboratory . 853

54 Apparent Operating Limits of Arc Heaters with

Respect to Total Enthalpy and Stagnation Pressure,

R. Richter, Electro-Optical Systems 879

55 Operating Characteristics of the Air Force Flight

Dynamics Laboratory Reentry Nose Tip (RENT)

Facility, John C. Beachler, Air Force Flight

Dynamics Laboratory 893

XV
Session X, Testing in Simulated Space Environments

Session Organizer - E. N. Borson

The Aerospace Corporation

Session Chairman - E. N. Borson

Session Co-Chairman - Fritz Benning

Spectrolab Div. of Textron Corp.

Sylmar, California

Paper

No. Page

56 Investigation of the Characteristics of Fluids

Vented into a Vacuum, R. Steddum, Dupree

Maples and F. M. Donovan, Louisiana State Univ. . . 905

57 Micrometeroroid Impact Simulation System,

R. A. Golub and J. R. Davidson, NASA Langley

Research Center 915

58 A Space Radiation Environment Simulator for the

Evaluation of Solar Cells, P. A. Newman,

J. J. Hirschfield and H. E. Wannemacher,

NASA Goddard Space Flight Center, and

M. Eck, H. G. Slack & Associates 935

59 Properties of Optical Materials in Space,

Fred W. Paul, NASA Goddard Space Flight Center . . 949

Workshop and Panel Discussion

Moderator - Charles H. Duncan

NASA Goddard Space Flight Center

Author Index 961

xvi
SPACE SIMULATION

ABSTRACT: This volume contains all of the papers

presented at the Fifth Space Simulation Conference

held at the National Bureau of Standards September l^-l6 , 1970, that were available for publication. The general scope of the conference "was the

effect of the space environment on materials, com- ponents, structures and man. The range of topics

is too wide to permit a simple classification, but

contamination, ablation, degradation of materials

by the space environment and predictive testing

account for approximately half of the papers presented. Other topics range from purely laboratory

problems such as radiometry standards, calibration

of vacuum gages, cryopumping and operation of

space simulation facilities to gravity simulation,

with neutral buoyancy for zero gravity and a man- rated centrifuge for high gravity, use of a drifting submarine to study the psychological aspects

of long-duration missions in a space station, and

simulation of (l) atmospheric balloon environments,

(2) radiation from nuclear power sources, (3) solar

wind, (U) mi cro-meteoroid bombardment , (5) soil

friction on the moon and (6) the Martian atmosphere. KEY WORDS: ablation, particulate radiation, reentry, space simulation, thermal radiation, vacuum, weightlessness

INTRODUCTION

The purpose of this conference, as stated in the Call for

Papers, was to provide a forum for the presentation of new information on the technology of space simulation. This technology encompasses all aspects of simulating the environment

of space and determining its effect on materials, components

and spacecraft. Specific aspects of the space environment

that are of interest include vacuum, solar and planetary radiation, particle radiation, meteoroids, weightlessness, magnetic

fields, temperature extremes, planetary atmospheres, and reentry. All of these topics are covered in the papers presented on the following pages. In addition, two other topics

are included in this conference that have not been adequately

covered in previous conferences; Contamination and Predictive

Testing.

The 59 papers accepted for presentation were separated

into 10 sessions and are grouped by sessions in these Proceedings . All papers were subjected to careful technical

xvii
review prior to final acceptance and comments of the reviewers were sent to the authors prior to preparation of the final

draft in form suitable for photo-offset reproduction, hut the

final drafts have not been reviewed. For this reason each

author must assume full responsibility for the content of his

paper

xviii
Paper No. 2

MOLECULAR CONTAMINATION IN ENVIRONMENTAL TESTING AT

GODDARD SPACE FLIGHT CENTER

J. C. Goldsmith^ and E. R. Nelson^

REFERENCE: Goldsmith, J. C. and Nelson, E. R. , "Molecular Contamination in Environmental Testing at Goddard Space Flight Center,"

ASTM/IES/AIAA Space Simulation Conference, 14-16 September 1970. ABSTRACT: A discussion of the various techniques used to detect

and analyze materials outgassing from a spacecraft during the environmental test is presented. In addition, a number of compounds that have been detected and their parent outgassing sources are listed. The two major methods for reducing the threat of contamination (i.e.

proper choice of spacecraft materials and proper operation of the test

facility) are also discussed.

KEYWORDS: contamination, outgassing, polymeric materials, residual gas analysis, vacuum, test facility, chemical analysis, organic

residues

This paper describes the techniques presently being used at Goddard Space Flight Center (GSFC) to detect and minimize molecular

contamination of spacecraft or spacecraft subsystems and experiments during the environmental test. The methods described are an outgrowth of work done to determine the test readiness of oil -diffusion-pumped test chambers for spacecraft having UV reflective optics.

Senior Engineer—Sperry Rand Corp. Space Support Division, Goddard Space Flight Center, Greenbelt, Md.

^Polymer Chemist—National Aeronautics and Space Administration,

Materials Research and Development Branch, Goddard Space Flight

Center, Greenbelt, Md.

1
DEFINITION

Molecular contamination can be defined simply as a deposition of molecules upon a surface that would affect some physical property of

that surface (i.e., reflectance, transmittance, absorptance, or surface resistivity) . These contaminants are chiefly composed of water

vapor and organic molecules of medium to high molecular weight (50

to 800 atomic mass units) and low vapor pressure. The low-molecular-weight organic compounds and residual gases are too volatile to adhere to any surface in a vacuum unless the surface temperature is in the order of -60 C or lower. Hence, these compounds do not ordinarily present a problem since spacecraft surface temperatures are

generally above this level. However, recently there have been some spacecraft experiments

flown that utilize detectors that are cooled to 0 C or lower. In this temperature range, some gases, volatile, low -molecular -weight

compounds, and water vapor will condense, thus obscuring the detector's view.

EFFECTS OF CONTAMINANTS

Molecular contamination can affect spacecraft performance in four major ways.

1. Loss of Optical Reflectance

A film of silicone oil approximately seven molecular layers (0.4

|jg/cm^) thick will cause a loss of approximately 8 percent in the reflectance measured at 1216A(1).^ Thus, the output of an optical experiment operating in the ultraviolet region in which the beam is reflected from several consecutive reflective surfaces can be seriously reduced.

2. Loss of Optical Transmittance

If the mass accretion upon an optically transmissive surface is sufficient, it can reduce optical and near infrared transmittance (0.4

to 18[im region). This may result in the degradation of the output of solar -cell arrays, the loss of sensitivity of optical experiments or sensors, or the loss of vision, as in the case of the Gemini 11 where

severe clouding of the spacecraft window occurred (2)

The numbers in parentheses refer to the list of references appended

to this paper.

2
3. Degradation of Thermal -Control Coatings

The molecular contaminants that condense on thermal-control

surfaces can alter the solar absorptance-to-emittance ratio. At this time, however, the relationships between the type and amount of molecular contaminant and the extent of degradation for polymeric materials are not clearly understood.

4. Malfunction or Failure of Electronic Systems

Outgassed contaminants condensing on a circuit board assembly,

on open contacts, or in waveguide subassemblies can alter effective circuit impedances to the point that malfunction or failure of an ex- periment or subsystem may result. In addition, molecules outgassing

in a confined enclosure may maintain pressure in the corona region.

Consequently, when sensitive experiments, transmitters, and highvoltage power supplies are commanded on, they can fail in orbit due

to high-voltage corona breakdown if these circuits are turned on before the volatized organics and residual gases have had a chance to diffuse into space.

SOURCES OF MOLECULAR CONTAMINATION

In any specific test there are two basic sources of contaminants:

the test facility and the spacecraft.

1. Possible Test Facility Contamination Sources

• Compounds adsorbed on chamber wall from previous tests.

• Cracked oils from the diffusion or mechanical pumping

system.

• Paint volatiles or cutting oils used in spacecraft mounting

fixtures.

• Outgassed compounds from cabling, potting compounds,

adhesives, flexible heaters, etc. , used to support the

test.

• Water or other compounds adsorbed from the ambient air

environment.

• Cleaning solvents.

2. Possible Spacecraft Contamination Sources

• Outgassed compounds from cabling, potting compounds,

sealants, circuit boards, adhesives, etc., in the electronic subsystems.

3
• Paint volatile s from improperly cured thermal coatings.

• Water or other compounds adsorbed from the ambient air

environment.

• Cleaning solvents. METHODS OF MEASUREMENT

The following complementary analytical techniques are available

for the quantitative and qualitative analysis of contaminants outgassed

during environmental test:

• Residual gas analysis (RGA).

• Infrared spectroscopy (IR).

• Gas chromatography (GC).

• Gas chromatography-mass spectrometry (GC-MS).

• Quartz -crystal microbalance gravimetry (QCMB).

• Ultraviolet reflectance. The use of all of these techniques may be necessary for the environmental test of a large spacecraft.

1. Residual Gas Analysis

A residual gas analyzer (Fig. 1). is used to quantitatively and

qualitatively identify the molecules present in the spacecraft test en- vironment. This analyzer is a mass spectrometer of relatively low

resolution and mass range that ionizes molecules present in the en- vironment and separates them by atomic mass. The data presented

by this instrument is in the form of a plot of the molecular weight of the molecules detected versus their relative abundance or partial

pressure. A typical RGA spectrum is shown in Fig. 2. A cursory inspection of the spectrum is all that is required to determine the presence and amounts of residual gases such as air,

nitrogen, and water because of their relatively simple molecular

formulas

However, to determine the presence of the higher molecular

weight contaminants, those molecules that originate from the spacecraft materials, one must have prior knowledge of the characteristic mass spectra or "fingerprint" spectra of all compounds likely to be

found in the test environment. The problem is further complicated by

the fact that the mass spectrum presentation represents a summation

of the mass spectra of each individual compound in the test environment at that particular instant. The resolution of this mass spectrum

4
into its individual compounds requires a sophisticated computer program. Such a program has recently been developed for this purpose

(3). Fingerprint mass spectra of known compounds are fitted to the unknown mass spectrum. The computer then determines the combination of fingerprint spectra that gives the best fit to the unknown

spectrum and prints out the results. The use of this program thus far has been limited only by the sensitivity of our analyzers and the limited amount of available fingerprint spectra of commercial compounds.

2 . Infrared Spectroscopy

A cold plate (Fig. 3) located in the test volume is operated at 77 K for the duration of the test, during which it collects a fraction of

all condensible contaminants in the test environment. At the conclusion of the test, the cold plate is washed with isopropanol solvent. The mixture of solvent and contaminant is then transferred to the analytical chemical laboratory where the solvent is evaporated and the residue (i.e. , contaminants) is analyzed, chiefly by infrared spectroscopy. In this technique of analysis, the infrared spectra of the contaminant is taken in the range from 2.5 to 25 [im. Figure 4 shows a

typical infrared spectrum obtained from the OAO thermal -vacuum test. The absorption pattern in this band is representative of the chemical

bonds between atoms of these molecules and is thus a fingerprint of the molecules.

However, as with RGA, the resulting infrared absorption spectrum represents a summation of all outgassed compounds that have

condensed on the cold plate. Thus, the mixture spectrum is capable

of giving general information on the functional groups present: ali- phatic, aromatic, hydrocarbon, silicone, ester, amide, etc. For a more complete identification of the individual compounds in the mixture, the mixture must either be separated for individual identifi- cation or a computer solution, similar to the one used for mass spectral identifications, must be applied to the mixture spectrum.

3. Gas Chromatography

The gas chromatograph is an analytical instrument that separates

a mixture of compounds into pure compounds. The sample obtained

from the residue from the cold plate is vaporized and bled into a stream of helium that then passes through a column of absorbing material. The column packing absorbs the less volatile compounds more

readily than it absorbs the more volatile compounds. Thus, the material will emerge from the column as a series of time -spaced peaks

5
in the helium stream with each peak representing a unique compound.

These peaks are usually detected by either flame ionization or thermal

conductivity detectors at the column exit or both. The results are displayed as a series of roughly gaussian peaks

plotted versus time on a strip-chart recorder. The data are taken as time of peak height (column traverse time) versus area under the curve, which is proportional to the amount of the compound. Though

this technique does give a very good fingerprint, it, unfortunately,

does not uniquely determine a given compound and may and usually

does vary somewhat from run to run. Hence, the gas chromatograph

cannot be used for identification without running reference compounds

or trapping the eluted compound for subsequent infrared or mass

spectrometric analysis.

4. Gas Chromatography -Mass Spectrometry

In addition to the infrared spectrometer and gas chromatograph,

the analytical laboratory utilizes a combination gas chromatograph- mass spectrometer for analysis of the contaminant residue. This

combination overcomes the difficulties of the use of the gas chromatograph alone. The combination enables one to separate the mixture

into its pure molecular constituents and to obtain a mass spectrum

for each constituent from 10 to 1000 atomic mass units. As a consequence of this capability, thirty or more mass spectra

may be generated from an average cold-plate sample containing a mixture of up to thirty compounds. For expediting the resolution of these spectra, automatic data reduction is a necessity. Figure 5 shows the gas chromatograph-mass spectrometer data acquisition

unit in use at GSFC.

5. Quartz -Crystal Microbalance Gravimetry

The QCMB measures gravimetrically the mass accretion of contaminant over the surface of an oscillating crystal. The change in the frequency of oscillation of the crystal is directly related to the mass

accreted upon it. With proper precalibration and temperature compensation, this instrument^ is capable of measuring as little as

01 fig/cm^.

^Hanyok, J., "Crystal Microbalance Vacuum Contamination Monitor,"

NASA X-Document X-324-69-404, National Aeronautics and Space

Administration, Goddard Space Flight Center, Greenbelt, Md.

Sept. 1969.

6
If the QCMB is placed at or near a critical surface, such as an

optical or thermal-control surface, and the QCMB crystal is configured so as to have the same "view" relationship and temperature as

the surface in question, then the mass accretion on the QCMB will approximate the mass accretion on that surface. Figure 6 shows the QCMB system.

6. Ultraviolet Reflectance

As mentioned previously, small amounts of condensed contaminants will cause a significant loss of reflectance in the ultraviolet region. Hence, with proper precalibration, a mirror sample (Fig. 7) can be used to measure directly contaminant levels on a critical surface. As in the case of the QCMB, the mirror must be configured so

that it has the same "view" and temperature relationship of the surface in question. Pretest and post-test reflectivity measurements

are made at 12 16 A with a vacuum monochromator . Most organic

chemicals absorb ultraviolet radiation at this wavelength but in varying amounts, depending on their chemical structure. Consequently, some prior knowledge of the type of contaminant is required for accurate measurement.

A change of reflectivity at 12 16 A is indicative of contaminant.

The exact level of contaminant deposited is, however, somewhat in doubt

THE RELATIONSHIP BETWEEN CONTAMINANT MEASUREMENTS

AND ACTUAL SPACECRAFT CONTAMINATION

Although each of the measurement methods previously described

will give some quantitative picture of the contaminant levels within the

test volume, the actual quantities of material outgassed or desorbed

from the spacecraft can only be inferred. This is due to the fact that,

in the average test, the majority of the outgassed compounds are pumped out and only a small percentage will return to the spacecraft

and condense as a surface contaminant. (Figure 8 shows schematically the relationship among these measurement methods.)

As shown by R. Chaun, ^ the fraction of outgassed material actually condensed on the spacecraft is a function of the chamber wall

capture coefficient, the spacecraft surface capture coefficient, and

the ratio of internal chamber area to total spacecraft surface area.

^Notes from "Vacuum Technology," a course at the University of California at Los Angeles (Extension), Feb. 14-18, 1966.

7
For a precise measurement of the amount of contaminant condensed

on a spacecraft surface (aside from physically wiping the surface

which has been done in a few cases, i.e. , OAO 2), it would be necessary to precalibrate the particular spacecraft -chamber system.

Modeling techniques can be used to determine the relationship of the material condensed on the spacecraft to the quantitative amounts

measured by the techniques described previously.

In solar-vacuum testing, the chamber shrouds are at 77 K or lower; therefore, a molecule that desorbs from the outgassing source

will be quickly cryopumped on the cold shroud. Consequently, the analysis of the cold finger sample may show only a small percentage

of the outgassed compound, and the RGA or reflective mirror sample

may indicate no trace of contaminant molecules, depending on their

respective "views" and capture coefficients relative to the outgassing source. In fact it has been found at GSFC that the cold finger will often pick up traces of compounds not detected by the RGA. This is due,

possibly, to the integrating effect of the cold plate as well as the limited mass range of the RGA.

In the thermal vacuum test, however, the outgassing molecule

will collide several times with the warm wall or spacecraft before

coming to rest, thus establishing a relatively uniform molecular flux throughout the test environment. Consequently, all contaminant sensors will "view" the outgassing compound(s). However these compound(s) may or may not condense, depending on their relative surface

capture coefficients.

CLASSES OF COMPOUNDS OBSERVED

A qualitative identification of the major compounds observed in the test environment is shown in Table 2. These compounds, collected chiefly from the cold plate, are grouped into three classes

according to their percentage of the total condensed sample.

In addition to the compounds listed in Table 2, traces of many

more compounds associated with plastics have been observed. Compounds such as 2, 6-di-t-butyl-p-cresol, biphenyl, and butyl ethyl

ether of ethylene glycol, have been observed in quantity but infrequently, and so far, it has not been possible to determine their source. Backstreaming silicones from the chamber pumping system are conspicuous by their absence. A positive gas chromatographic technique has been developed that is able to detect Ip-g or less of tetra- methyltetraphenyl-trisiloxane compound (DC 704). With this technique, backstreaming silicones have been observed in only six tests, with over six hundred samples thus far analyzed.

8
In general, outgassing levels of polymeric compoimds, will be

seriously affected by the following factors: their "cure" cycle (i.e.

the correct time -temperature relationship following their preparation),

the amounts of plasticizer and antioxidents added to the parent compound to modify its basic physical properties to fit the application,

and the operating temperature of the compound.

METHODS OF REDUCING CONTAMINATION

1. Choice of Materials

The ideal method for reducing contaminant levels is to reduce the outgassing rate at the source. This can best be accomplished by the proper choice of materials in the design phase.

The criterion that is used at GSFC for nonmetallic spacecraft

materials is based on the work of Muraca and Whittick (4) at the Stanford Research Institute for the Jet Propulsion Laboratory.

These criteria limit the use of materials to those with less than 1- percent total weight loss (TWL) and 0.1 -percent vacuum condensible

material (VCM) . The VCM value tends to be far more important than

the TWL value because the difference between VCM and TWL is accounted for by the light gases and water. Normally, if a material has

a high TWL and a low VCM, it still can be used on a spacecraft if the material is free to outgas into space. At present, the Material R&D Branch at GSFC is continuing the certification of additional nonmetallic materials for inhouse application. This is being done with a highly automated material outgassing

facility developed at GSFC and based on the original design of Muraca

and Whittick.^

Where a specific requirement can only be met by a material with

a high outgassing rate, it is recommended that the material or subassembly be preprocessed (i.e. , exposed to a high -temperature and vac- uum environment before integration into the spacecraft) . Thus , the material whose initial gas load is composed largely of gases, water,

and lower-molecular weight fractions is "cleaned out." At the present

state of the art, it is not possible to prepare a spacecraft free of all sources of molecular contaminants, but the contaminant hazard can be

significantly reduced by the use of the correct nonmetallic materials.

^Fisher, A. and Marmelstein, B., "Micro -Volatile Condensable Material System for Polymer Outgassing Studies," NASA X-Document X735-69-47, National Aeronautics and Space Administration, Goddard

Space Flight Center, Greenbelt, Md. , Oct 1969.

9
2. Operation of The Vacuum Chamber

Several practical steps may be taken to minimize the contamination threat to a spacecraft during environmental test:

• Pretest bake -out . This will remove contaminants condensed

on the chamber wall from previous tests. This can best

be accomplished by a high-temperature long-term bakeout (100 C for 48 h is a recommended minimum level).

• Maximum pump-speed to test -volume ratio . Cryopanels and

high-capacity pumping system will maximize chamber

capture coefficients, thus effecting the rapid removal of the contaminant from the test volume.

• Minimum spacecraft-chamber temperature differential . The

temperature differential between the chamber wall and

the spacecraft during cold-to-hot temperature transition

should be held to a minimum, preferably 10 C or less. This will extend transition time; however, it will de- crease chamber-to-spacecraft contamination.

• Backfill chamber . In order to hold contaminants that have

been trapped on the cryopanel, it is desirable to backfill the test volume to approximately 0. 1 torr with dry GN2

when concluding a test in which large cryopanels are

used. This will reduce the molecular mean free path in the test volume to less than 2.5 mm while still maintaining the environment below the region of convective heat transfer. Thus, the probability that a trapped contaminant released by the cryopanel during the warm-up phase

will reach the spacecraft is considerably reduced.

• Protective cryopanels A relatively small cryopanel interposed between a sensitive experiment and the chamber

cryopanel will occlude the experiment's "view" of the larger chamber cryopanel, thus reducing the contaminant hazard due to outgassing from the chamber wall. In addition, larger cryopanel surfaces that are kept cold

during chamber warmup will act as a molecular sink

for molecules released from the chamber wall.

• Sweep purging Backstreaming of a diffusion pumping system

is at a maximum when the diffusion pump is operating in a transition mode (i.e. , pumpdown or repressurization)

A small leak or purge of nitrogen bled into the chamber

during this transition period will stabilize molecular flow

outward (i.e. , from the test volume to the atmosphere)

(5), thus minimizing the risk of backstreaming.

10
CASE HISTORY-AN EXPLORER SPACECRAFT

A contaminant case history is given in the succeeding paragraphs,

for one of the Explorer series spacecraft.

1. Spacecraft Description

The spacecraft had a mass of approximately 140 kg and carried

six experiments. The nonmetallic materials used in the fabrication of the spacecraft were

• Polyurethane foam potting compounds.

• Epoxy and silicone potting compounds.

• Silicone and ester lubricants.

• Silicone adhesives.

• Epoxy adhesives.

2. Test Procedure

The spacecraft was subjected to ten-day thermal -vacuum test fol- lowed by a six -day solar-vacuum test. A three -day thermal -vacuum

retest was performed finally to assure that malfunctions that had oc- curred during the first test had been corrected.

3. Test Instrumentation

• Residual gas analyzers

Consolidated Electrodynamics Corp. 21-613

General Electric Monopole

• Cold plate sample analysis

Perkin Elmer 621 Spectrophotometer

Hewlett Packard 5750 Gas Chromatograph

Perkin Elmer 270 Gas Chromatograph - Mass Spectrometer

4. Analysis of Outgassed Compounds

A complex mixture of almost 2.5 g of outgassing products was

found on the condenser plate following the first thermal-vacuum test. This sample was analyzed by infrared, gas chromatographic, and mass spectrometric techniques. Figure 9 shows an RGA spectrum

taken during the second retest phase. Fig. 10 shows the IR spectrum

of the cold plate sample mixture, and Fig. 11 shows the gas chromatogram of the cold plate sample mixture.

The mixture of at least fifty components, twenty of them major in quantity, was found to have the following composition: a complex mix11
ture of aliphatic or alicylic hydrocarbons or both, 50 to 60 percent; a

series of phenylmethyl siloxanes, 5 to 10 percent; six members of a homologous series of dimethyl silicones, 15 to 20 percent; and pure

bis(2 -ethyl hexyl) phthalate, 20 percent. The hydrocarbons could come

from several sources such as lubricating oil (i.e. , a petroleum -based

lubricant) or perhaps from the outgassing of some O -rings, which

quite often use plasticizer with white oils of this type. The dimethylsilicones and phenylmethyl silicones are most likely due to outgassing

of RTV and other silicone resins. The bis (2-ethyl hexyl) phthalate is one of the most common plasticizers used in the United States and thus could come from outgassing of various compounds. It should be noted

that the use of such compounds is discouraged at GSFC.

This mixture did not contain any of the oil from either the chamber diffusion pumps (Dow Corning 704) or from the mechanical pumps

(Sunvis 931). These results were obtained by running Dow Corning

704 and Sunvis 931 through a gas chromatograph under the same conditions as those of the outgassing mixture. The fingerprint obtained

did not match any peaks in the outgassing mixture. During these thermal vacuum tests, a sun gun (quartz -iodine photoflood lamp) was used

to excite the spacecraft solar -aspect sensor. This sun gun was suspected of being a source of the hydrocarbon contaminants seen in the analysis because of its high operating temperature. Consequently, a

separate outgassing test was performed on the sun gun. Gas chromatographic and infrared spectra of the condensed cold plate residue

failed to show any similarity with spectra from the thermal -vacuum

test. Hence, the sun gun was ruled out as a possible contributor of hydrocarbons.

For the subsequent solar-vacuum test and thermal-vacuum retest, the condensor plate samples taken after these tests showed al- most the same infrared patterns and gas -chromatographic patterns as had been obtained from the thermal-vacuum test, only in much less

quantity. Results of the RGA scans were inconclusive during both the

initial thermal -vacuum and initial solar-vacuum tests because of a

loss of resolution during the thermal -vacuum test and a loss of sensitivity in the solar-vacuum test. In the solar-vacuum test, no peak

above atomic mass 57 could be detected. SUMMARY AND CONCLUSIONS

With the measurement techniques described, it has been possible

to partially identify and roughly approximate the contaminants out- gassed in a specific test. However, a great deal more analytical and

experimental work must be done before exact quantitative and qualita12
tive contaminant data can be determined. Two general observations

may be made:

(1) It has been found that, in most of the tests, the test item is the prime source of contaminants.

(2) Contamination traceable to the chamber diffusion-pump system occurred in a very small percentage of the tests

(0 . 1 percent)

Future spacecraft missions involving extended orbital life such as

the Grand Tour deep -space probe and the next generation of communication satellites and orbiting observatories will require a functioning

lifetime of three years and longer. This imposes a severe constraint on the spacecraft designer in his choice of materials.

Longer orbital lifetimes also will mean an extended environmental-testing program; consequently, it will be necessary that the environmental test engineer be assured that his test facility is clean and

that his methods of operation minimize the contaminant hazard.

13
REFERENCES

1. Shapiro, H., "Monomolecular Contamination of Optical Surfaces,"

NASA Technical Note D-4612, National Aeronautics and Space

Administration, Washington, D.C., June 1968.

2. Blome, J. C. , and Upton, B. E. , in "Effects of the Space Environment on Materials," Society of Aerospace Material and Process

Engineers National Symposium and Exhibit, 11th Symposium,

Science of Advanced Materials and Process Engineering Proceedings, Vol. 11, p. 217, 19YQAE.

3. Babst, R. W. , "Computerized Analysis of RGA Scans," Final Report SP 923-0054, Sperry Rand Corporation, April 1970,

4. Muraca, R. F., and Whittick, J. S., "Polymers for Spacecraft

Application Final Report," SRI Project ASD-5046, JPL Contract No. 950745, Menlo Park, California, Sept. 15, 1967.

5. "Technology and Space Simulation," NASA Publication SP-105,

National Aeronautics and Space Administration, Washington,

D.C., 1966.

14
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Paper No. 3

CONTAMINATION MEASUREMENT IN SPACE ENVIRONMENT SIMULATION

CHAMBERS

Calvin M. Wolff and Melvin L. Ritter"^

ABSTRACT

Contamination Tneas-urement is classified according to the requirements; i.e., the identity, soircces , and effects of contamination "by active and passive sampling. The chemical natiire of

the most common contaminants foiind in space environment simulation is discussed. Qualitative and quantitative techjiiques for

contaminants analyses are described and evaluated. These techniques include mass spectrometry, quartz crystal oscillator

mass monitors, infrared spectrometry, and gas-liquid chromatography. Also, the uses of optical surfaces and thermal control

coatings as contaminants effects monitors are described.

KEY WORDS

Space environment simulation chambers, mass spectrometry, infrared spectrometry, gas -liquid chromatography, contamination collection units, quartz crystal oscillator mass monitors, reflectometry, radiometry, thermal control coatings.

INTRODUCTION

The increasing sophistication and flight duration of space vehicles and related hardware is imposing more stringent cleanliness requirements of space environment testing facilities. The

need of controlling contamination and measuring its effects is thus "becoming of great importance. Measurement of contamination may yield the following information:

1. The source and nature of contamination

2. The identity and quantity of contamination

3. The effects of contamination on test articles.

*Brown & Root-Northrop, Inc., SESL, P.O. Box 34416, Houston, Tx.

25
THE SOimCE Am NATUEE OF CONTAMINATION

Contamination in space environment chamlDers may be considered

as twofold: particulate and molecular. Particulate contamination, such as lint and dust, is, in most instances, of less concern than molecular contamination, because the particulates

can be easily and completely removed. Particulate removal may

be made by forced air or other gas, by bombardment with soft

cKrparticles (anti-static technique) , or by brushing or wiping

if contact is permissible. Also, particulate contamination affects only extremely small fractions of test article surface

area; however, particulate contamination may be dangerous when it

is not found and removed, for it degrades delicate electronic

contacts, miniature mechanical systems, and blocks slits of

spectrometers. Particulate contamination in space environment

testing can occur at all times when tne test article is at ambient conditions, during the initial pumpdown phase, and during the

final (high pressure) repressurization phase. Particulate con- tamination is very improbable under high vacuum conditions , being

that there is no external source of particles and no matrix for

the mobility of particles, save for the gravitational field.

The effects of particulate contamination on thermal-vacuum test data are negligible because the fraction of affected area involved in heat transfer is minute. Likewise, the effects of particles on facility heat transfer surfaces , such as cryopanelling

and solar simulator and radiometer optics, are negligible. Several references are available describing detection and removal

of particulate contamination (references 1 thru 8).

Molecular contamination is a serious problem in thermal-vacuum

testing. Molecular contamination may be defined as that which

migrates and deposits in the test volume as individual molecules,

rather than the relatively massive molecular congomerates and

polymers that we refer to as particulates. Molecular contamination may emanate by evaporation, sublimation, chemical decomposition, or desorption. These processes are described in references 9 thru 21

Because molecules migrate individually and randomly, they also

will deposit in a random manner, that is, tend to cover an ex- posed surface uniformly. Since molecular masses are so low, surfaces are quickly coated with monomolecular layers of contaminants , even though the partial pressures of contaminants may be

quite small. For example, one milligram of molecular contaminant will deposit, roughly, one monolayer over a surface area of ten square feet. The time of a contaminant remaining on a sur26
face after it has deposited is governed by desorption rate parameters. The contamination level on a surface is dependent on

the rate of deposition and desorption or evaporation. The con- tamination level will be high for any combination of a high and

continuous deposition rate, a very high tenacity of the contaminants on the surface, and a very low surface temperature. The

situation becomes worse when the ultraviolet radiation of solar

simulation breaks intramolecular bonds of contaminants on test article surfaces and which in turn may bond chemically with the

test article surface and each other. Since the chemical adsorption (chemisorption) energies are much higher than other types,

the desorption rate of contaminants on the test article would be

greatly reduced, causing essentially permanent contamination.

One important facet of desorption is that when a surface is ex- posed to a gas on the order of one atmosphere pressure, as in

"ambient" conditions, desorbed molecules are reflected back

onto the contaminated surface by the gas molecules to an extent

that elevating the surface's temperature is not sufficient for

decontamination; that is, elevated temperatures and vacuum are

required. Further, the adsorption energies of many molecules on

a foreign surface are generally much greater than the heat of

vaporization or sublimation of that molecule, such that the

final monolayer of contaminant is very difficult to remove.

This monolayer must be removed either by vacuum "bake-out" or by

chemical decomposition (such as oxidation) of the contaminant

into gaseous and particulate material.

The "permanent" gases (oxygen, nitrogen, hydrogen, helium, methane, carbon monoxide, argon, and neon) will always be present in vacuum chambers; however, they are of little concern as contaminants because even if they should adsorb, their effects on thermal

and optical properties of test surfaces are negligible and quantities greater than a monolayer will desorb above cryogenic temperatures. Carbon dioxide and moisture (water vapor) will always

be present, since they will be adsorbed in large quantities on

the chamber and test surfaces during ambient conditions and migrate extensively during pumpdown and repressurization phases.

If sufficient time is allowed during pumpdown to degas and pump

away most of the carbon dioxide and water, and if leakage is controlled, these contaminants should be of no consequence.

However, if these gases are allowed to migrate from warm surfaces

or enter the chamber when the test article is cold, then con- tamination may build up (sheets of ice and frozen CO2) to a point as to seriously affect heat transfer at thermal control

surfaces. These materials have no absorption in the near ultraviolet and visible regions of the spectrum, so that portion of

solar absorptance would not be affected; however, they have con27
siderable infrared absorption and can affect the emittance of the

coatings. These contaminants could be removed by solar simulation without permanently affecting thermal or optical properties.

In fact, the saving grace of carbon dioxide and moisture contamination is that they are reversible; i.e., they are removed, ex- cept for a few monolayers., when the test article is returned to ambient temperatures.

The vapor pressures of ionic and of most metallic compounds (I4) are sufficiently low as not to be of general concern as contaminants. There are exceptions, of course, such as mercury and cadmium, but these materials should be restricted from use in space

environment facilities. The standard inorganic materials recom- mended for vacuum usage, such as stainless steel, copper, aluminum, gold, quartz, pyrex glass and porcelain, are so stable and

of such low vapor pressure as not to be of concern as contaminants in space environment simulation testing.

Organic materials, however, are the most troublesome contaminators. Organic materials must often be used in space environment

chambers where non-brittle adhesives and flexible electrical and

thermal insulators are required, and are also used as vacuum

sealants, lubricants, insulative structural materials, hydraulic

oils, diffusion and mechanical pump oils, and thermal control

coatings. To minimize contamination, polymer materials are preferred. Since polymer molecules (many relatively small molecules linked together) are very heavy (molecular weights in the

range of several thousand or greater), their mobilities are very

small, and the great number of chemical bonds present furnish

many modes in which to dissipate and distribute thermal energy

other than in evaporation. When non-polymers must be used, they are of high molecular weight (several hundred or more), to minimize evaporation. The contamination source of polymer and nonpolymer organic compounds is not the very heavy molecules , but

the lighter impurities initially present (16). Most organic materials used in vacuum work show continually decreasing vapor

pressures as the time of vacuum exposure increases. The materials outgassed from polymers are usually the monomers, i.e.,

the "building block" molecules or link of the polymer chain and

also catalyst compounds (19, 20, 21 ). In non-polymers, the out- gassed materials are usually lower boiling fractions or impurities in the oil, which may be initially present, and may also be

generated by very slow "cracking," or decomposition, of the

larger molecules into smaller more volatile molecules. These

materials will eventually outgas themselves to a state at which

they are quite involatile, and the only outgassing would be

from very slow decomposition. Common classes of organic materials are briefly discussed below.

28.
Heavy Hydrocar"bons , Oils, and Esters

Commercial heavy hydrocarbons in vacuum use are high boiling petroleum fractions. Their infrared (36) and mass spectra are es- sentially identical, as are their chemical behavior, including

their effects as contaminants. A large number of hydraulic oils,

mineral oil, motor oils, lubricating oils, and vacuum oils,

greases, and waxes (Apiezon) are of this nature. Many natural

organic oils are not hydrocarbons but are fatty acids or their

ester derivatives. The fatty acids have a ten-to-thirty carbon

atom chain which terminates in carboxyl (O = C-OH) group (22).

The existence of the polar (-G=0) (carbonyl) group and of possible double bonds in the carbon chain cause the compound to remain in the liquid state despite the high molecular weight. Exters of these oils (hydrocarbon chain terminating in -C00(CH2) nCH^ group, n = 0, 1, 2, 3) have even lower melting points, that

is, have more of a liquid nature. When the size of the carbon

chain becomes large (ten or more), the chemical and physical nature of hydrocarbons, acids, and oils become similar. These

types of compounds often fluoresce, making them detectable upon

visual inspection with black (ultraviolet) light, and are easily

soluble in organic solvents and detergents. They may form stubborn tars when exposed to solar simulation, however.

Halocarb ons

Halocarbons are organic compounds containing carbon and at least

one halogen atom. Halogens are fluorine (f) , chlorine (Cl),

bromine (Br), and iodine (l).

The most important halocarb on in vacuum use is Teflon, a perfluoroethylene polymer whose repeating monomer is (-CF2-CF2)^;

Teflon is the least troublesome of the organi cs , as the outgassings, nearly 1 0(^ monomer, C2F. , are stable and are reluctant to adsorb even on very cold surfaces (25). Halocarbons containing

fluorine and some other halogen atom, such as Kel-F and freons

are of more concern, because their outgassings are decomposed by

ultraviolet radiation from solar simulation, the decomposition

products being very reactive with and detrimental to first surface aluminum, SiO, and Si02 coated mirrors of solar simulators

(24).

Silicones

Silicones are silicon containing organic compounds usually con- taining several continuous silicon-oxygen-silicon linkages (sil29
oxanes). The silicone oils, greases, and rulDbers -used in vac-Qiim

work are commonly polymers of dimethyl siloxane, -^0-Si (CHs)^^.

These compounds have common and very characteristic mass and infrared spectra (27,31 )• T^iey are difficult to dissolve with any

solvent and are very stubhornly adsorhed on foreign surfaces.

They are known to "creep" over very large continuous surfaces,

forming a very fast monomolecular layer of silicone in their

wake. When light silicone polymers and monomers are outgassed

and deposited on foreign material, they may easily repolymerize

especially with the aid of ultraviolet light from solar silulation.

The DC-700 series silicone oils are used frequently as diffusion

pump oils. They are non-polymers, heing phenylmethyl trisiloxanes , that is, R^Si-0-SiR2-0-SiR^ , where the R's are methyl or phenyl ("benzene) groups in varying distrihutions ah out the molecule . ,

Epoxies

Epoxy polymers vary greatly in chemical structure, but they all were created from monomers containing the epoxide group,

_ C - C -, which decomposes and cross-links to form the polymer.

Many of the high omittance thermal control coatings used in space

technology are epoxy paints, whose monomer is an ester containing the epoxy group. Heavy esters, alkyd resins, drying oils,

and alcoholic ethers ( cellosolves ) are used as matrices for these

paints and are the major outgassed materials (34). The esters and epoxy monomers are quite soluble in most organic solvents,

but are also subject to tar formation and polymerization by ultraviolet light (35)

THE SAMPLING OF CONTAMINANTS

Contaminants may be analyzed according to their quantity, identity, directionality, and effects. They can be analyzed on an

in situ, real time basis (active), or samples can be removed

from the test site after testing and analyzed in the laboratory

(passive). An important factor in either active or passive

sampling is that the sampling surface be at or near the same temperature as the test article surface of concern. It has been

found, for example, that some fairly heavy oils will not deposit

on surfaces above 4° C. Also of importance is the condition

that the sampling surface be of the same material as that of the test article. An outstanding example of this requirement is that many compounds will not adsorb on Teflon regardless of its

temperature, and conversely, methane and carbon dioxide are re30
tained on highly activated charcoal, alumina, and silica at room temperatures . Another parameter that is of concern in

sampling is that the sampling surface "be as near as possible to

the test article and have the same directionality and field of

view as the test surface. The most desirable sampling conditions would he, then, to sample the test article surface itself

on a real time "basis for identity and quantity of contaminants

and for the effects of the contaminants on the test surfaces.

Unfortunately, this is not feasible, and the sampling conditions

must be selected to satisfy the most important requirements of

contaminant analysis. For example, it is often important to know the sources of contamination in a vacuum chamber, so that

they may be found and eliminated. After they are eliminated,

it is desirable to evacuate the chamber again and repeat the measurements to determine the effect of the cleaning. In this

case, passive devices at cryogenic temperatures, and at locations all about the chamber, are sufficient. On the other hand,

when it is important to know the effects of contamination during testing, as in long duration simulated flights, only active

effects measuring devices with close matching of surfaces, temperatures, locations, and field of view will suffice. The various modes of sampling and analysis techniques available shall

be described to assist in the satisfaction of contamination

measurement requirements

Mass Spectrometry

This active or passive sampling technique has the greatest analysis potential; however, a great amount of effort is required to make it usable, and caution is recommended in considering its

application.

Mass spectrometers , with the exception of the leak detection

type instrument which is "tuned" to one peak, were formerly used

in a more passive contamination sampling role better suited to discrete samples and slower analysis scan times not related to real time test monitoring. With the recent advent of fast scan- ning and recording mass spectrometers , it is now advantageous to use the instrument on a real time basis.

Generally the mass spectrometer can be coupled either externally

or internally to the test volume under consideration. For ex- ternal use, the mass spectrometer has an ultra-high vacuum system which draws an in-chamber atmospheric sample through a heated sample line into a heated mass analyzer tube. The sampling line should extend into the chamber test volume in order

to obtain the most significant sample and to avoid sample con31
densation on the coldwalls. Externally coupled instruments are generally useful only for pumpdown and repressurization test phases, when sample pressures are sufficient for migration

through the line to the instrument. Each sample line will have

its own time constant, i.e., the time necessary for the entrance

of the sample into the line until its exit at the entrance to the ionization source region of the mass spectrometer. The time constant of the sample line will depend on line length, sample

pressure, line diameter, sample line temperature, and material

of tuhe wall . , . - .

At high sample pressures, the fastest response is ohtained with

a small diameter (3 to 7 mm.) line, which keeps the gas volumes

low. At lower sample pressures, where an auxiliary piimping line

which ordinarily helps reduce the gas flow into the analyzer

would not he used, a large diameter (7 to 25 mm.) line would

give the highest pumping speed and fastest time response. The

sample line design must he optimized for best response with any

given set of operation conditions, be it a single line with variable orifices or a series of sample lines valved in at differ- . ing pressure ranges. An internal mass spectrometer is a type known as the Remote

Residual Gas Analyzer (RRGA) which is essentially a mass analyzer mounted in the test volume and depends on the chamber itself for vacuum. Of necessity, this instrument can operate

only when the chamber pressure falls below the 10~4 torr region

and must be provided with automatic filament protection should

there be a sudden pressure surge. Since the instrument control

consoles are located external to the chamber, the necessary preamplifiers must also be located with the analyzer to enhance the

signal-to-noise ratio. The RRGA mount may allow rotation through

two mutually perpendicular axes to provide directional sampling

capability. The usual magnetic sector instrument mounted inchamber requires its own magnet, and its weight can present a mounting problem in directional sampling. A recent popular

trend in test volume mass spectrometers has been the small,

lightweight monopole and quadropole type instruments which

lessen mounting constraint and provide good linear spectral display (magnetic. type mass spectrometer mass scans are usually

non-linear with respect to mass numbers J

For either ^internal or external instruments, once the sample is obtained in the ion-source region and ionization of the gas is effected, the types of analyzer methods used may be widely differing. Upon electron impact, the ions produced are electrically acceleratedand filtered from the source region into the an- alyzer region, where the ion sorting may be accomplished by use

of any combination of magnetic, electrostatic, or electrodynamic

32
fields. Various mass analyzer types are the magnetic (180°

Dempster, 60° and 90° sectors), quadropole, and monopole types

already mentioned, and also the cycloidol, omegatron, RF, and

pulsed time-of-flight analyzers which are less commonly used.

Each of the above types has its own advantages and disadvantages,

and the choice among them for use in chamber sampling depends on

the requirements of resolution, scan speed, mass range, sensitivity, clean-up time, and inlet system. A very good listing and

specification comparison of commercially available mass spectrometers is contained in reference 32.

The mass spectrometer is an "absolute" analyzer which can be

made both qualitative and quantitative. One recurring problem

with the present rapid scan instruments is that there is too much useful data, and complex data reduction methods have to be

employed. Analyses of mixtures of compounds, without appropriate working conditions and on a poor resolution instrument, can

be futile. The problem is the indeterminability of percentage

of contribution of each component for a selected mass value.

Identification of possible chamber contaminants requires a very

extensive program of library referencing or "fingerprinting"

spectra of the suspected single components. If possible, these

single components should be run as series of mixtures of various

compositions under differing simulated chamber temperature and

pressure environments to serve as a standard for the instrument

sensitivity, resolution, and times of appearance. The end result

of the previous work will be an in situ calibration, a formidable task, indeed, for multicomponent compounds such as oils,

paint samples, or surface outgassings. Published fingerprint

data are not very useful, as a mass spectrum of a compound is unique for any mass spectrometer and may vary for different

sampling systems for the same instrument. However, general rules

of fractionation of molecules in mass spectrometer ionizing chambers are applicable in determining the type of compound present.

This is a very delicate exercise, however, since mass peaks from

several different types of molecules tend to make each other's

spectra ambiguous, i.e., there are often too many peaks. To begin, mass numbers of unknown peaks must be matched to, or interpolated between, mass peaks from a known mass marker. One then

searches for the mass peak of the unfractionated molecular ion,

if possible, and proceeds to smaller mass peaks, trying to recognize any patterns representative of certain functional groups.

For example, m/e peaks of 29, 39, 42, 43, 55, and 57 are representative of hydrocarbons; m/e 31, 45, or 59 peaks of alcohols; m/e 45 and 60 of carboxylic acids; m/e at 30 + (I4) of

amines; m/e 30 and 46 of nitro compounds; and m/e 65, 77, 78, 79,

91 + (H)^ for aromatic hydrocarbons. The above functional group

33
patterns are only a very few of many. Several excellent texts are availalole for functional group identification and description of typical cracking modes in mass spectrometry (30» 3"! » 32,

33); however, considerable expertise must be developed in reading mass spectra before this technique can be useful.

Mass spectrometers are now being linked directly to gas chromatographs , so that individual components of a mixture may be analyzed, thus eliminating ambiguities caused by the presence of

unknown mixtures. Gas chromatographic sampling of gases whose

vapor pressures are less than 10~5 torr is not possible, however, so that this powerful separating technique cannot be used

with mass spectrometers for real time contaminant analyses.

Again, one is cautioned in considering mass spectrometry as a contaminant analysis technique, because, though it is a very

powerful method, it also requires very great effort and expertise to analyze the wealth of data from each mass scan.

Quartz Crystal Mass Monitors

These devices measure the mass of contaminants deposited and are capable of active or real time sampling with directionality,

field of view, and temperature control. They consist of small,

thin quartz crystal pieces cut at the AT plane. The crystals

have aluminum or noble metal films deposited on each face which

are electrodes. An electric field is applied across the faces

of a crystal, which distorts piezo-electrically in a direction

perpendicular to the field (parallel to the faces). The natural

oscillating frequency is dependent upon the mass and temperature

of the crystal, although it is quite temperature insensitive in

the range 0° to 70° C. Thus with controlled temperature, the

natural frequency of the crystal will decrease as the crystal becomes contaminated; i.e., its mass will increase, and the effect

is analogous of adding mass to a vibrating spring. The crystal

mass is on the order of 0.1 grams, and its natural oscillating

frequency is in the range of 5 x I06 Hz. The frequency (f) of an oscillator, within limits, is inversely proportional to its mass

(m), f = c/m, where c is a constant. In differential terms df/f = dm/m, (5) which can be approximated to Am = -m Af/f, (6); as long as Af/f is small, the approximation is quite accurate.

The valid range of is about Hz, and the typical stability and sampling area of the system are such that contaminants

can be measured in the range of 10~4 to 10""^ g/m^, or approximately 0.1 to 100 monolayers of a heavy organic material.

34
The frequency of the oscillator may be monitored directly with

an electronic counter, but problems are generally encountered

in transmitting the weak radio frequency signal through the chamber; a preferred method is to beat the samping crystal's natural

oscillating frequency against that of a reference crystal that

is completely shielded from contamination and located only a few

centimeters distance from the sample crystal. This results in an audio frequency signal which is easily transmitted and measured with an electronic counter. It is interesting to note that

electronic counters, which are very stable, must use an oscilating quartz crystal as a clock, so that the beat method actually involves three crystals. Several sample crystals may be

beat against a single reference crystal in a multidirectional

apparatus. It is important when using the beat method that the

sample and reference crystal temperatures be identical, at least to within 1° C. An interesting phenomenon in the use of

quartz crystal mass monitors is that during the pumpdown phase,

the frequency may often increase, due to desorption of CO^ and

H^O from the crystal faces.

The above technique cannot discern the nature of the contaminants , and the sampling surfaces are limited to the type of

electrode material (metallic) that can be employed. The crystals begin to lose their linear frequency-mass relationship

when the contaminant becomes too heavy (/d f > 1o5 Hz), and this

limit may be lower if the contaminant is in the liquid, rather

than the solid, state. The crystals are apt to change oscillating planes, and thus the principal frequency, if mechanical

stresses develop. The crystals may be cleaned with organic solvents and reused. To remove the final monolayer of contaminant,

the crystal may be boiled in an aqueous solution of 3 percent

hydrogen peroxide and 1 percent ammonia. This type of mass monitor is in common use as rate indicators in thin film vacuum deposition processes.

In Situ Reflectometers

This active sampling technique provides crude relative information on the quantity of contamination and is of most benefit in

measuring the effect of contamination on specularly reflecting

surfaces. The device is comprised simply of a light source, a mirror, and a detector, which are mounted as a unit in the space

environment chamber and monitored throughout a test. A decrease

in detector signal will result from contamination on a mirror.

Whereas transmitting optics are quite insensitive to contamination, specular reflectance of mirrors is very much affected.

The reflectometer system may be sophisticated and tailored to satisfy the requirements of the investigation. Sources, detectors, and optical filters may be selected to view only a given

35
iDand of the spectrum. For complete spectral analysis, a sophisticated scanning monochromator system might be employed; however, increasing the complexity of a system increases the number

of susceptible optical surfaces, and also increases the probability of self-contamination. A sophistication of a non-scanning

system is monitoring the intensity of the incident beam, so that

the ratio of reflected to incident light, or specular reflectance, may be monitored directly. This avoids errors due to drifts in source intensity and detector sensitivity. An important factor in the design and employ of in situ reflectometers

is to assure that the obstruction of the view of the mirrored

surface by the instrument itself is minimized. Temperature control of the mirror is desirable. An additional detector may be

included that receives light reflected from the mirror that is not specular (diffusely reflected light). This is a delicate

system because the aperture of the detector must be low to avoid

view obstruction of the (sampling) mirror; the scattered, or diffusely reflected, light from small amounts of contamination

would be very weak. However, diffuse reflectance increases dramatically with contamination and is more sensitive to contamination than specularly reflected light. The only value of diffuse

reflectance measurement, though, would be to monitor the relative extent of contamination.

A typical decrease of specular reflectance at 253 • 7 wave- length due to two or three monomolecular layers of heavy oil

would be on the order of 1 percent. The increase of diffuse

reflectance of a mirrored surface, which is initially very low,

for the same contamination would be around 50 percent. The mirrors used in reflectance monitoring must be first surface coated.

Painted Radiometers

This is an active, effects measuring technique which also serves

as a radiative heat transfer monitoring method. This approach

involves placing a nude radiometer or calorimeter, whose coating

is identical to that of the test surface of concern, on the test surface and in good thermal contact with it. The radiometer must be so constructed that the temperature differential between its

back and sensing surface is less than 0.5° C. A thermocouple

must be located on the radiometer body to monitor radiometer tem- perature, Tj^. The output of the radiometer, V, is directly proportional to the radiative heat transfer at the test article

surface by the equation (T)^

M

(7)

36
where

S = solar a"bsorptance (for solar simulation)

Eg = solar simulation irradiance

= total hemispherical emittance for radiometer tem- perature T„K

= Stefam-Boltzman constant, 5'67 x 10 ^ wm ^ deg K

^

S-p, = sensitivity of radiometer "based on absorbed flux

M , ^ = net radiative heat transfer at the surface, rad , net

If the radiometer is monitored when the solar simulation is off, ES the emittance may he directly measured ^ = R (s)

This value may be substituted into equation (7) to determine o<'g

the absorptivity of the surface for solar simulation. Of course,

the value of solar simulator irradiance at the location of the

painted radiometer must be known. The best data the painted ra- diometer technique will afford, however, are changes in net ra- diative heat transfer at the test surface. Provided radiometer

temperature measurement is accurate, these data will indicate the

effects of solar simulation spectrum and irradiance, and of con- tamination on the test surface; however, it cannot discern between the two effects unless solar is off, at which time the effects of contamination on emittance can be measured. Should the

coating material itself photochemically degrade, however, emittance monitoring would not distinguish this degradation from

that of contamination.

Contamination Collection Units (CCIJ's)

This is a passive sampling, quantitative and qualitative analysis

method. It can provide bountiful information regarding identity,

extent, and source of contamination. The sampling is very simple

and inexpensive. This technique collects low levels of contamination over large surface area in order to obtain enough sample

to analyze. For example, if one monomolecular layer of heavy oil

is approximately 10"^ g/m^, then about 0.1 milligram will be

collected from a sampling area of one square foot, or 929 cm^.

Microgram quantities of organic materials can be analyzed with

gas chromatographs (with hydrogen flame ionization detectors)

and with mass spectrometers . Quantities as small as ten micro37
grams can loe analyzed with infrared spectrophotometers using

"beam condensing attachments.

A contamination collection unit is simply a ^00 x ^00 mm glass

or polished metal plate with 25 mm diameter quartz discs at- tached to each face. A thermocouple is attached to the plate

to monitor its temperatures during a test in the space environment chamber. These units may he placed anywhere in the chamber,

e.g., near pump ports, test articles, and solar simulators.

These units may be sophisticated by placing them in thermally

controlled mounts with covers that can be removed and replaced

during various phases of a test in a space environment chamber;

for example, one might be exposed only during the pumpdown

phase, another only during the test conditions, and still another only during repressurization. After they are removed from

the chamber, the quartz discs are detached from each face, and

their transmittances are measured at 200 nm on an ultraviolet

spectrophotometer. They are cleaned with ethyl alcohol or car- bon tetrachloride, and their transmittances at 200 nm are measured. The log of the ratio of transmittance of the cleaned to that of the contaminated disc (absorbance) at 200 nm affords a quick indicator of the relative contamination on that disc.

This may be very crudely converted to a quantitative value if a molar extinction coefficient of lO^m^/mole (typical of many or- ganic compounds at 200 nm) is used in the relationship:

fA

- 300

^^^

g

where ^ is the mass per unit area

To and T are the transmittances at 200 nm of the cleaned

and contaminated quartz disc, respectively ^ is the molar extinction coefficient

300 is the assumed molecular weight of a typical heavy

organic contaminant

The assumptions of extinction coefficient and molecular weight

make the quantitative values obtained from the quartz discs very

uncertain; however, a comparison of absorbances of the quartz

discs provides a quick look evaluation of the relative severity

of contamination about the chamber.

Each face of the CCU plate is rinsed down with 100 cm of spectroquality carbon tetrachloride. The rinsing has been shown to

38
'be "better than 90 percent effective in reinoving contaminants in excess of one monolayer. One monolayer of some organic material

will remain on the plate surface. This monolayer may be subject

to competitive displacement by other more adsorptive materials.

If the existing, pretest monolayer is of a sufficiently stubborn material, such as silicones, then at least 90 percent of all

the contaminants accrued during testing will be collected by the

rinsings. The rinsings of each face are concentrated to a volume of 3 cm5, and the absorbances (log of the inverse of the

transmittance) of the contaminants in the solvent are measured

in the region 2.5 to 10 micrometers on an infrared spectrophotometer. The optical path is from one to two millimeters.

The quantitative estimate of contamination is determined from

absorbance and extinction coefficients similarly to the method

used for quartz discs. The calculation is from the formula

-1 ^"^10 m /T . Y A

where To- and T are the transmittances of the solvent (CCl^)

and CCl^ solution containing the contaminant,

respectively, at wavelength

is the extinction coefficient of the contaminant, on a mass rather than molar basis, at wavelength "X . is the optical path of the infrared absorption cell (1 to 2 mm).

A,"CCU is the area of one face of a CCU. V is the total volume of solution (typically

3 cm^).

Some values of ^> for various compounds, with "X selected at minima of bands characteristic of the compound, are:

39
COMPOIMD A 1 ,jim -1

2 -1 2

6^^2 Nujol Mineral Oil 3.4 3. 18 SunVis 706

Hydraulic Oil 3.4 2.79

DC-705 Silicone Oil 3.2 7.0 18.5 7.9

Houghtosafe IO55

Tricresyl Phosphate

Oil 8.6 7.5

Polydimethyl Siloxane

oil, grease, or rubher 8.6 12.0

i The sensitivity of this method is about O.3 monomolecular layers

of heavy oil on a 3OO x J>00 miD CCU sin?face. A different infra- • red solvent may he used, such as carbon disulfide, if contaminant absorbances in the region 7»5 to 1 5 1^ are sought.

The solvent, CCl^, is very transparent in the region 2.5 to 6.0 yus where absorption of 0-H, N-H, C-H, and 0=0 bonds occur. From

these data, an estimate of the total hydrocarbons, alcohol and

amines, and carbonyl (C = O) containing compounds can be made.

The solution is then evaporated to dryness on a small (12 x 12 mm) piece of KRS-5 infrared transmitting material (non-hygroscopic) or in 25 milligrams of potassium bromide (KBr) pellet

powder and pelletized in a microdie (1.5 millimeter diameter).

The sample - KBr powder mixture may be mechanically mixed, or the sample may be transferred to the KBr by the lyophilization

(freeze drying) technique (28). Lyophilization is a time consum- ing and tedious process, but affords quantitative spectra because

of resulting homogeneity of the sample in the pellet (29). The

KRS-5 piece or the pellet is placed in a beam condensing attachment of an infrared spectrophotometer, and the spectrum can be

measured from 2.5 to 40 micrometers without interference from

the media. If the KRS-5 piece is used, the sensitivity of the

technique is about 10 micrograms, corresponding to about 0.1 monolayer of a heavy oil on a ^00 x ^00 CCU. The spectra may

be matched against "fingerprint" spectra of known materials

used in and about the space environment facility. Figure 1 depicts infrared spectra of materials commonly used in vacuum chambers. In some cases the contaminants may be readily identified;

i.e., their spectra are unique, as in epoxy paint outgassings

(alkydresins) and DC-7OO series oils (27, 34) • In other in40
stances, where the contaminants are basically of the same compositions snch as petroleum based fluids and polydimethyl siloxane

compounds (silicone rubber and many silicone oils and greases),

identifications would be restricted to the type of compoumd (petroleum based hydrocarbon, polydimethyl siloxane, etc.) (27, 3^). One very misleading consequence in matching contaminant infrared spectra to fingerprints is that perhaps the volatile component of a contaminating compound may constitute only a very small

fraction of the compound fingerprinted, in which case this com- pound would be erroneously exonerated by fingerprint matching

as a contamination source. It is preferable, if possible, to collect only the outgassings of materials for fingerprinting,

in order to avoid this pitfall. Identification of contaminants becomes difficult when small quantities of some contaminants are present with large quantities of the others. In this

case, the infrared spectra of the lesser contaminants becomes

obscured by those of the more dominant ones. It is preferable

to fractionate, or separate, the contaminants, and this is facilitated in a very sensitive fashion by gas chromatography.

If the contaminants concentrated on the KRS-5 piece or in the KBr pellet are redissolved in an organic solvent, such as CCl.

they may be injected in a gas chromatograph. This instrument

separates the contaminants according to chemical composition

(fractionates) from a stream of inert carrier gas flowing

through a long, narrow "column" (usually coiled tubing) packed

with a solid material of uniform mesh size, each particle of

which is coated with a low vapor pressure organic material

known as the stationary phase. The effluent of the column

passes through an electronic detector whose current is sensitive

to impurities (fractions of contaminants) in the carrier gas. For most detectors , change in detector current is directly proportional to the amount of the sample's fractions in the effluent stream. The gas chromatograph is strictly a relative instrument, that is, the time between sample injection and occurrence of a fraction of the sample at the detector is a function

of the individual column used, flow rate of carrier gas, and

column temperature, which is often increased at reproducible

rates after injection. Thus if materials are to be "fingerprinted" for gas chromatography, the procedure must be performed

for each type of column, preferably each individual column, used.

A general purpose system consists of a dual column, dual hydrogen flame ionization detector chromatograph using 2 meter long

columns of 2 to 4 millimeters internal diameter packed with 60

to 70 mesh solid coated with 5 percent (by weight) SE-30 silicone gum rubber and operated at 3OO , 310, 315° C for column, detector, and injector, respectively (37). This technique provides an accurate quantitative analysis of contaminants if the

fractions can be reliably matched to fingerprints. The sensitivity of gas chromatography is now typically in the sub-microgram range, so that quantities of contaminants much less than

41
0.1 monolayer on a CCU face can be identified and their masses measured. It is also possible, but with great difficulty, to individually collect the effluent fractions of the CCU contaminants from a gas chromatograph. When this is done, their infrared spectra may be recorded (if the sample is heavy enough)

and also their mass spectra may be measured. This is a very

powerful technique, as it allows the infrared and mass spectra

of individual compounds to be unambiguously measured, under

which conditions the infrared and mass spectra are very capable

of identifying the fractions, to the extent of determining most

of the molecular structure, including molecular weight of the

compounds. The linking of a gas chromatograph effluent stream

to a mass spectrometer inlet system is now becoming a common

practice, and this combination is highly recommended in CCU analysis. The combination GC-MS eliminates the one great disadvantage of mass spectrometry in identifying compounds from mult

component mixtures

Optical Samples

Optical samples may be mirrored, transparent, or opaque and nonspecular. The use of first surface mirrors permits the determination of the effects of contamination on similar mirrored

surfaces. For this application, it is recommended that the coat

ing, view factor, directionality, location, and temperature of

the sample mirrors be as similar as possible to the same parameters of the test surfaces of concern. The specular reflectance

and diffuse reflectance spectra of the sample mirrors and of one

or more laboratory control mirrors are measured immediately before the environmental test begins and again immediately after

the test is concluded. The control samples are maintained in a sealed container in the measurement laboratory and serve to indicate changes in the reflectometer and errors in handling technique that occur between pre- and post-exposure measurements.

The wavelength range of specular reflectance measurement should

be selected to match that range in which the test surfaces are intended for use (soft X-ray, vacuum ultraviolet, near-ultraviolet, invisible, near infrared, infrared, or far infrared). The

diffuse reflectance spectrum is most conveniently and advantageously measured in the range 200 to 750 nm (near ultraviolet and

invisible) and is a sensitive indicator of the extent of contamination. Integrating sphere attachments for spectrophotometers

using magnesium oxide or barium sulfate coatings are commercially available. Likewise, specular reflectance attachments are available for vacuum ultraviolet, ultraviolet, visible, and infrared spectrophotometers. Absolute accuracy is not necessary

in this type of measurement, the important information required

is only the extent of change, or degradation, of the optical surfaces due to environmental testing. The same approach is applicable to transparent samples , wherein these samples are matched t

42
the pertinent test surfaces, and their transmittance spectra, and

that of control samples , are measured before and after exposure

to simulated space environment.

Samples coated with thermal control coatings identical to those

of the test article may he also employed in the same manner. Integrating sphere attachments for spectrophotometers also lend

themselves to ahsorptance measurements in the solar region, and

precise emmisometers can he employed for the measurement of normal total emittances (in the infrared region) of such samples.

It may he commendable to control the exposure of optical samples

to selected phases of the environmental test by remote, con- trolled shuttering. The shutters and actuating mechanisms

should be very thoroughly cleaned, preferably including vacuum

bakeout, in order to avoid their own contamination of the optical samples

Optical samples will indiscernably indicate degradation by environmental conditions other than contamination, such as ultraviolet radiation and thermal shocking, and thus contamination may

be erroneously blamed for such degradation when optical samples

are employed.

Microscope Slides

This is a passive technique for the determination of particulate

contamination. Microscope slides, preferably with calibrated

grids imprinted on them, are cleaned and are photographed through

a microscope at significant times before being placed in the en- vironmental chamber. The slides are again microphotographed at significant times after removal from the chamber and the amounts

of particles in successive size ranges are counted. Control

slides are photographed and counted at the same time the sample

slides are, except they are maintained in a particle-free envi- ronment while the slides are being exposed. The exposure of the

samples may be controlled by remote shutterings. Cleanliness in

this type of mea.surement is understandably critical (1, 7> 8)»

43
REFERENCES

Particulate Contamination

1. Sivinski, H. D.; Whitfield, W. J.; Pauhamus , J. A.: Contamination Control Principles , NASA Special Publication

SP-5045, Natural Aeronautics and Space Administration,

Washington, D. C, I967, PP. 41-46.

2. Method of Test for Continuous Counting and Sizing of Airhorne Particles in Dust-Controlled Areas hy the Light

Scattering Principle (for Electronic and Similar Applications) , ASTM F30, American Society of Testing Materials,

1916 Race St., Philadelphia, Pa.

3. Whitfield, W. J.: Monitoring a Class 100 Clean Room , SC-R-956, Sandia Corp., Oct. I966.

4 . Procedure for the Determination of Particulate Contamination

of Air in Dust Controlled Spaces by the Particle Count Method , SAE-ARP-743A, Society of Automotive Engineers, Inc., New

York.

5. Johnson, J. D.: Contamination Analysis and Monitoring , Proceedings of the National Conference on Spacecraft Sterilization Technology, NASA Special Publication SP108, National

Aeronautics and Space Administration, Washington, D. C,

„. 1967, pp. 293. . -

6. Dwyer, J. L.: Contamination Analysis and Control , Reinhold,

New York, I966.

7 . Handbook for Contamination Control on the Apollo Program , NASA Handbook NHB 53OO.3, National Aeronautics and Space

Administration, Washington, D. C, Aug. I966.

8. Sivinski, H. D.: Contamination Control Handbook , NASA Special Publication SP-5076, National Aeronautics and Space Administration, Washington, D. C, 1969> PP. 63-82.

44
Desorption, Suljlimation and Evaporation

9. Moe Iwyn-Hughes , E. A.: Physical Chemistry , 2nd Ed., Pergamon

Oxford, 1961, pp. 921-983, 1051-1113.

10. Santeler, D. J.; Holkeboer, D. H. ; Jones, L. N.; Pagano , P.:

Vacuun] Technology and Space Simulation , NASA Special Puhlication SP-IO5, National Aeronautics and Space Administration, Washington, D. C, I966, pp. 175-213.

11. Deboer, J. H. : Dynamical Character of Adsorption , Oxford,

Toronto, 1953.

12. Dayton, B. B.: Outgassing Rate of Contaminated M6tal Surfaces , Transactions of Eight National Vacuum Symposium and

the Second International Congress on Vacuum Science and

Technology, Vol. I , Pergamon, 19^2, pp. 42-57.

13. Dushman, S.: Scientific Foundation of Vacuum Technique , Wiley, New York, I962, pp. 491-587-

14. Gray, D. E. (Ed): American Institute of Physics Handbook , McGraw-Hill, New York, I957, pp. 4-I3I ff.

15. Ehrlich, G.: Molecular Processes in Adsorption on Metals,

Transactions of the Eight National Vacuum Symposium and the

Second International Congress on Vacuum Science and Technol - ogy, Vol. I , Pergamon, I962, pp. 126-1 45.

16. Espe, W.I Materials of High Vacuum Technology, Vol. 3 , Pergamon, Oxford, I968, pp. 235 ff.

17. Fulk, M. M. ; and Horr, K. S.: Sublimation of Some Polymeric

Materials in Vacuum, Transactions of the Ninth National

Vacuum Symposium , Macmillan, 19^2, pp. 324-333.

18. Boebel, C. P.; Mackie , N. A.; and Quaintance , C.C.: Outgassing Studies of Space Materials, Transactions of the

Ninth National Vacuum Symposium , MacMillan, I962, pp. 3O7-

310.

45
Decomposition and Chemistry

19. Brandrup, J.; and Immergut , E. H. (Ed): Polymer Hand"book , Wiley, New York, I966, pp. V-5 - V-12.

20. Lenz, R. I) . : Organic Chemistry of Synthetic High Polymers , Interscience, New York, I967, PP. 733-76?.

21. Madorsky, S. L. : Thermal Degradation of Organic Polymers , Interscience, New York, 1964.

22. Fieser, L. F.; and Fieser, M. : Organic Chemistry , Heath,

Boston, 1956, pp. 399-4O8, 412-4^6.

23. Madorsky, S. L., Ibid , p. I32.

24. The Boeing Co. , Kent, Washington, June ^0 , I966. Results

of an Investigation on the Effects of Solar Simulator

Ultraviolet Irradiated Freons on First Surface, SiOx Overcoated Aluminized Solar Simulator Mirrors.

Infrared Spectroscopy

25. Silverstein, R. M. ; and Bassler, E.C.: Spectrometric Identification of Organic Compounds , 2nd Ed., Wiley, New York,

1967, pp. 64-109.

26. Henniker, J. C: Infrared Spectrometry of Industrial Polymers , Academic, New York, I967.

27. Hummel, D. 0.: Infrared Spectra of Polymers , Interscience,

New York, I966, pp. 152-153.

28. Instructions for KBr Microdie, Perkin-Elmer Corp., Norwalk,

Conn., 1966.

29. White, R. G.: Handbook of Industrial Infrared Analysis , Plenum, New York, I964, pp. I23.

46
50. Silverstein, R. M. ; andBassler, E.G.: Ibid , pp. 4-63.

31. Bieman, K. : Mass Spectrometry, Organic Chemical Applications , McGraw-Hill, New York, I962, pp. I7I-I72,

32. Industrial Research, I97O Instruments Specifier , Industrial

Research, Inc., Beverly Shores, Ind., 1970, pp. 32, 39-41-

33. McLafferty, P. W. (Ed): Mass Spectrometry of Organic Ions , Academic, New York, I963.

34. Hummel, D. 0..: Ihid, pp. 112-113.

35. Lenz, R. W,: Ihid , pp. 307.

36. Hummel, D. 0.: Ihid, pp. I79.

37. Langdon, W. M. ; and Ivanuski, V.R.: Backstreaming from Oil

Diffusion Pumps , Report No. IITRI-C6O3O-I 0 , IIT Research

Institute, Technology Center, Chicago, 111,, July I966,

pp. 17, 163-170.

47
Figure lA. Infrared Spectrum of Petroleum Based Oil or Grease (Hydrocarbon)

41

Figure IB, Infrared Spectrum of Tall Oil Alkyd - Used in

Paints - Outgassed From 3M Nexte3®Paints

4a
Figutr ID. Infrared Spectrum of DO700^ Series Oils

(Methyl-phenyltrisiloxanes)

49

Paper No. U

CONTAMINATION STUDIES IN A SPACE SIMULATED ENVIRONMENT

David Lc Enlow, Research Physicist, Environmental Sciences Laboratory, Re-entry and Environmental Systems Division, General

Electric Co= , Phila, Pa.

ABSTRACT: Biological and optical contamination problems can be

best investigated experimentally. Studying the dynamics of microorganisms in a vacuum environment showed that shock forces in the order of a 1000 g's will liberate 20 percent of the previously bound

particles and the effect of simulated attitude control gas jet firing

will remove more than 50 percent of the bound particles.

A potential source for optical degradation to infrared systems is identified as the absorption phenomena of condensed volatile products

from certain poljrtneric materials. The effects of this condensation were analyzed by a combination of surface photography and IR spectroscopy. These preliminary investigations point out the need for

proper materials selection for far infrared space-borne optical

systems.

KEY WORDS: planetary quarantine, shock force, fluid dynamic force,

microorganisms, contamination control, testing, evaluation, space

simulation, optical properties, condensation, volatile products, poly- mers, far infrared, absorption.

INTRODUCTION

Contamination problems in previous space system programs have

to varying extents endangered full mission objective realization. In particular, the problems of biological and optical contamination have

in the past and will be in the future, increasingly important. This is evidenced by the planetary constraints which are placed on the current Viking Mars Landing Program and the ever increasing level of sophistication and operational sensing requirements of space-borne

optics. This paper will discuss anal;^i;ical and experimental investigations for both of these areas of contamination.

51
1.0 BIOLOGICAL CONTAMINATION

Biological contamination is of prime consideration for the Viking

Mars Landing program in which a sterile lander will investigate the existence of life on Mars. The overall Martian planetary biological

integrity will be insured by the provision of an International Planetary

Quarantine Agreement which specifies the overall probability contaminating Mars within the next twenty years shall be less than

1 X 10" 3. The constraints imposed by the Planetary Quarantine Requirements on the Viking Program will consist of designing, fabricating, and launching a sterile lander which will be protected during

the trans -planetary cruise by a bioshield. To realize these requirements, extensive experimental evaluation (we will discuss sterile lander recontamination only) under space sinrilated, as well as spacement mission event simulated environments, must be performed„

1. 1 Analytical Studies

Recontamination refers to the phenomena of the sterile lander

becoming contaminated via any source at any time after terminal

lander sterilization. Detailed analytical studies on the recontamination phenomena for a Mars Lander Mission were performed. First, a definition of recontamination mechanisms was generated (see Fig, 1). The attitude control gas particle transport mechanism is one whereby

particulates may be removed from non-sterile portions of the spacecraft and transported to sterile areas. The availability of microorganisms is determined by the interactions of bound particles and

micro-meteoroid impacts, shock and vibration events

To determine the importance of each mechanism, they were

singularly analyzed in terms of their controlling forces. The controlling forces were found to be orderable in three main areas

(1) Removal forces (input forces ) - shock and vibration, fluid dynamic, surface sublimation, and impact; (2) Adhesive forces (force

bounding particles to surface) - capillary and electrostatic; (3) Dynamic Forces (forces controlling unbound particles) - electrical

gravitational, and fluid dynamic. Attempts to quantify these forces

analytically turned out to be tractable in some cases but for the most

part very difficult » In summarizing the analytical work [1] , it should

be pointed out that the key properties to understanding the recontamination phenomena were the adhesion and electrostatic characteristics of small particles and their interaction with surfaces; that

[1] The number in brackets refer to the list of references appended

to this paper.

52
analytical analyses was not sufficient to define the controlling forces;

and that experimental investigations must be performed under simulated space flight conditions o

lo2 Experimental Program

1.2.1 Introduction

An experimental test program was designed to investigate the adhesive and electrostatic surface interactions for microorganisms

on surfaces [2]. The experimental approach (Fig. 2) consisted of developing a partulate deposition technique to enable uniform and repeatable depositions be performed on various types of surfaces. To

determine the number of microorganisms removed or remaining on

a test surface, bio-assay methods were selected and validated for high removal efficiency and repeatability » To investigate the adhesion phenomena of potential contaminating particles, force produce

test assembly systems were designed and fabricated to simiilate specific forces; i.e. , centrifugical , shock impulse, and fluid dynamic

flow. The necessity for space simulation in the investigations required that force producing systems be implemented under vacuum conditions, and therefore must be vacuum system compatible.

The aerosol deposition technique (Fig. 3) was finally selected

after many variations , consisted of placing 100 mg. of microorganism

B, subtilis var. niger in the spore form into the aerosol head behind

a paper diaphram. Pressure was applied to the head and the dispersed

cloud was allowed to circulate in the chamber for a half an hour, the planchets (1 in. diameter test metal discs) were then exposed for several hours. This technique enabled a deposition uniformity of +15

percent to be attained. The microbe-deposited planchets were used

as the test specimens. The ultrasonic rinse bioassay technique was

selected because of its high removal efficiency (>97 percent repeatability) o The two major recontamination events to be simulated were the shock environment and the thruster (altitude control firing) flowo The

techniques used to simulate these mission events in terms of evaluating their importance as a recontamination mechamisms were greatly controlled by the requirement for vacuum environment.

1. 2o 2 Vacuum Shock Experiment

To simulate the shock environment, that is, impart a maximum

of 1000 g's (1 millisecond duration) to a test surface under vacuum

53
proved difficult. Since the problem of absorbing the shock recoil

energy by any vacuum chamber, its connections, and seals, is a very

serious threat to maintaining vacuum integrity, an approach was taken

to connect a small highly durable test assembly to a vacuum chamber

by a flexible connector, with the test assembly mounted rigidly to a shock test machinco Several designs were conceived; the final as- sembly (shown in Fig. 4a) was made of stainless steel 304, with re- movable top lid and vacuum gasket of Vitono It was connected to the vacuum system by a stainless steel flex hose which was found to be

far superior to earlier used thick wall tygon (shown in Figure 4a,

which was found to present a wall diffused gas load thus preventing a

test low base pressure. ) The sterilizable test chamber accommodated four planchets at one time. The supporting hardware (Fig. 4b)

shows the flex hose mounted to the vacuum chamber via a pneumatic

safety valve, connected to the test assembly which in turn is mounted

to the shock test machine. The vacuum pressure was measured at the test assembly by a hot cathode ionization gauge which was con- nected to the test assembly via a valve so that the tube could be re- moved at the time of the shock pulse. The shock pulse was measured

for permanent record by a microminiature accelerometer , charge

amplifier, peak-hold meter, and oscilloscope/camera. The vacuum

system used in those tests and in all the biological contamination ex- periments was a Kinney Vacuum System (Fig. 5). It consisted of a

6 in. diffusion pump (400 i/s) backed by a 0. 9 i/s holding pump.

The pump down from ambient was done by a (7. 1 i/s ) Roughing pump.

The diffusion pump had both water and LN^ trapping, with the LN^

level maintained by a controller. The pressure was measured at several pump locations by thermocouples and the vacuum chamber *by an ionization tube, with an identical tube and gauge at the end of the

test assembly. Since the testing required reasonably long term durations , a pneumatic safety valve was mounted between the vacuum

chamber and test assembly. Since the testing required reasonably

long term durations , a pneumatic safety valve was mounted between

the vacuum chamber and test assembly to prevent excessive backstreaming in the event of a power failure. The pressure profiles were maintained on both the vacuum chamber and test chamber, a

typical run is shown in Fig. 6. The pressure in the test assembly was in high 10"6 torr range, whereas the vacuum chamber pressure was at 5 x 10~8 torr. This large difference is due to the low conductance of the 1 in. diameter connection hose.

The procedure for the vacuum shock experiments was to place

four planchets into the previous sterilized test assembly, mount it to the shock machine, connect the hose and vacuum pump to a pressure

in the 10"^ torr range and allow the system to remain under vacuum

for varying lengths of time; then close the valve to the test assembly

54
ionization tube, remove the tube, and shock the assembly o To re- move the test assembly, a noninterfering bleed-up technique was

required. The turbulant flow conditions can occur during equilibrium

balancing of a test chamber [3]. The results of analysis and an

attempt to analytically determine bleed-up effect on microbial re- moval showed that an experimental evaluation was required. To determine the appropriate bleed conditions , the shock test procedure was repeated without imparting a shock impulse and the microbes re- moved as function of time duration of bleed-up (which was controlled

by a precision leak valve; sensitivity 10" 13 cm^/s at standard at- mosphere). It was found that if the time exceeded one-half hour there were virtually no microbes dislodged. Therefore, all shock tests were bleed up to ambient over an hour duration to insure isolation of the removal forces (Fig. 7). The shock tests measured the percent removal of microorganisms as a fxmction vacuum level duration, the pulse impulse was

taken as worst case - 1000 g's (1 millisecond sine pulse). The summarized results of tens of tests are shown in Fig. 8o The removal

was shown to be dependent on vacuum level duration and that the re- movals are not large; these two factors point up that the prediction

that the strong capillary forces are in effect as the water at the particle surface is probably still present. It is well known that water

vapor can remain in vacuum systems after long pumping [3]. For

this reason, future experiments will have to either be longer term

( > 200 hours) and/or include a moderate non-interfering thermal

bake of the test surface (less than 100 C). Conclusions from our limited investigations showed that shock impulse removal forces did not present a major removal force thus

shock generating spacecraft sequencing events did not present a clear

threat as a recontamination mechanism.

1.2.3 Fluid Flow Vacuum Tests The next potential recontamination event to be simulated was the thruster (attitude control gas jet) firing and its effect on the removal

of microorganisms. At various points in a trans -planetary cruise,

the thrusters on a spacecraft will fire to modify the trajectory, lock on to earth signals, or stabilize. The analytical evaluation of the

effect of a gas jet plume expanding into a vacuum, colliding with particles on a surface , was found to be very difficult without proper experimental data. An experimental program was planned, designed,

and performed. A chamber was designed that would accomodate one

planchet, be sterilizable, and vacuum compatible. The chamber

(Fig. 9) was fabricated from a small stainless steel cylinder with

55
the central portion bored out„ The planchet was placed looking downward and into the vacuum o The test chamber was connected to a fil- ter assembly which trapped unbound microbes moving during a molecular flow condition o The attitude control jet firing was simulated by

a rapid pressure excursion controlled by a precision leak valve. To

indicate why this analogy and simulation was selected, the topic of free expansion in a vacuum environment is discussed.

The expansion of a gas jet plume in space has been investigated

by NASA-Langley [4] and others [5, 6] initially to increase understanding of the lunar landing problems. This work can be directly

applied to the same phenomena in interplanetary space since the near surround density is very low relative to the jet plume

The expansion of a gas jet into the vacuum environment is such

that the majority of mass and momentum of the plume are in its center section in which the density decreases along and normal to the centerline axiSo The mathematical investigations of this phenomena

have led to the following expressions (4):

P = P (l+yM ^ (1 +-^—^ M ^

r o e 2 e

e

k M

^ - exhaust gas kinetic energy

y-1 e exhaust gas internal energy

where y = C^/C^ ratio of ideal gas specific heats, = Mach

number, h = distance along streamline from nozzle exit, r

raciial coordinate normal to axis, r = nozzle exit radius, ft =

streamline inclination angle relative to axis, P = static pressure on

surface beneath nozzle, = normal shock recovery pressure at nozzle exist, P = nozzle stagnation pressure, and maximum

angle of plume offcenterline o See Figo 10 for the Nozzle Geometry

Configuration.

Using the NASA-AMES report on compressible flow, expressions

for P (pressure at nozzle exit) in terms of P^ (pressure in the jet

chamBer) were generated [7]o Fig. 11 shows a listing of calculated

exit pressures and surface pressures considering nozzle configurations for exit to throat ratios of 10, 100, and 1000 and a chamber

pressure of 100 times ambient pressure 10^ torr). Most attitude control gas jets have a nozzle ratio of 5 to 100; therefore, looking

at Figo 11, this corresponds to a surface pressure rise from interplanetary vacuum, 10" 1^ torr to 2 x 10-2 torr (for 30 degrees off the

k+4

(cos ft) where

56
centerline and 3o05m from the nozzle exit) for A /a = 10, for Ag/A^=

100 the pressure at the plume impinged surface would rise to 1. 7 x

10"^ torr. The analytical predictions of the equations for the flow

field of an exhaust plume were shown to correspond very well with ex- perimental studies [5] when the distance from nozzle exit (h) is great- er than re (i.e. , h> 50 r^). It was shown, however, that the approximation does hold to a reasonable extent even for hs 10 re [4]. Considering the validity of the analysis it can be stated that the ex- perimental technique should follow a procedure whereby the pressure

is allowed to rise to levels near 2 x 10~2 torr in varying time durations. One difference in the two events is that the time of the pressure excursion under a jet firing is probably very fast (<1 s) whereas

the laboratory experiment was found to be limited to a 5 second

minimum.

The fluid flow experiments procedure was similar to that of the vacuum shock tests except no flex-hose was used with the test as- sembly being mounted directly to the pneumatic valve; and the force

producing system was a precision leak valve with the entrance port

parallel to the planchet test surface (see Fig. 9)o The experiments were performed determining the percentage of particles removed as

a function of vacuum level duration and time duration for the pressure excursion (typical 10~5 to 10"^ torr to 10~2 torr), A characteristic pump down pressure profile and bleed-up, with the pressure

measured by an ionization tube or thermocouple, is given in Fig, 12 and 13 respectively. Fig, 14 and 15 show the results of these ex- periments. In the dependence of the percent microbes removed on

the vacuum duration is shown. As expected, the longer the duration

the larger the number removed. This is understandable since the adhesive forces are decreasing as the capillary force component is being diminished by desorbing vapors at the particle-surface interface. More positive discussion of this phenomena would be possible

if heating of the test surface could be performed on a microbiological

non-interference basis. Fig. 15 shows the dependence of the time

excursion of the percent removed. The shorter the time the more

removed. All runs were done under similar vacuum conditions to accurately isolate the variables. The shorter time duration for the pressure excursion corresponds to a comparatively larger mass rate

flow, i.e. , m/dA where m = dm/dt. This would imply a large dp/dt,

where p is the composite momentum which is imparted to the surface

or particles on the surface. This is assuming that for the time domain of interest no significant boiindary layer is established at the surface of test planchets. The increasing dp/dt corresponds to an

increasing imparted force. This analog is simplistic but of probable

validity for free molecular flow regimes.

57
l„2o4 Conclusion: Biological Contamination Experiments

It can be observed that the particle removal by shock forces are

not large as was predicted. This may be due to the still present

bound surface water allowing the adhesive force to remain large. The

potential recontamination threat presented by shock impulse producing events (bioshield separation, staging, positioning of panels) is not as significant as previously proposedo This is not to preclude the necessity for longer duration vacuum testing.

The effect of the fluid flow forces on microbial removal under a simulated space environment show the importance of attitude control gas jet firing as potential recontamination mechanismo There are

critical events during a Mars Landing Mission when this phenomena

may present its greatest threat (ioC. , at the lime of bioshield separation) o In considering these results one must evaluate the simple

direct approach of simulation and determine how well this approach

meets the real case.

2.0 OPTICAL CONTAMINATION

Contamination effects on space-borne optical systems have presented problems that can degrade a mission performance. Optical

contamination may be defined as any surface phenomena which causes

the incident radiation's optical path to be deviated or intensity attenu- ated from the design specifications for the optical system. The effects of optical contamination have been observed on Gemini through

Apollo, as well as some military programs o The cause stems from

several sources: liquid dumps, plume interactions, cryogenic dumps

and space thermal vacuum effects on polymeric materials o Most of the past contamination studies have considered optical systems which

operate in the ultraviolet and visible sensing region of the spectrum

Discussed herein will be optical contamination for optical systems

which will operate in the mid to far infrared. Especially those systems requiring thermal control at cryogenic temperatures » The

rationale for such a study is that the previously known second order

material effects due to the space thermal-vacuum environment may

well become first order considerations. The program objective is, to analytically and experimentally investigate the physical phenomena

of optical degradation for cooled space -borne optical sensing systems

detecting in the mid to far infrared. Fig. 16 shows flow chart outlining the approach taken and will be discussed in the following sections c

58
2ol Analytical Studies

To best evaluate and understand the occurrence and effect of contamination on optical systems (cooled, mid to far infrared) analytical

studies were initiated. Considered are the contaminating events that can occur due to the sublimation and outgassing of bonding, sealing,

and encapsulating polymeric materials that selectively condense on

critical optical surfaces (for temperatures lower than the materials).

To do this task, the past studies on space vacuum effects on materials were reviewed. One of the most comprehensive reports (8), investigated the outgassing behavior (i.e, , total weight loss and volatile condensables) of some 350 polymeric materials. These experiments were performed under the thermal-vacuum conditions of 125 C and

approximately 10-6 torr. The total weight loss for the materials

tested (at a temperature of 125 C) ranged up to 30 percent for some

sealentSo The effect of curing conditions was found to be significant

with the weight loss decreasing for increasing thermal (>25 C) duration curing (see Fig. 16) [8,9], This is due to the various chemical

curing mechanisms that are occurring for each specific cure condition.

Long term curing cycles or elevated temperatures will minimize the various effects. The sample size has been shown to be important,

with thicker samples (greater than l/2 in. cube) exhibiting long term

outgassing [8,9]. In Reference 9, volatile condensable material

(VCM) experiments were performed in which the percent of volatile condensable mass collected at25C to sample total mass was determinedo For most sealents and adhesives the VCM was under two percent with the majority under one percent (see Fig. 17). Information

of this type is most significant for optical contamination effects be- cause it is the condensable materials that may cause degradation. To

identify the constituents of the VCM and outgassed materials mass

spectroscopy and qualitative infrared spectrophotometry were performed. The IR spectral analysis was found to be useful for determining optical effects of these potential contaminates and will be discussed later. Other studies have been performed that determine total weight loss and VCM for thermal vacuum conditions of <10-'^ torr and < 200C with the VCM collectors at 25C [10]. Consider now the outgassing effects coupled with the phenomena

of condensation, nucleation and growth. One of the most important

factors in the investigation of optical contamination is the amount of condensable materials that is collected on cooled optical surfaces.

This determination depends on differences in the rates of outgassing

(evaporation) of materials.

59
For a pure compound the expression for loss of material by evaporation and/or sublimation is: W~ 6x10"! P (M/T)-'-^^ where

W = rate of evaporation in kg/m~^/s"-*^, M = molecular weight, T =

absolute temperature, P(T) = vapor pressure (in torr), and A, B =

constants for each material.

Fig. 19 shows the vapor pressure of pure compounds is a function

of its temperature and may be expressed by Log P = A - B/T or P = A — B /T" 10 . It also points out the obvious that as the temperature is in- creased, the vapor pressure increases, and similarly the evaporation

rate increases. The evaporation rates for silicone polymers (shown

in Fig. 20) shows the dependence on temperature as well as molecular

weight (given by n-the number of dimethyl-siloxane groups in a molecule). Thus one can see that low molecular weight polymers have a much greater evaporation rate hence the expulsion of low n polymers

are first to be observed under thermal vacuum conditions. This effect does not follow directly the above given expression for the rate of evaporation. For polymeric materials composed of various molecular

weight molecules it is not simple to estimate the rate of evaporation

because of the effects of diffusion of particular molecular species

through the bulk material, distribution of stages of curing, and nonideal and non-equilibrium vapor pressures.

The absorption or condensation of the outgassed material onto a cooler surface is a commonly observed phenomena since the high vacuum pumping speeds at pressures < 10~6 torr use cryosorption

techniques as a standard practice. The build up of molecules on a surface may be expressed by: N = nt (molecules /m^) [12] where

N = molecules /unit area, n = number of molecules striking a unit surface/unit time, and t = average remaining time. The number of molecules striking a surface 1 m^ per second is: n = 3. 52 X 102 P (MT)

-l/2 where M = molecular weight, T = absolute temperature (K), P = vapor pressure of material (torr). The average sitting time (t) of a molecule on a specific surface can be expressed by the Frenkel

equation : t = to exp (Q/RT) (i2), where t^ ^ 10-12 - 10-14 g and is related to the lattice vibrations of the solid surface, where: Q =

heat of adsorption (J), R = gas constant (8. 314 x 10^ j/Kgmol.K and

T = absolute temperature of the surface (K). A listing of some characteristic values for this expression and the significance of cryogenic

temperatures for molecular trapping are shown in Fig. 21. Another approach to the condensation and adsorption phenomena

is by surface energy considerations. When a molecule hits a surface

three conditions are possible: (1) the molecule may return to the gas phase, (2) it may be trapped in the shallow potential well (minimum

for physical adsorption), (3) it may be trapped in the deep potential well (minimum for chemisorption) [13]. The possibility that item (2) is the result of a collision is defined as the condensation coefficient c

60
and the possibility (3) is the result of a collision is defined as the

sticking probability s. The effect of substrate temperature on a c and

s is shown in Reference 13. For example, H2 and Ne gas depositing on glass both show a 30 percent increase in c for a temperature decrease 100 C down to OC . Similar quantitative data is not known to exist or be available for polymeric condensable products. However,

the overall trend of increasing c and s values for decreasing collector substrate temperatures (T <273 K) are realistic for these materials.

Similarly little information can be found describing the nucleation

and growth of a polymeric film for temperatures less than 273 The

experimental program to be discussed in the next section will hopefully help to increase the understanding of these phenomena at least,

in particular, for the effect on optical components. However, Muraca

[8] has shown that for some materials collected on a cold collector (298 K),

they evaporate slowly, and that the weight of material on the cold collector is never stable. This does not apply in general to specific polymeric bonding materials that may have spacecraft application because

of the fact that the volatile condensed material which is different from

the initial bulk source may polymerize or chemisorb to varying extents.

2.2 Optical Effects Mid to far infrared optical systems are of two basic types: (1) refractive -utilizing lenses, and (2) reflective-utilizing mirrors. Each

approach has its own advantages and disadvantages for each specific

sensing application. The performance of an IR system to accurately sense and detect radiation at required levels of sensitivities is a function of the individual system component performances which is shown

simply in Fig. 22. Item number one (Fig. 22) lists the transmission of IR radiation through windows or lens of Irtran , Germanium , or other

similar materials would be affected by a contaminant build up on the component's surface. Analytically, the transmission may be expressed by: A

T = [14]

1 + r2 rjL exp (-2 i<y^) where T is the transmittance and r-j^, r2,^1,^2 are the Fresnel coefficients at the riQ/n-^ and ni/n2 interfaces, (j =[2 t/x] n^ t cos (J, is the phase thickness of the film, and X is the radiation wavelength in vacuum (see Fig. 23) This expression assumes a parallel-sided, isotopic filmo For the case of reflective optics , the reflectance is ex- pressed by: r^ + r exp (-2 ia-^) For the real case of an

I ^ '^2-'"l (~2ioi)

1

1 61
absorbing medium must be expressed in complex form n-^ - ik where k is the extinction coefficient. From these expressions one can

see how the film absorbance characteristics, film thickness, and film

morphology play a controlling role. It is difficult however, to analytically compute the specific T and R values for various contaminant

films on surfaces because of the many unknown required input factors

Our solution to this problem is an experimental investigation. In addition to transmission and reflection property changes, the presence

of a contaminant film, which is probably not planor, will scatter radiation. For a transmitting surface the effect of scattering sites (Fig. 24a)

not only reduces further the amount of radiation emergent on opposite side

of the lens but also disperses the directionality. For the reflective

optics case the effect is enhanced, as shown in Fig. 24b. The reflected radiation will be scattered and absorbed going into the film as well as coming out of the film. This will cause a further reduction in signal intensity and directionality « These factors will present serious

stray radiation problems (off-axis inputs, increased detector crosstalk) as well as reductions in resolution and contrast capability for an

Infrared System.

To determine the effect of scattering one must know the film-surface characteristics, i.e. , film morphlogy, thickness, and index of refraction. These explicit factors have been found unavailable for specific materials in question, and it appears that an experimental

approach is best. The extent of this effect will be further evaluated

in the near future.,

2. 3 Experimental Program

To evaluate optical contamination due to polymeric material thermal vacuum effects two basic experiments were planned. The first would involve photomicroscopic analysis of contaminant build-up.

This would be done by first determining the size, thickness, and where

possible, composition of the condensed volatile material. Then predict the corresponding optical degradation using analytical analysis developed as a function of optics temperature, bonding material, its temperature, and vacuum level duration. The materials selected for evaluation include single and double component methyl-siloxanes of both the condensation and addition type. These experiments would be

done to evaluate the previously discussed transmission loss, scattering increases, and corresponding effect on resolution. It should be

pointed out that a computer simulation capability already exists at GE-RESD to predict IR system performance as a function of operational conditions (i.e. , selective adsorption, increased detector noise due

to thermal loads , etc. ).

62
Type two experiment will be in situ-infrared transmissionproperty

measurements. That is, the transmission properties of IR optical

materials (Irtran 4, 6, Germanium, and Chalcogenide glass) subjected

to a simulated contaminant environment. The transmission properties would be evaluated as a function of the following parameters

optics materials, bonding material, temperature of optics and bonding

material, and vacuum level duration » The performance of these ex- periments would establish directly the effect of contaminant build-up on the transmission properties of IR collection optics, thereby, clarifying the significance of proper materials selection.

To determine the transmission properties of volatile condensable

products a vacuum distillation system with a cold trap for collecting

VCP was fabricated and utilized. A diagram of this simple system is shown in Figo 25. Showing the glass assembly connected directly to the Kinney Vacuum System (Fig. 5). The flask connecting the polypieric material was placed in an oil bath to provide the 125 C temperature. The glass valving (Apiezan T lubricant was used) is such that the outgassed material must pass through the tubing connecting the tube in the LN„ bath and the flask. The VCP would therefore collect

in the base of the tube. The LN2 level was monitored by a controller

and the vacuum level by a gauge at the far end of the assembly. This

assembly has been used for Methyl phenyl siloxane RTV bonding material and Platinum vinyl siloxane. The procedures used consisted of placing 30 g's of the material chopped in 2 cm pieces (approximately

0. 8 cm thick) into a flask after various curing cycles. The first tests were performed with methyl siloxane RTV for both the accelerated

cure of 8 hours at 65 C, and the seven day room temperature cure. Each batch of at least 100 g's was mixed the same: dip coat clean

petri dishes in a primer and allow to dry, add a given weight of the bonding material along with 0. 1 percent curing agent (dibutyl tin dila- urate) and mix thoroughly and separated for different curing. For the Platinum vinyl siloxane it was mixed according to manufacturers di- rections, separated for a four hour (65 C) accerlerated cure, and

cured at room temperature for seven days.

The results of the initial tests (the program will consist of six to eight different polymeric materials) which remained under vacuum for 24 hours (~10-5 torr) showed a 0. 5 percent mass loss for the methyl

siloxane (accelerated cure). The collected condensable material when

first taken out of the LN2 bath was semi-solid, as it was allowed to warm up (isolated from rest of the system) two liquid phases were

observed. Later IR analysis showed this to be water to a large extent which was attributed to the fiUer used in this RTV. For the vinyl

siloxane (accelerated cure) and seven day room temperature, under

similar test conditions , smaller amounts of condensable products were

collected. There was not a two phase liquid visible. Each sample

63
was analyzed by Infrared Spectroscopy (Perkin Elmer) and the corresponding spectra are shown in Fig. 26 and 27. The spectra was obtained with the samples at a temperature of 25C. Plans are being

made to adapt the experiment to perform low-temperature IR spectral

effect of contaminants absorbing radiation on cooled optical window.

The spectra in Fig. 26 shows the large selective absorption for the 9 to 10 micron, 12 to 13, and 14 to 15 micron bands. The samples were

run in the pure form with cell absorbance spectrally subtracted out of the results. Fig. 27 shows a similar spectrum in the original

absorbance data form for the vinyl siloxance. However, the thickness

of film are not the same with the latter being (one half the former)

approximately 1 microno Selective absorption of this type if realized on an operational spaceborne IR optical system would cause serious

loss of sensing capability; for example, for Earth's limb measurements the resulting IR output spectra would show false absorptions in the peaks previously discussed.

To assess how significant this selective absorption might be, the

size, thickness, and surface state of condensed off-gassed polymeric

products must be determined. This would have to be performed in a space simulated ultra- clean vacuum facility. The requirements for such an experimental facility are the following: (1) multi-test capability; (2) thermal control and monitoring (77K to 673K); (3) elimination of synergistic vacuum system contamination; (4) high vacuum

pumping to maintain simulation requirements during initial contaminant source outgassing; (5) bakeable for rapid clean-up; (6) ultra- high vacuum capability (cryopanneled) ; (7) provision for in-situ

optical measurements. The design of a system to meet these requirements is shown in Fig, 28. It is a stainless steel chamber approximately 46cm X 46cm x 25.4cm. The rotatable optical specimen mount

(which is thermally controlled by strip heater underneath the plate and

cryogenics on top of the plate) accommodates four test optical materials. Each specimen can be rotated into the visible or IR optical

train as the specific experiment requires. This platform is removable from the chamber and will fit through the entrance port. A detailed view of the optical specimen mount is shown in Fig. 29. Each

of the four holders is cryogenically cooled by boring out the center of each holder's sides (as shown). The LN2 is supplied by a flexible stainless steel hose that will rotate as the whole table is turned. The

contaminant source configuration is also graphically shown. The

materials will be easy to change as it is placed in a removable boat

for each specific experiment.

The walls are cryopaneled (LN2) with the fin-type panel which has

been shown to have the best trapping efficiency. The outside walls

have strip heaters attached to facilitate rapid bake-out up to

64
The thermal excursions that occur at the various locations in the as- sembly are monitored by Copper-Constantan thermocouples.

There are various support feedthroughs to provide LN2, voltage

for heaters, and thermocouple connections. The major vacuum pumping will be done by an 30 l/s ion pump (after the base pressure has

reached 10"^ torr via a diffusion pump). The mating of the test as- sembly to main vacuum system (see Fig. 30) is by a pneumatic safety valve. This is required between chambers as well as between the test chamber and ion pump to effectively isolate the components in the event of a power failure. The major vacuum systems consists of a 2000 l/s 11 in. diffusionpump, cryotrapped; cryopanelled (LN2) cham

ber walls, circular heaters for the main chamber exterior, and a 19//s roughing pump. The pressure level is measured by the ion

pump controller and by ionization gauges at various locations. The

performance of this system with test chamber isolated and cryocooled walls after a moderate bake-out of 250 C for four hours is in the high 10~8 torr range. The ion pumped test assembly will achieve

pressures in the low 10" 8 torr range with moderate bake-out.

The in-situ optical measurements can be performed for either the

visible or IR range of the electromagnetic spectrum. The visible

optics investigations are possible through metal-gasketed viewports

which are aligned across from each other, with the test specimen

being positionable in the viewport field-of-view. The in-situ surface

photography utilizes a stereo microscope which is focused outside the chamber on the optical component being tested. A camera is attached

to the microscope for a permanent records The visible light source

will be a mercury arc lamp (500 watt) which provides sufficient il- lumination for low power magnitications . Difficulty has been experienced in gaining sufficient resolution for film thickness determination

for the long working distances ~ 6 inches). This problem is currently being addressed to enable in-situ thickness determination as well

as surface morphology definition. As an intermediate approach,

microphotographys were taken under ambient conditions of condensed

material from methyl siloxane based polymer using the set-up as shown in Fig. 25 by placing glass slide into the tube (discussed in Appendix A). The infrared observations will use specially designed IR viewports (Irtran 6). The radiation source for the infrared (yet to be acquired) will be a precision blackbody radiator, with the detector

(also yet to be acquired) being a thermopile.

The vacuum test assembly (Fig. 31) in its present state of development is operational for the surface photography investigations.

The infrared transmission experiments will be commenced shortly

upon receipt of the IR source and detector.

65
2.4 Conclusions

An experimental program has been defined and developed to in- vestigate the optical degradation effects due to the thermal vacuum

performance of polymeric materials that are used on spacecraft programs o It has been shown analjrtically that the presence of cold

(cryogenic) collection surfaces do enhance the contaminant build-up.

In addition, a major degrading effect may be due to the far infrared

absorption properties of condensed volatile out-gassed materials.

The scattering effects are predicted to cause serious off-axis rejection problems as well as loss of detection sensitivity. From these

preliminary investigations, it is apparent that care must be exercised

in the selection of a far infrared space-borne optical subsystem bonding materials as the degradation of the resulting contaminating film

may severely decrease the signal/noise ratio for an operational opti- cal sensing subsystem and thereby minimize the overall system

sensitivity o ACKNOWLEDGEMENTS

The author wishes to acknowledge the assistance of Dr. A.E.

Binks of the Environmental Sciences Laboratory in the area of polymer

chemistry and Dr, E.L. G. Odell also of the Environmental Sciences

Laboratory in the area of optical physics, Frank Pancoast, Vacuum

Technologist in the area of vacuum system fabrication, performance

and evaluation, and B. Schein for graphical displays.

APPENDIX A

A. 1 Preliminary Photomicroscopic Investigation of Contaminate

Build-up

In the contamination collection vacuum assembly, glass slides were placed in the bottom of the collection tube which was submerged

in LN20 After 24 hours with the polymer material (methyl siloxane)

at 125 C at 10~5 torr pressure, the collection tube with the enclosed

slide was isolated and removed from the LN2. The sample was al- lowed to warm up under reduced pressure to minimize the ambient

vapor condensation effect when it was removed for microscopic in- vestigations . The slide was then removed and analyzed under a reflected light (metallurgical) microscope. This microscope was operated in the 10 OX to 500X magnafication range with the attached camera

assembly (Bausch-Lomb L Camera and Polaroid Film Holder)

66
operating in the 50X to 200X range o The light source was a mercury

arc lamp filtered to provide good color match for the film. Some of the first photos taken are shown in Fig, A-1. They show the variation in contaminate morphology. The size of the droplets was determined by a calibration with a precision stage micrometer and the corresponding linear distances are recorded on the figure. The thick- ness of the condensed material was determined by interference techniques. That is for an index of refraction for the film nf such that

i^air ^f <iiglass' 2 tn = (m + l/2) X for a minimum, 2 tn = mXfor

a maximum m = 0, 1, 2, 3, t = thickness. In Plate a) at least 8 fringes can be counted on the large droplet assuming nf =

1.4; ngiass - 1*5 and X= 0o45 microns we get t = 1. 3 microns with

its linear size being approximately 100 microns. Clearly this is a

large formation and was not observed in general. The majority of the cases were as shown in c) an agglomerate of smaller particles. It was observed that as the substrate was allowed to ramain at ambient,

there was a tendency for the droplet to coalesce; however, a continu- ous film was not observed in the cases of this particular RTV.

REFERENCES

[1] Enlow, D. , "A Physical Study of Microbe Surface Interactions",

Re-entry and Environmental Systems Division, General Electric

Company, Technical Memorandum 8126-15, October 1967.

[2] Enlow, D. and Cheater, D. , "Experimental Investigations of Microbe Surface Interactions", Re-entry and Environmental

Systems Division, General Electric Company, Technical Memorandum 9821-006, February 1970.

[3] Vacuum Science and Engineering, CM. Van Atta, McGraw HiU,

New York, p. 328-334,1965.

[4] Roberts, L. , and J. C. South, Jr. , "Comments on Exhaust Flow

Fluid and Surface Impingement", ALAA Journal, Volume2, #5,

p. 971-973, May 1964.

[5] Eastman, D. , and Radtke, L. , "Flow Field of an Exhaust Plume

Impinging on a Simulated Lunar Surface", AIAA Journal, Vol- ume 1, #6, p. 1430-1431.

[6] Sibulkin, M. , and Gallaher, W.H., "Far-Field Approximation

for a NozzleBc hausting into a Vacuum", AIAA Journal, Volume

1, #6, p. 1452-1453.

67
[7] Equations, Tables and Charts for Compressible Flow, NASAAMES Report 1135, 1953.

[8] Muraca, R.F., and Whittick, J.S., "Polymers for Spacecraft

Hardware", Final Report, Stanford Research Institute, SRI

Project #ASD 5046, September 1967.

[9] Whipple, C.L. , and Curtindale, E.G., "Performance of Silicone

Materials in Thermal Vacuum Environments", Trans. 10th National Vacuum Symposium, 1963, p. 63-68, MacMillan Company, New York.

[10] Poehlmann, H.C., "Vacuum Weight- Loss and Contamination

Tests of Some Materials for Space Application", Proceedings of the Fourth International Vacuum Congress, 1968, p. 809-813.

[11] Vacuum Science and Engineering, CM. VanAtta, McGraw Hill Publ. , 1965, p. 410. "

[12] "ATM Contamination Study Task IV", Final Report Ball

Brothers Research Corporation, Report F67-06, March 1967.

[13] The Physical Basis of Ultrahigh Vacuum, P. A. Redhead, J. P. Hobson, E.V. Kornelsen; Chapman and Hall Publ. , London,

1968, p. 59-80.

[14] Thin Film Phenomena , K.L. Chopra, McGraw Hill Publ. , New

York, 1969, p. 723-725.

68

Fig. 8. Shock Test Results

70
71
Figo 15. Fluid Flow: Percent Removed as a Fimction

of Duration of Pressure Excursion (^)

DEFINE AND

ANALYTICALLY

INVESTIGATE PHENOMENA

OF OPTICAL DEGRADATION

LABORATORY

REQUIREMENTS AND DEVELOPMENT FOR ULTRA-CLEAN SPACE SIMULATON VACUUM

SYSTEM

VISIBLE OPTICS

INVESTIGATION OF CONTAMINATION "FILM" (SURFACE MICROPHOTOGRAPHY)

INFRARED OPTICS INVESTIGATION OF

TRANSMISSION PROPERTIES

STUDY CONCLUSIONS RECOMMENDATIONS

FOR MINIMIZING OPTICAL CONTAMINATION

Fig. 16. Approach to Optical Contamination

Investigation: Flow Diagram

72
73
74
75



Paper No. 5

EXTERNAL SPACECRAFT CONTAMINATION MODELING

AND COUNTERMEASURES

Theodore Baurer"^, M. H. Bortner"^, Irwin M. Pikus"^, and

Arnold M. Cooper^

REFERENCE: Baurer, Theodore, Bortner, M. H. , Pikus,

Irwin M. , and Cooper, Arnold M. , "External Spacecraft Contamination Modeling and Countermeasures, " ASTM/IES/AIAA Space

Simulation Conference, 14-16 September 1970.

ABSTRACT: The problem of spacecraft external contamination is a

function of both the spacecraft and the receptors, in relation to the natures and quaatities of contaminants involved, i.e. , the problem

depends on both design and mission. The relevant phenomenology

may be conceptualized in a model which assumes that a mass-transfer

mechanism is operative, including sources, transport modes, and

sinks. Typical sources include engine debris, jettisoned wastes,

outgassing, and cabin leakage. Gaseous contaminant transport is based on the complex interaction of gas-surface interaction

potentials and hydrodynamic, electrostatic, and magnetic fields. Condensed-phase materials are subject to hydrodynamic coupling,

induced polarization, and electrostatic field effects. The sink or

deposition mechanism assumes differential charge development on

conducting (e.g. vehicle skin) and dielectric (e. g. windows, lenses)

external surfaces , such as to cause preferential attraction to the

latter. The net electrostatic potentials at the dielectric surfaces cause charge neutralization of impinging ions, followed by bonding,

scattering, and desorption. Impacting particulates adhere or

scatter at the dielectric surface depending on the relative magnitudes

of local surface potential, strength of induced dipoles, and particle

^Space Division, General Electric Company, Philadelphia, Pa. Research Dept. , Grumman Aerospace Corp. , Bethpage, N. Y.

79
velocity vectors normal to the surface. Flowing and adhering

contaminants continually cover the surface, preserving its dielectric nature. This model suggests the applicability of at least three

complementary approaches to the problem of counteracting contamination, viz. , passive prevention of the appearance of contaminants,

active removal measures during the transport of contaminants which

do appear, and means of counteracting or discounting the effects of contaminants which complete their transit to sensitive surfaces.

KEY WORDS: contamination, spacecraft, mass transfer

1: INTRODUCTION

The external environmental contamination of manned spacecraft

has been recognized as a problem since the actual sighting of contaminant particles was first reported by the early Mercury

astronauts (Refs. 1, 2), ^ and the origin of these particles as effluents from the vehicle was then shortly confirmed (Ref. 3). Subsequently,

similar phenomena were observed photographically on unmanned

rockets (Ref. 4) and satellites (Ref. 5). Gaseous environmental

contamination was observed as well (Ref. 6). The implications of a self- generated contaminated environment

immediately adjacent to manned or unmanned spacecraft, in terms

of its effects upon the ability of men or instruments to perform

scientific missions in space, were explored analytically three years

ago (Refs, 7-9). However, in the absence of definitive experimental

verification, the results of these early analyses were inconclusive.

Indeed, a specific item of experience brought into these considerations (Ref. 7), viz. , the reported inability of the astronauts to see

stars in the daytime, was attributed elsewhere (Ref. 10) to factors

having more to do with human ocular adaptability to varying background luminance than with any effects of the spacecraft environment.

The whole problem came into fresh focus with the well-publicized window contamination observed on Gemini spacecraft (Ref. 11). Specific experiments (Ref. 12) carried out in conjunction with the

later Gemini missions, and involving exposures to the space environments of test coupons, reaffirmed the gravity of the situation. Still more recent experience under the Apollo program (Ref. 13) has

3The numbers in parentheses refer to the list of references appended

to this paper.

m
tended to reinforce the general concern over contamination.

It is now well recognized that, whereas certain manned and

unmanned satellite-based experiments projected for this and later decades have as their primary objective the observation and measure- ment of radiation emanating from solar, stellar, and other nonterrestrial sources, unimpeded by atmospheric absorption, this aim now faces the possibility of severe compromise. Specific

dangers arising from the presence of external contaminants include:

(1) the development about the experimental vehicle of its own

localized "atmosphere" comprising off-gas and cabin leakage products,

human and fuel cell wastes, and engine effluent particles and gases;

(2) the deposition and adhesion of both particulate matter and gaseous

condensates from this "atmosphere" upon sensitive optical, thermalcontrol, solar cell, and other external surfaces; and (3) the interference, through light absorption and scattering, with signal

reception aboard the spacecraft, owing to the passage of gaseous and

particulate contaminants across optical paths. These effects, if permitted to proceed unchecked, would tend to destroy the very basis

of the several experimental packages now projected, by virtue of both

the presence of undetermined foreign matter in the respective

optical fields, and the actual fouling of the receptor elements themselves. The present concern is with the reduction of the external spacecraft contamination phenomenology to a model which assumes that the mechanism is essentially one of mass transfer, including sources,

transport modes, and sinks. This model, which is based insofar

as possible upon information available from the NASA manned spaceflight program, serves in turn as the basis for the consideration of countercontamination concepts.

2: THE MODEL: DEVELOPMENT

As indicated above, the GE/SSL effort to model the external

contamination phenomenology of spacecraft has been based on the

simple concept that the entire process is an example of classical mass transfer, under highly specialized conditions. The practical

approach to modeling, based on this concept, has comprised preliminary definition of the specific sources of contaminant materials

insofar as they are known, followed by the attempted description of probable transport modes, and identification of the sink or deposition

mechanisms, while relating the totality of these efforts to actual experience, as derived from chemical analyses of contaminant

deposits returned from space.

81
2. 1: CHEMICAL COMPOSITION AND SOURCES OF CONTAMINANTS

2.1.1: GEMINI

The results of semi- quantitative emission-spectrographic

chemical analyses (Ref. 13) of window contaminant material from

manned Gemini (GT4-7, 9-12) space capsules were examined (Ref.

14), and an estimate of the percent of each element was derived

therefrom. These estimates are presented in Table 1. A number of possible sources of Gemini window contamination were examined

critically (Ref. 14), and their probable importance was evaluated.

This information is also contained in Table 1. While several relatively unimportant constituents of the window

deposits may not be fully accounted for, such elements contribute no more than 2. 5% by mass to the overall contamination analyses,

while the others are attributed in Table 1 to just five major sources,

viz. , the throat ablator, propellant impurities, waste dumps,

window seals, and the sea. Needless to say, while the sea may be

accountable as a partial source of the deposit material as analyzed,

it certainly is not a source of anything in space.

2.1.2: APOLLO

The available analytical data relating to contaminant composition on the Apollo windows were compared (Ref. 14) with the corresponding information elucidated with respect to the Gemini missions.

In this effort, analytical data were obtained for several unmanned

flights (I-IV and VI), not all of which were orbital missions, and one manned flight (XI). Table 2 presents a summary of the analytical data from the

flights enumerated above, along with presumptive sources, as deduced by NASA/MSC persomiel (Ref. 13), which apply only to the

five unmanned flights. These sources are of interest, inasmuch

as they coincide, to a degree, with the conclusions expressed above,

with respect to Gemini. One additional major source is listed, viz. , the launch escape system (LES), and the launch escape tower

(LET) components, which had not been considered in evaluating the Gemini deposits.

The Apollo XI analyses in Table 2 roughly parallel those of Apollo II-IV, which were unmanned orbiting or high-heat reentry

vehicles. Thus, the residual contaminants found on the windows of both manned and unmanned orbiting spacecraft may have come from

the same sources. However, those contaminants which may have

impacted the windows and bounced away, or which may have washed

away on splashdown, could have had other sources, e.g. waste

dumps, propellant impurities.

82
An important distinction which must be drawn in comparing

Gemini and Apollo is to be found in the different number (two and

five, respectively) and disposition of windows in the two vehicles.

The Apollo window configuration comprises the hatch window situated

in the middle, two flanking rendezvous windows, and two outlying

side windows. Of these, the hatch and side windows are flush with

the conical vehicle surface and are reputedly the most vulnerable

to contamination in flight; the rendezvous windows are inset facing

the apex of the cone, i.e. , in a forv/ard flight direction, and have

remained relatively clean.

2.1.3: PASSIVE PREVENTIVE MEASURES AT THE SOURCE

The above discussion and those of Ref. 14 relate to a variety of conceivable contaminant sources, the importance of which is a function

of both the operational characteristics of a spacecraft and its mdssion.

For example, where deposition largely of a silicaceous nature

(Table 1) was markedly noted upon the Gemini windows (Ref. 11) and

upon test coupons (Ref. 12), it appears that such gross sources as the phenolic-Refrasil ablatively- cooled attitude- control thrusters

and silicone-based window sealants were among the probable primary

sources of the visual interference noted by the astronauts. Those

lesser sources such as surface outgassing and cabin leakage which

may have produced as much as 2-3% of the contaminant mass were

relatively less important within the context of these missions. On

the other hand, the optical surfaces associated with projected flights of space experiments such as the Apollo Telescope Mount (ATM)

cluster are likely to be much more sensitive then were the Gemini

windows to minor forms of contamination. Therefore, every conceivable source of foreign matter must be regarded, a priori, as a potential threat to the latter type of mission.

It is tempting, having identified the principal contaminants, and

their major and minor sources, to cast about in an effort to find means of suppressing the contaminants at the source by the application

of control measures which are essentially passive in nature. Table 3 lists the passive measures which might be utilized in efforts to control

contamination in space at the source. Some of the remedies suggested in Table 3 appear to be relatively easy to implement, e. g. cleaning up the engine fuels, modifying the waste handling procedures.

Others would require major vehicle redesigns, e.g. substituting

radiatively for ablatively cooled engines, retaining window and lens coverage until well after LET jettison. Still others require extensive

materials searches, evaluations, and trade-off studies, e.g. for

83
new window sealants and for the several minor external surfaces

listed.In the case of the window sealants, a change in procedure, viz.

the introduction of a pre-launch bakeout, and the substitution of a newer, less volatile silicone product, have been credited (Refs. 13,

16) with a partial improvement in the contaminant picture as regards

silicon. However, still further improvement is possible, and it has

been argued (Ref . 17) that the utilization of Viton, a synthetic

fluorinated heat-molded elastomer, in place of the silicones, could

be expected to eliminate or strongly reduce this aspect of the contamination problem. This is based on cumulative experience with

Viton as a laboratory vacuum seal, and as a test substance in materials degradation experiments (Ref. 17). As for the concept of passive controls generally, and notwithstanding the worth of some of the remedial approaches suggested in Table 3, or discussed above in relation to window sealants, it does seem evident that the number of conceivable sources, for unwanted

matter capable of condensing on windows and optical surfaces, is so

great, and their variety is so diverse, as virtually to preclude the successful application of passive control techniques at every single source. Accordingly, it is suggested that no reliance be placed at this time on passive control as an overall remedial measure or cureall for the contamination problem. At best, this type of technique

may be relied upon to reduce somewhat the level of contaminant

emissions at the source, but not necessarily to the point of threshold

tolerance by experimental instruments whose protection is of utmost

importance to a given space mission.

2.2: TRANSPORT OF CONTAMINANTS

The transport of gaseous contaminant matter is dependent upon

the superposition of thermal and photoionization processes upon the ambient flowfields, followed by the complex interaction of gas-surface

interaction potentials and hydrodynamic , electrostatic, and magnetic

field effects. Particulate matter is dependent for transport on phase

equilibration followed by hydrodynamic coupling, induced polarization,

and electrostatic field effects. Thus, the number and variety of forces involved are such that the complete quantitative definition of these processes is very difficult. However, certain aspects of the transport mechanisms for gaseous and particulate contaminants have

been examined individually (Ref. 14). They include the ionization and

field interactions of gases, and the two-phase flow and interactions

of solids, and will be summarized here.

84
2.2.1: IONIZATION OF GASEOUS METAL ATOMS

The transport of gaseous contaminants, e.g. metal vapors, is dependent upon their ionization prior to electrostatic field interaction

with the charged surface of the vehicle. A variety of ionization

mechanisms may be postulated with respect to the metallic constituents found in Gemini and Apollo window deposits. The most

obvious of these mechanisms include:

a. Direct solar photoionization, e.g. M +hv > M +e, where M represents a metal atom and hi/ a photon of appropriate

energy.

b. Thermal ionization, e.g. in the course of OAMS or other

engine combustion processes.

c. Indirect solar photoionization, e.g. via metastable- state formation and subsequent solar-flux photoionization:

M + A E > M* followed by

M* + hv (E. - AE) *M +e, where

M*is the metastable state mto which quantities of M are first pumped, and having an energy AE above that of the ground-state M;

or via metastable -state formation and subsequent autoionization:

M + AE ^ M* as above, followed by

M*+ (E. - AE + §E) > W^, where

5 E is a small increment of energy representing the difference in ionization limits between two different state series available to M,

and M** is M in the limiting state for one (the higher) of the two

series, then spontaneously, ^

M^t^ >M^+ e + 5E.

d. Metallic vaporization as the oxide, either in the original

emission, via oxidation in the vehicle environment, or on the surface after deposition; followed by ionization of the gaseous oxide, e.g. MO + hv *• MO + e, or MO + e '•MO + e + e.

It was concluded (Ref. 14) that, pending further experimental

investigation, none of these processes appear to be important, and

that the metallic constituents reported in spacecraft window deposits

probably arrived in condensed form, as solid particles or liquid

droplets, and not in the gas phase.

2.2.2: FIELD INTERACTIONS RELEVANT TO THE TRANSPORT OF

IONIZED GASEOUS CONTAMINANTS

Significant quantities of ionized gases are nevertheless present

85
in the vehicle flowfield, e.g. ambient ionic species such as O . Aside from purely hydrodynamic considerations such as were treated

in Ref. 8, other forces appear to play a possibly significant role in determining the transport modes of gaseous contaminant ions. These

include the forces arising from ionic interactions with local electrostatic fields, and to a lesser extent with the earth's magnetic field. The development of a surface electrostatic charge upon conducting surfaces of the vehicle (Refs. 18-27) may arise from:

a. The transfer of positive charge to the surface, e.g. through

trapping of ions and by Auger emission of electrons;

b. Electron ejection through the absorption by a sunlit surface

of sufficientl}^ energetic photons; and

c. Subsidiary mechanisms, e.g. motion of the vehicle through

the earth's magnetic field which gives rise to an effective electric

field in the vehicle's coordinate system, and the operation of electrical equipment inside the vehicle which can produce a change

of surface charge state if there is a conducting path to the surface.

The field produced about the vehicle by the charge upon its surface is weakened and distorted by the flowing plasma environment.

The details of this problem have been considered for a number of years (Refs. 28, 29). Even assuming that the charge on the vehicle

is uniform or that the vehicle is an equipotential surface, the field in the vicinity of the vehicle is not easily described. The principal

features of these results which seem applicable to the problem are

the following: The negative charge on the front of the vehicle is well

shielded by the incoming positive ions. Toward the rear, however,

the potential is much less effectively screened. Furthermore, there

is significant distortion of the field caused by the effect of the earth's

magnetic field on the surface charge distribution.

It was concluded (Ref. 14) that, with respect to the transport of contaminant gaseous ions through an electrostatic mechanism:

a. Significant charge differences can be established between

dielectric and conducting surfaces.

b. Windows facing out of the flow, especially those facing rear- ward, probably achieve the most negative potentials on the spacecraft.

c. Thruster firing not only supplies new contaminant material

but also changes the charge state of the vehicle at least during the

firing period.

2.2.3: THE TRANSPORT OF CONDENSED-PHASE CONTAMINANTS

The forces under which solid and liquid contaminant material

acts include: 86
a. Those due to the ambient conditions, e.g. aerodynamic drag,

radiation pressure, solar wind, and interactions with the space

magneto -plasma (Refs. 8, 30). b. Those due to plume interactions, e.g. two-phase coupling

(Ref. 31).

c. Those due to the presence of the vehicle, e.g. the electrostatic force arising from the existence of charge upon the vehicle

surface, and charged contaminant material.

The ambient forces were found (Ref. 14) to be negligible in comparison with the other effects considered.

The problem of purely gaseous plume interactions is not understood sufficientlj^ w^ell for the purposes of impingement predictions.

A purely continuum formulation of plume expansion would provide a

solution which would obtain in the vicinity of the gas exit port.

However, far from the exit any continuum formulation breaks down

since the mean free path of the molecules in the plume increases

downstream, while the length scale associated with the spatial

gradients of the flowfield, as predicted b}^ the continuum solution,

is scaled by the exit diameter. That is to say, far from the source,

situations exist for which the local mean free path is comparable

in magnitude to the scale of the local flow gradients. Then the molecular distribution function can no longer be near equilibrium,

and any continuum formulation breaks down, necessitating recourse

to a kinetic formulation.

In recent years there has been substantial progress in understanding rarefaction effects in underexpanded free jets. However,

this progress has been concentrated in the vicinity of the plume

axis where the basic symmetry of the flow enables an accurate modeling of the plume by a supersonic spherical source expansion

(Ref. 32). Utilizing the spherical source model, numerous theoretical

studies (Refs. 33-35) have provided predictions of the nature and

location of the rarefaction effects in the centerline region. These

studies have been for pure monatomic and polyatomic gases and for binary gas mixtures.

The problem of predicting particle migration in the plume

flowfield is more formidable in that the general case presents an extremely complicated coupled two-phase flow. The nature of the resultant flowfield is dependent upon a number of fundamental

parameters, e.g. the size distribution of particles, the concentration

field at the exit, and the average mass density ratio of particles

to gas. This last parameter determines the nature and degree of coupling between gas and the particles in an overall sense. The details

of particle size distribution are necessary to determine the distribution

of this coupling over the particle size distribution. There are basically two coupling situations which are important

87
in gas-particle plumes, viz. , strong and weak coupling. Strong

coupling occurs when the mass densities of the two phases are comparable and the resulting flow is such that the gas and particle

characteristics are dependent upon each other. Weak coupling occurs when the particle loading is very low. In this case the gas- phase characteristics are independent of the particle phase, but the particle-phase characteristics are determined by interaction with

the gas phase. The weakly coupled case is, of course, the more

convenient for calculation since the gas flowfield can be determined

independently and can then be utilized to produce the required

particle field. Finally, a preliminary examination of electrostatic effects between a vehicle and a charged droplet was undertaken (Ref. 14)

because of the importance of droplets in window contamination on

spacecraft. The mechanism proposed for the interaction involved

summing the coulombic forces between charges on the vehicle and

the droplet, and the forces between the vehicle and an induced dipole

moment on the droplet. In order to compute these forces it is necessary to know the vehicle charge, and the charge and polarizability of the droplet. The latter problem is one of considerable

difficulty since the chemical com_position of the droplet is essentially

unknown, the physical state is unknown, the equilibrium charge

on a small droplet cannot be extrapolated from the calculation of vehicle charge because of the differences in composition and size,

and the effective polarizabilities of mixed-phase (solid and liquid)

droplets of various geometries and compositions are not available nor easily computed. Model calculations of the interaction can, however, be made by first assuming values of the necessary properties.

2.2.4: ACTIVE PREVENTIVE MEASURES IN TRANSPORT

It is implicit in the concept of ionized gases, polarized liquids,

and charged solids traveling in predictable trajectories defined by

known or readily estimated force fields, that the magnitudes and

directions of artificially imposed counterforces which would

comprise the basis of active countercontamination measures are at least calculable. Whether such measures are also realizable in terms of actual devices has yet to be determined. However, despite

the lack of demonstrated feasibility of such devices at this time,

it is nevertheless possible to indicate several general principles upon

which they might be based, such as to implement the general idea

of force-counterforce measures. These principles include:

outright expulsion from the total spacecraft environment, deflection

88
to nonsensitive surfaces comprising integral parts of the vehicle

structure, deflection to sacrificial surfaces deliberately added to the spacecraft, and trapping upon grids or other special filtering

devices imposed across the contaminant paths. Each of these

approaches offers the potential of providing a credible means of countercontaminant control. Their further analytical and experimental development is strongly recommended.

2.3: DEPOSITION OF CONTAMINANTS

The sink or deposition mechanism assumes differential charge

development, with the resultant generation of surface currents, on

different external surfaces of the vehicle, i.e. , the conducting

vehicle skin vs. the essentially dielectric functional areas, e.g. windows, lenses, thermal- control coatings. The resulting net

electrostatic potentials at the dielectric surfaces may cause charge

neutralization of impinging gaseous ions, followed alternatively by

physical and chemical bonding, scattering, and desorption phenomena.

Charged particles and polarized droplets impacting the vehicle adhere

or scatter depending on the relative magnitudes of the local surface

potential, the strength of induced dipoles, and the particle velocity , vectors normal to the surface. Some of the factors affecting this phenomenology were examined (Ref. 14).

2.3.1: CONTAMINANT-SUE FACE ENCOUNTERS

Contaminant material, in atomic or molecular form,

originating at the spacecraft, may encounter the vehicle surface even in the absence of a long-range attractive force. Its speed at encounter depends upon its initial thermodynamic state and the degree to which it has thermalized in the ambient medium.

Gas-surface interactions have been investigated (Refs. 36-38)

where the surface was metallic and the gas was a monatomic or diatomic species. A marked dependence of capture probability upon

surface temperature has been found, in addition to the expected

dependence on molecular velocity (Refs. 36, 38). Studies of the trapping of gas on an impure surface have indicated that the coupling

between impurity and substrate is very important. For certain

combinations of molecule and surface species the interaction

potential at encounter is of dominant importance (Ref. 37). In such

chemisorption interactions it frequently has been found that there

exist a number of binding states between molecule and surface.

Sometimes the adsorbed molecule has sufficient mobility to travel

89
along the surface until it meets a location even more favorable

energetically. In certain cases it is expected that the trapped

molecule will react at the surface, either dissociating or combining

with other surface species, to form new structures. The impingement of contaminant particles on spacecraft

surfaces is predicted from considerations of contaminant transport,

and the prediction is confirmed by observations. Particles generally are expected to inpiige at relatively low speeds. Thus, the trapping

of particles is highly probable. Solid particles, even at low speeds, can produce surface damage, i.e. , defects in surface structure,

before coming to rest. Liquid droplets may spread out over a considerable portion of the surface or may remain as globules. This

is determined by the relative importance of surface tension of the droplet and forces of interaction with the surface. The tendency for small droplets to coalesce into larger ones, and the likely distri- bution of droplet sizes on the surface, are also determined by these

factors.Interaction modes between trapped particulates and the surface

material include chemical bonding and surface alloying. However,

even ui the absence of these very strong interactions, other modes

of interaction exist. In particular, Vander Waals forces, which are

relatively long-range on an atomic scale, and electrostatic (or

double-layer) forces are important (Ref. 39). These can give rise

to trapping of the solid particles and to the viscous creeping of

liquid droplets, leading to their gradual spread over the surface.

2.3.2: CONTAMINANT-SURFACE BONDING

It is necessary to take account here of the competing forces

which may cause on the one hand, bonding or sticking, or alternatively,

reflection or scattering, of positive ions approaching a negativelycharged dielectric surface. A rigorous solution of the problem is precluded at this time. Furthermore, virtually all the previous

efforts in this area have been concerned with gas-surface bonding.

Therefore the subject of bonding between condensed-phase contaminants and spacecraft surfaces has not been considered, beyond the few remarks on interaction modes which concluded the previous

paragraph.

The interaction of fast gaseous species, both neutral and

charged, impinging on solid surfaces, has been a matter of concern

for some years (Refs. 38, 40-43). The experimental approach has

been based largely on the results of molecular and atomic beam

testing. A second area of phenomenology bearing on the gas-surface

interaction problem is that arising from studies of sorption, both

90
physical and chemical (Refs. 38,44). However, much of the work in this area deals essentially with equilibrated systems, and hence is of limited applicability to the present problem. On the other hand,

the results of these investigations do shed some light on the nature

of bound states resulting from djmamic encounters. A number of analytical approaches to the problem have been

reported as well. For example, one model assumed simple,

classical "hard-cube" scattering (Ref. 45), which in one version

(Ref. 38, p. 392) was treated in conjunction with an increasingly

detailed lattice model of the solid, such as to project possible

generalizations to quantum interaction models. This model was

succeeded by a more sophisticated concept based on exponential

interaction between a gas atom and a single one-dimensional

oscillator, at a flat surface assumed to consist of "soft-cube" atoms

(Ref. 46).

Another classical approach was based on lattice dynamics.

This treated a two-dimensional independent - oscillator lattice,

with harmonic potentials connecting the lattice sites, and Morsetype potential functions operating between the incident gaseous atom

and each lattice atom (Ref. 47). The model was subsequently

expanded to a three-dimensional independent-oscillator lattice; nine

movable lattice sites were each connected both to the incident gas atom by a Morse potential, and to fixed lattice points by harmonic

springs (Ref. 48).

One further modeling concept (Ref. 37) considers, as a first approximation, a triangular interaction potential with impulsive

energy transfer in a rigid-spheres high-speed head-on collision, after the attractive potential has acted. Here, the target atom is coupled to the solid through an effective spring constant, and the

lattice model is assumed to be two-dimensional in depth, i.e.

straight down into the solid, rather than in breadth, i.e. , across the surface of the solid, as in other models (cf. Ref. 47). Threedimensionalization was contemplated in a projected extension of this study (Ref. 37), which unfortunately was not completed. This was to have included considerations of off-center, as well as head-on,

collisions. With regard to all the lines of investigation reviewed above,

they all are concerned with crystalline solid surfaces. Thus their

applicability to interaction systems involving glassy surfaces is very much open to question. In the present area of concern, the stress is upon the latter type of surface, as exemplified primarily

by the spacecraft windows. A second significant limitation of the current state of the art is the emphasis upon neutral gaseous atoms

as the colliding species. The problem presently under consideration

involves the impingement of gaseous ions. Thus most of the previous

91
experimental and analytical work in gas-solid interactions is only

marginally applicable to the current area of interest.

2.3.3: REMEDIAL MEASURES AT THE POINT OF DEPOSITION

By contrast with the passive and active preventive measures

suggested above for implementation, respectively, at the source and

in transport, many of the countercontamination procedures which

might be brought into operation at the point of deposition would more

likely be compensatory or remedial in nature. For example, a totally permissive approach to countercontamination would be one in which the amount of deposit accumulated upon

a given window or lens was routinely monitored up to some limit of acceptable performance loss, after which remedial procedures would

be undertaken. These might include cleaning of the affected surfaces,

either by some mechanism controlled from within the spacecraft, e.g. using an external N - methanol jet (Ref. 16), or during extra- vehicular activity (EVA) by an astronaut. Or contaminated optical

components might be replaced by retracting the instrument within the vehicle for overhaul. By the same token, a temporary flux of contaminant gases passing through a given optical path might be

deemed tolerable if the extent of light absorption and/or scattering

caused thereby did not exceed some acceptable noise level. If such

a limit were exceeded, observation via the affected light path would

be shut down for the duration of the contamination episode.

On the other hand, a preventive-permissive approach would

involve advance knowledge of major contaminating events, e.g. engine firings, waste dumps, plus advance calculation of the resulting

contaminant trajectories. Then all sensitive surfaces likely to be

affected could be shut down temporarily, i.e. , protected by retractable covers, or drawn within protective housings or the spacecraft itself. This would last until normal operations could be resumed safely,

according to the indications of strategically emplaced monitoring

instruments. Minor (and unpredictable) contaminating events , e.g. the loss of surface material by outgassing or upon micrometeorite

impact, would then be treated in a totally permissive manner. The difficulty with adopting a strategy of tolerance and/or

remedial treatment after the fact is that in a large and complex

future spacecraft, many potentially contaminating events are likely

to be occurring simultaneously. Hence the problem of predicting the

total contaminant flux at any one point and at all times is likely to become rather involved. Furthermore, the time that a given

sensitive element spends out of service may be excessive. Likewise,

a commitment to the totally permissive approach may turn out to

92
require unacceptable levels of "down-time" for cleaning or replacement of contaminated sensitive surfaces, and corresponding onerous burdens on logistics systems responsible for supplying the necessary

cleaning materials or replacement optics.

3: THE MODEL: SUMMAEY AND CONCLUSION

A conceptual model has been tentatively developed to describe

the phenomenology of spacecraft external contamination. This model

is based on the processes classically involved in mass transfer,

i.e. , sources, transport modes, and sinks or deposition points.

3.1: SOURCE MECHANISMS

The primary contaminant sources arise from engine firings,

waste dumps, and flowing or outgassing window sealants. Secondary

sources include outgassing paints, solders, and other structural

surface materials, and leakage from the spacecraft interior. These

amy be summarized as follows:

a. Engine firings (e.g. CAMS, RCS, LES, SPS)

(1) Propellant impurities

(2) Ablation products

(3) Ablator impurities

(4) Combustion products and intermediates

(5) Nozzle and explosive-bolt materials

b. Jettisoned wastes

(1) Human and animal excreta

(2) Fuel- cell effluents

c. Window mounting materials

(1) Sealants

(2) Gaskets

(3) Insulation

(4) Grommets, brackets

d. Outgassing of vehicle surfaces

(1) Structural components, e.g. LET, structural ablators

(2) Paints and lacquers

(3) Sealants and binders

(4) Flanges and gaskets, e.g. cork liners, docking flanges

(5) Wire, solders, weldments

e. Cabin atmosphere leakage and venting

(1) "Normal" cabin atmosphere

(2) Cabin interior contaminants

93
3.2: TRANSPORT MECHANISMS

a. Regarding the transport of gaseous matter:

(1) Uncharged gaseous metal atoms emitted in the ground

electronic state undergo neither thermal nor photoionization to any

significant extent and are therefore unrelated to window contamination,

but may be responsible for absorption and scattering of light in the

optical paths of instruments in space.

(2) Uncharged gaseous metal atoms emitted in excited

electronic states may be capable of photoionization or autoionization

in the solar flux, especially if the states concerned are metastable;

otherwise, they are more likely to return to the ground state, with

the loss of a photon. Those ions which may be formed in this manner

would be subject to the effects of electrostatic fields originating on

the spacecraft surface itself.

(3) Gaseous metal oxides may be subject to photodissociation

in the solar flux; in some instances, this may produce metastablestate metal atoms. In no case, however, are ionized gaseous metallic

oxides important components of the contaminant flux.

(4) Significant charge differences can be established

between dielectric and conducting surfaces on the spacecraft, which

may affect the transport of ionized gases.

(5) Dielectric surfaces facing out of the flow, especially

those facing rearward, probably achieve the most negative potentials, of the order of -10 to -15 volts, on the spacecraft.

(6) The firing of thrusters changes the charge state of the vehicle for at least the duration of the firing period.

(7) The net transport motion of gaseous ions with respect

to the spacecraft is a highly complex aggregate fimction of hydrodynamic forces characterizing the vehicle flow field, a gas-surface

interaction potential, the superimposed electrostatic field potential,

and possibly forces arising from interaction with the earth's magnetic

field. b. Regarding the transport of particulate matter:

(1) The metallic constituents reported analytically among

Gemini and Apollo window deposits probably were transported there

by entrainment within, or as an integral part of, particulate

contaminants

(2) Liquid matter, e.g. as droplets, undergoes phase

equilibration with respect to the corresponding solid and gaseous

states of the substances concerned.

(3) Solid matter in two-phase flow is coupled hydrodynamically to the enveloping gas; this coupling may be strong or weak, depending on the relative mass densities of the two interacting

phases.

94
(4) Therefore the net transport motion of particles with

respect to the spacecraft is determined either by interactions between

roughly equivalent mass densities of the two phases (strong coupling), or by significant dependence of the particle flow upon a dominant

gaseous flowfield (weak coupling).

(5) Particles approaching the spacecraft may undergo an induced polarization effect, regardless of their net charge, by electrostatic interaction with the vehicle; if this happens, the resulting field effects are superimposed upon the purely hydrodynamic effects.

3.3: DEPOSITION MECHANISMS

a. To recapitulate pertinent aspects regarding electrostatic

attraction:

(1) A net negative charge develops on the vehicle surfaces,

both conducting and nonconducting; magnitudes of the potentials on

nonconducting surfaces are dependent on their orientation with respect

to the flow, but are greatest (of the order of -10 volts or more) on

rearAvard-facing areas.

(2) The charge state of the vehicle is affected by thruster

firing.

(3) Currents are generated through conducting surfaces by

virtue of potential gradients.

(4) Gaseous ions are attracted electrostaticalh^ to the

vehicle, and especially to sideward or rearward-facing dielectric

surfaces.

(5) Particles carried toward the vehicle by flowfield

coupling to the gaseous envelope may be subject to induced polarization, with subsequent attraction to the vehicle, and especially

to the sideward or rearward-facing dielectric surfaces.

b. Regarding the impingement and adhesion of contaminants:

(1) Gaseous ions impacting the vehicle are neutralized;

both physical and chemical bonding processes ma\^ be active but some

of the impinging species may be lost bj^ scattering or desorption.

(2) Particulate matter impacting the vehicle adheres or

scatters depending on the relative magnitudes of the local vehicle

potential, the strength of the induced dipole, and the velocity vector

of each particle relative to the surface.

(3) Liquids spread or coalesce depending on their surface

tension in comparison to their interaction forces relative to the surface.

(4) Flowing window sealants and/or adhering particulates cover the surface and seal previous layers of deposited materials in

95
place; the dielectric nature of the contaminated surface is preserved

for repetitions of the mechanism.

4: COUNTERCONTAMINATION: SUMMARY AND CONCLUSIONS

The above contamination model suggests the applicability of certain corrective or countercontamination measures, for it is implicit in the notion of predictable mass transfer that the process

may be interrupted at any stage, given a suitable procedure or device.

Thus there appear to be at least three types of approach to the practical problem of counteracting contamination, viz. , prevention

of the appearance of potential contaminants by passive measures,

active removal measures to prevent or inhibit mass transfer for contaminants which do appear, and compensatory means of counteracting or discounting the degrading effects of contaminants which

complete their transit to sensitive surfaces.

4.1: PASSIVE PREVENTION

The number of conceivable contaminant sources is so great,

and their variety so diverse, as practically to preclude successful

passive prevention at each source. Passive prevention appears to be at best a means of reducing contaminant emissions, not necessarily

to the point of threshold tolerance by sensitive optics, but at least to the level of subsequent manageability by other means of control, e.g., active prevention.

4.2: ACTIVE PREVENTION

The interception of contaminant materials in transit between

the various sources and the sensitive optical and other instruments

which are to be protected appears to be technically feasible. Furthermore, since the point of interception can be chosen an3rwhere in the contaminant flow-field, convenience of operation will dictate that

this type of control be applied where it is most needed, in the immediate vicinity of the surface being protected. In this manner,

the effects of active devices should be readily predictable in terms

of actual effectiveness (as determined by suitable monitoring instruments) within the operational locus of the protected sensitive surfaces.

It remains to translate the concept of active countercontamination

into the construction and testing of hardware.
4.3: REMEDIATION

If all else fails, there is always the final resort of a remedial

approach, i.e. , either the toleraace of deposition (with subsequent

cleanup), or the shutdown of functional surfaces during major contaminating events (and permissiveness toward minor ones). However,

for complex spacecraft reliance on this strategy alone is expected to lead to excessive "down-time" of the instrumentation being protected.

4.4: A GENERAL STRATEGY OF COUNTERCONTAMINATION

It may be concluded from the foregoing discussions, that each

of the three approaches to countercontamination has something to recommend it, and each has certain drawbacks. Furthermore, one may also conclude that the three in fact tend to complement one

another, such that an effective countercontamination strategy ought

to be based upon their integration into one system.

Thus, passive prevention would be applied, in the selection of components and materials during the design and construction of spacecraft, such as to minimize the efflux of contaminants in space.

In order to compensate for those events which, despite all precautions,

would still release unacceptable amounts of contaminants into the spacecraft environment, active controls accompanied by suitable

local monitoring stations would be mounted in the vicinitj^ of all functional instruments. Finally, remedial procedures would be

devised and standardized for utilization either on the rare occasions

when a particular contaminating event might be of such a magnitude

as to overcome all preventive measures, or, equally rarely, at periodic intervals when the aggregate deposits of subthreshold amounts

of contaminants slipping by the preventive controls have accumulated

to an unacceptable degree and must be removed.

5: REFERENCES

1. Glenn, J. H. , Pilot's Flight Report in Results of the First

United States Manned Orbital Space Flight, February 20, 1962;

(National Aeronautics and Space Administration, Washington, 1962).

2. Carpenter, M. S. , Pilot's Flight Report in Results of the 2nd US Manned Orbital Space Flight, May 24, 1962 (National

Aeronautics and Space Administration, Washington, 1962).

3. O'Keefe, J. A. , "Preliminary Report on the Results of the MA-6 Flight in the Field of Space Science", in Results of the First

97
United States Manned Orbital Space Flight, February 20, 1962;

(National Aeronautics and Space Administration, Washington, 1962).

4. Tousey, R., Koomen, M. J., McCullough, R. E. , and

Seal, R. T. , "Photographs of Particles from an Aerobee Rocket at 130- to 204- km Altitude", Meteor Orbits and Dust, ed. G. S. Hawkins, NASA SP-135, SCA Vol. E, 333-344 (1967).

5. Grenda, R., Neste, S. , and Soberman, R. K. , "Contaminant Particle Trajectories near a Spacecraft", Space Research, Vol. DC, eds. K. S. Champion, P. A. Smith, and R. L. Smith-Rose,

North Holland Publishing Co. , Amsterdam (1969).

6. Hinton, B. B. , Leite, R. J., and Mason, C. J., "Neutral

Particle and Positive Ion Data Obtained from OGO-C", Trans. AGU

48, 74 (1967).

7. Ney, E. P. , and Huch, W. F. , "Optical Environment in Gemini Space Flights", Science 153 , 297 (1966),

8. Newkirk, G. , Jr. , "The Optical Environment of Manned

Spacecraft", Planet. Space Sci. 15, 1267 (1967).

9. Osborne, F. J. F. , and Kasha, M. A., "The VXB

Interaction of a Satellite with its Environment", Can. J. Phys. 45,

263 (1967).

10. Schmidt, I. , "Optical Environment in Gemini Space Flight,"

Science 155, 1136 (1967).

11. Tilly, M. J. , "Foreign Deposits on Gemini XI Optical

Windows, " Final Report, McDonnell Aircraft Corp. , NASA-MSC-GR-66-8, Supplemental Report 7, August 1967.

12. Muscari, J. A., and Cunningham, A. C. , "Gemini 12 Contamination Study, " Martin Company, Report R-67-2, January

1967.

13. Leger, L. J. , NASA Manned Spacecraft Center, private

communication (1969).

14. Baurer, T. , et al. , "External Spacecraft Contamination

Modeling and Countermeasures", GE TIS R 70 SD 260, June 1970 . 15. Dushman, S. , Scientific Foundations of Vacuum Technique,

John Wiley and Sons, New York, 1949.

16. Gibson, W. C, and Modisette, J. L. , "A System for Removing Contaminants from Spacecraft Optical Systems", J. Spacecraft Rockets 7, 353 (1970). 17. Goldstein, H. W. , GE Space Sciences Laboratory, Private

communication (1968). 18. Belton, M. J. S. , "Dynamics of Interplanetary Dust",

Science 151, 35 (1966).

19. Wyatt, S. P., "The Electrostatic Charge of Interplanetary

Grains", Planet. Space Sci. 17, 155 (1969).

20. Brundin, C. L. , "Effects of Charged Particles on the Motion

of an Earth Satellite", AIAA J. 1. 2529 (1963).

98
21. Liu, V. C, "Ionospheric Gas Dynamics of Satellites and

Diagnostic Probes", Space Sci. Eev. 9, 423 (1969). 22. Jen, N. C. , "Satellite Potential in an Ionized Atmosphere",

AIAA J. 3, 714 (1965). 23. Samir, U. , and Wren, G. L. , "The Dependence of Charge

and Potential Distribution Around a Spacecraft on Ionic Composition",

Planet. Space Sci. !]_, 693 (1969). 24. Beard, D. B., and Johnson, F. S. , "Ionospheric Limitations on Attainable Satellite Potential", J. Geophys. Res. 66, 4113

(1961).25. Sehnal, L. , "The Motion of a Charged Satellite in the Earth's Magnetic Field", Smithsonian Astrophysical Observatory

Special Report 271, April 1968.

26. Singer, S. F. , ed. Interactions of Space Vehicles with an

Ionized Atmosphere, Pergamon Press, New York (1965). 27. Al'pert, Ja. L. , Gurevich, A. V. , and Pitaevskii, L. P.

Space Physics with Artificial Satellites, Consultants Bureau, New

York (1965). 28. Shay, J. J., "Collisionless Plasma Flow Around a Conducting Sphere in a Magnetic Field, " Rarefied Gas Dynamics

5th Symposium, Academic Press, New York (1967); P. 1671.

29. Maslennikov, M. V., and Sigov, Yu.S. , "Discrete Model

of the Medium, in a Problem on Rarefied Plasma Stream Interaction

with a Charged Body, " Rarefied Gas Dynamics 5th Symposium,

Academic Press, New York (1967); p. 1657.

30. Kovar, N. S. , Kovar, R. P., and Bonner, G. P., "Light

Scattering by Manned Spacecraft Atmospheres", Planet. Space Sci. 143 (1969). 31. Carlson, D. J., Lewis, C. H. , Jr., and Bartky, C. D.

"Solid Propellant Rocket Exhaust Plume Study" (Unclassified), J. Missile Defense Research, Vol. IV, No. 1, Summer 1966.

32. Ashkenas, H. , and Sherman, F. S. , "The Structure and Utilization of Supersonic Free Jets in Low Density Wind Tunnels",

Rarefied Gas Dynamics , ed. de Leeuw, Supplement 3, Vol. 11, Academic Press, New York (1965). 33. Hamel, B. B. , and Willis, D. R. , "Kinetic Theory of Source Flow Expansion with Application to the Free Jet", Phys.

Fluids 9, 829 (1966). 34. Willis, D. R., and Hamel, B. B. , "Nonequilibrium Effects

in Spherical Expansions of Polyatomic Gases and Gas Mixtures",

Rarefied Gas Dynamics, Supplement 4, Vol. I, Academic Press, New

York (1967). 35. Cooper, A. L. , and Bienkowski, G. K. , "Asymptotic

Theory for Steady Source Expansion of a Binary Gas Mixture",

99
Rarefied Gas Dynamics , Supplement 4, Vol. I, Academic Press, New

York (1967). 36. Ehrlich, G. , "Atomic Processes at Solid Surfaces," GE

TIS 66-C-023, February 1966; "Atomistics of Metal Surfaces",

GE TIS 67-C-432, November 1967.

37. Pikus, I. M. , "High Speed Gas-Surface Interactions, 1. Preliminary Model, " GE TIS R67SD16, April 1967.

38. Saltsburg, H. , Smith, J. N. , Jr. , and Rogers, M.

Fundamentals of Gas- Surface Interactions, Academic Press, New

York (1967). 39. Krupp, H. , "Particle Adhesion Theory and Experiment",

Advan. Colloid Interface Sci. 1, 111 (1967). 40. Proceedings of the Atomic and Molecular Beams

Conference, University of Denver, Denver, Colo. , 20-22 June 1960.

41. Wise, H. , "Energy Exchange between Gases and Solids",

Project Squid, Technical Report SRI-15-P, Stanford Research

Institute, February 1963.

42. Saltsburg, H. , and Smith, J. N. , Jr. , "Molecular-Beam

Scattering from the (111) Plane of Silver", J. Chem. Phys. 45, 2175

(1966).43. Harris, R. E., "On the Determination of Thermal

Accommodation Coefficients in the Temperature - Jump Region", J. Chem. Phys. 46, 3217 (1967). 44. "The Role of the Adsorbed State in Heterogeneous Catalysis",

Disc. Far. Soc. No. 41 (1966); entire volume.

45. Logan, R. M. , and Stickney, R. E. , "Simple Classical

Model for the Scattering of Gas Atoms from a Solid Surface", J. Chem. Phys. 44, 195 (1966). 46. Logan, R. M. , and Keck, J. C. , "Classical Theory for

the Interaction of Gas Atoms with Solid Surfaces", J. Chem. Phys.

49, 860 (1968). 47. Raff, L. M. , Lorenzen, J., and McCoy, B. C. , "Theoretical Investigations of Gas-Solid Interaction Phenomena. I", J. Chem. Phys. 46, 4265 (1967). 48. Lorenzen, J., and Raff, L. M. , "Theoretical Investigations

of Gas-Solid Interaction Phenomena. II. Three-Dimensional Treatment", J. Chem. Phys. 49, 1165 (1968).

100
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TABLE 3 -- Available passive control measures

Contaminant Source Passive Control

Engine Throat Ablators

Engine Propellant Impurities

Waste Dumping

Window Seals

LES and LET Components

Paints, Solders, Wire,

Liners, Binders, Structural

Ablators, All Other External

Surfaces

Use Radiatively Cooled Engines

Clean up Propellants

Desiccate, Bag, and Save Wastes; Recycle Water; Modify Dumping

Procedures

Change Sealants; Utilize NonFlowing, Non-Volatile Materials

Keep Windows Covered and Delay

Deployment of Optical Experiment

until Complete Jettison of LET.

Undertake Careful Materials

Selection; Find Trade- Offs between Functional Performance and

Contaminant Emission Behavior

for Each External Area

Considered.

103

Paper No. 6

PARTICULATE CONTAMINANT MEASUREMENT BY

QUARTZ CRYSTAL MICROBALANCE

R. L. Chuan"^

REFERENCE: Chuan, R. L. "Particulate Contaminant

Measurement by Quartz Crystal Microbalance, " ASTM/IES/

AIAA Space Simulation Conference, 14-16 September 1970.

ABSTRACT: A piezo-electric crystal executing vibrations in a shear mode can be used as a mass balance with very high sensitivity. Various methods have been used by investigators in

diverse fields to bring the mass to be measured to such a bal- ance. Some of these are: absorption of a gas into the surface

of the crystal; condensation of a gas onto a cooled crystal;

chemical reaction with material bonded to the crystal surface;

and impaction of particulate matter onto the crystal surface.

The present oaper deals with this last mas s -accretion mechanism as applied to the measurement of contamination in the form

of particulate matter present in space environment (actual and

simulated)

The QCM (quartz crystal microbalance) responds to a mass

change on its surface typically in the following manner: Af = -Cf^Am, where Af is the change in the crystal resonant frequency

f due to the addition of massAm, and C is a constant whose magnitude is about 10^. For a nominal frequency of lO''' Hz the

sensitivity, Af/Am, is of the order of 10^ Hz / gm, from which

it is seen that the mass resolution of the QCM depends directly on frequency resolution. Frequency resolution, in any given

oscillator, depends on the characteristic time within which

measurements are taken - the resolution generally degrading

as the time increases. Thus for small mass change occurring

in a long time (by which we mean seconds or tens of seconds)

the frequency resolution is commonly not better than about 0. 1 to 1 Hz for a lo''' Hz oscillator. However, if the characteristic

measurement time is less than, say, 10"^ sec, the frequency

resolution can be quite good - being of the order of 10""^ Hz or

1Mpnager, Advanced Technology Group, Atlantic Research

Corporation, Costa Mesa, California.

105
better. This consideration lends itself to the sensitive and

accurate measurement of discrete particulate mass, because

when such a mass is deposited on the crystal there is an almost

instantaneous, though small, frequency shift. By utilizing

this very high rate of frequency shift it is possible to measure

particulate mass down to the order of 10" ''^ gm. This is equivalent to a single particulate of 2 micron diameter with mass

density 2 gm cm~-^.

In addition to measuring the mass of particulate matter, the

technique also provides information on whether the material

is volatile. If subsequent to its impaction onto the crystal

surface, with an accompanying frequency shift, the particulate

material evaporates, there will not only be a loss of mass but,

more impoartantly, there will be cooling of the crystal. The

crystal can be chosen so as to exhibit a frequency change of a sign opposite to that caused by mass addition. Then the instantaneous frequency change caused by the impaction of particulate

will be followed closely by an opposite frequency change due to evaportation.

Experimental data will be presented showing the response of the QCM to non- volatile and volatile particulate matters.

KEY WORDS: spacecraft contamination, particulate sampling,

quartz crystal microbalance

INTRODUCTION

The recognition in recent years that a spacecraft can con- taminate itself both in actual space environment and in laboratory simulated space environment has led to a variety of devices with high sensitivity to detect the presence of contaminants. As more data becomes available, from both flight and

laboratory experiments, it appears that a significant amount

of the contaminants released by a spacecrart (particularly on manned spacecraft) is in the form of particulates rather than

vapors. This paper describes an instrument system which measures the flux of vapor and particulate contaminants , and can distinguish between the two types as well as measure the mass of discrete particulates. In addition, it can distinguish

between solid and volatile particulate matters.

What is required basically as a pertinent measure of spacecraft contamination is the determination - directly, in

real time, if possible - of the mass flux per unit area per unit

time incident on the surface in question. Depending on the

criticality of the effects of contamination, the resolution in such

a mass flux measurement may be between 10"-'^*-' kg/sm^ and

10"^kg/sm^ (those corresponding roughly to monolayer

106
formation times of 3 hours to 1 second). If the contaminant

is in particulate form, the resolution in discrete mass measurement need be of the order of 10" ''^ kg (corresponding to the mass of a 2 micron diameter particle of mass density 2 x 10-^ kg/m^.

The quartz crystal microbalance can be used quite adequately for such measurements. The QCM is a piezo-electric

crystal vibrating in a shear mode such that the addition of mass on its surface results in a change in its resonant frequency. The device is compact and operates largely free from

environmental influences such as pressure, temperature and

gravity, and thus is well suited for use on a spacecraft as well

as inside a laboratory chamber. Several configurations of the

QCM are described in this paper, together with typical experimental results.

CHARACTERISTICS OF THE QCM

The basic sensor is a small disc of piezo-electric crystal

(about 1 cm in diameter and 0. 014 cm in thickness) which

forms part of a resonant circuit whose frequency is controlled

by the mechanical resonant frequency of the crystal. In our

application the crystal, cut from quartz operates in a shear

mode whereby the opposite faces execute parallel displacements. In this oscillation mode the frequency is affected by

the mass on the crystal surface. For small amplitudes of oscillation and small variation in mass, the resonant frequency

is inversely proportional to mass:

where f is the change in frequency caused by a mass change

Am on an active area A of the crystal, and C is a constant

which in our case has the nominal value of 2 x 10"^, with f expressed in Hertz, and m in kg and A in square meters.

Figure 1 shows a pair of crystals with nominal frequency

of lO"^ Hz. The sensitivity, by the above formula, is then

Af = -2 X 10^

^

Hz

A

In practice this change in resonant frequency of the sensing

crystal is detected by the change in the beat frequency created

by mixing the frequency of the sensing crystal with that of a reference crystal with slightly higher frequency, as shown

schematically in Figure 2. The beat frequency, being much

lower than the sensing frequency, makes it much easier to measure a small frequency change. Instead of measuring a change, of, say, 10"^ Hz in 10^ Hz, it is only necessary to measure 10"^ Hz out of 10^ Hz or less. An increase in mass

shows as an increase in the beat frequency. The type of quartz crystal used exhibits a slight sensitivity to temperature

variation (about 5 Hz k-1). By placing the reference crystal in

107
the same thermal environment as that of the sensing crystal

(but without being exposed to the impaction air jet) temperature

compensation is automatically achieved.

DISCRETE MASS SENSING

In order to erlsure that particulates incident on the crystal

surface will attach to the surface and not bounce off; a thin

layer (about 1 micron thick) of adhesive substance, which is inert and has very low vapor pressure, is aoplied over the

crystal. The arrival of a particulate mass on the crystal

results in a step-wise rise in the beat frequency. If the beat

frequency is converted to an analog voltage and then differentiated, the step-wise frequency rise becomes a pulse whose

height is proportional to the mass of the particulate. The

sensitivity attainable for such a means of measuring particulate

mass is estimated in the following manner.

With conventional circuitry it is possible to resolve a. frequency change, over a short time interval, of the order of 10"^ Hz. Referring to the crystal sensitivity formula given

above, it is seen that a frequency shift of lO"^' Hz results

from a mass change per unit sensing area of Am r- 1^-10, , 2

= 5x10 kg/m

A

-5 2 2 The area typically employed is about 10 m (10 nm ), so

that the attainable mass resolution is 5 x lO-lS^g. Actual

test results to date have reached a mass resolution of about

1 X 10" '•4 kg. This is the mass of an equivalent sphere of diameter 2 microns with a mass density of 2 x iO^ kg/m-^.

Figure 3 shows an actual recording trace of a differentiated analog signal from the beat frequency change. The five

readily discernible pulses represent discrete masses of from

5 X 10" ''^ to 6 X 10" '^ kg. These are produced by particulates

of mass density 3 x 10-^kg/m^; thus the largest pulse represents a single particulate of diameter 7. 5 microns.

If the particulate matter incident on the crystal is of a

volatile material, some of the mass initially captured may be

lost due to evaporation, which is also accompanied by some

cooling of the crystal surface. The resulting response in the

differentiated analog signal is a pulse followed by a negative

pulse - the former due to the arrival of the mass, and the

latter due in part to mass loss and in part to cooling. An

example of this is shown in Figure 4. The pulses here are produced by water droplets - the difference from the signals

in Figure 3 are quite apparent. The largest pulse here

represents a water droplet of diameter 13 microns.

EXAMPLES OF PARTICULATE SENSORS

Figure 5 shows a sensor used primarily for space chamber

monitoring. The main feature of it is the crystal sandwich

108
in which the sensing and reference crystals are bonded to- gether so as to achieve good thernnal compensation, as well as

to conserve space. The two oscillators and the mixer are

built into a single board closely behind the crystal sandwich.

The output beat frequency is brought out of the space chamber

for further processing.

Another example of the particulate sensor is one designed

for use on a spacecraft to protect sensitive optical surfaces.

Figure 6 shows a unit built for the Apollo Telescope Mount.

A thermister is attached to the crystal sandwich to monitor

its temperature so that corrections can be made to compensate

for slight shifts in frequency over a wide range of temperature

variation (from about -10° C to +100°C). There is a heater

behind the crystal assembly which is used to clean off volatile

materials if too much accumulates.

Figure 1. Quartz Crystals

109
Reference

Crystal

Beat Frequency f, = 10 Hz

Reference

Oscillator

f„ = 10^ + 10-3

Figure 2. Schematic Diagram of Particulate Instrument

ax 1

. 4.» X 10

Figure 3. Solid Particulates

Figure 4. Liquid Particulates

110
Figure 5. Microbalance for Space Chamber

Contamination Monitoring

nLTtR cmcinr

10 CM

Figvire 6. Microbalance for Spacecraft Contamination Monitoring

111

Paner fvo, 7

SPACE MEASUREMENTS OF THE CONTAMINATION OF SURFACES BY OGO-6

OUTGASSING AND THEIR CLEANING BY SPUTTERING AND DESORPTION^

D. McKeown and V/. E. Corbin, Jr."*"

REFERENCE: McKeovm, D. and Corbin, W.E. Jr., "Space Measurements o£ the Contamination of Surfaces by OGO-6 Outgassing and

Their Cleaning by Sputtering and Desorption',' ASTM/IES/AIAA

Space Simulation Conference, 14-16 September 1970.

ABSTRACT: Results of the contamination of surfaces by outgassing of the OGO-6 satellite and the rates at which these surfaces are now being cleaned by sputtering and desorption after

being in space for five months are given. The contamination

measurements are being made with a quartz crystal microbalance

to define the conditions of Al and Au surfaces used in an ex- periment to measure gas -surface energy transfer. It was found

that the primary source of outgassing on the satellite was its two solar panels baking out in the sun. The time constant for

the exponential decay of the outgassing is 1,000 hours. The

maximum amount of contamination adsorbed by the surfaces ex- posed to the outgassing was reached after five months in orbit

and is 96 mg/m^ for the Al surface and 52 mg/m^ for the Au surface. The contamination has a desorption activation energy of

26 kcal/g mol which falls into the energy range of materials

such as, epoxies and vacuum oils. The surfaces are now undergoing cleaning by desorption at 1.2 x lO'^g/m^s and sputtering

by upper atmospheric neutral impacts at 2.3 x lO'^g/m^s. A

method to greatly increase the cleaning rate by removal of the

contamination by sputtering with upper atmospheric ions is discussed. KEY WORDS: Surface contamination, space, outgassing, solar

panels, sputtering, desorption, cleaning, ions, quartz crystal

microbalance

Staff Members, Faraday Laboratories, La Jolla, California

2 This work was supported by NASA Goddard Space Flight Center

under Contract NAS5- 11163 and the Office of Naval Research un- der Contract NOOO14-68-C-0373.

113
INTRODUCTION: With the planned orbiting of large manned space

laboratories in the 70 's, the space environment has taken on a new dimension. Not only must problems be considered concerned

with the effects of the environment on the laboratory but also

the effects of the laboratory on the environment. One area of

concern is the contamination of surfaces , such as , telescope

mirrors and windows by the outgassing of materials used in the

construction of the craft. The contamination problem is es- pecially troublesome because of the free molecular flow conditions that exist in the hard vacuum of space. Molecules issuing from the contamination source flow freely in straight

lines to exposed surfaces. If these contaminated surfaces al- so point towards the sun, solar UV will compound the problem

by polymerizing the oil and grease contaminants and a hard

carbon glaze will fom resulting in significant changes in the

optical properties of the surfaces and in the accommodation

coefficients for gas -surface energy transfer.

At present we are conducting a space experiment to mea- sure gas -surface energy transfer by upper atmospheric atomic

and molecular impacts to determine how satellite drag is effected by various surface materials (1)-^. The experiment is

being flown on OGO-6 launched on 5 June 1969 into a polar or- bit with a perigee of 397 km and an apogee of 1098 km. The

experiment uses four quartz crystal probes as dual microbalances and energy transfer probes. One important part of the

experiment is to monitor the rate at which material is outgassing from the satellite, how much of this material is adsorbed

by the experimental surfaces, and after the outgassing rate drops to low levels, at what rate do the surfaces clean by desorption and sputtering erosion? Because the rates at which

surfaces are contaminated by satellite outgassing and cleaned

by desorption and sputtering are of importance to the understanding of surface contamination on other spacecraft, we would like to present here the results of measurements on the

effects of a space environment on surfaces obtained from

OGO-6.

INSTRUMENTATION: The flight experimental instrumentation is

shown in Fig. 1 and is 11.4 x 17.8 x 19.1 cm, weighs 2.2 kg,

and operates at 28 Vdc on 3.5 r/. The experiment has been described in early work (1,2) and only the general operation of

the instrumentation will be given here. The surfaces used in

the experiment are housed in the four gold plated probes located behind the shutter wheel. The experiment is pointed along

the velocity vector of the satellite into the air stream. To

1 The numbers in parentheses refer to the list of references

appended to this paper.

114
minimize weight in the OPEP appendage of OGO-6 which has attitude control for pointing into the upper atmospheric stream,

the experiment was flown in two parts. The four probe sensor package is located in the OPEP and connected by a 3 m cable to the L- shaped electronics package located in the main body of

the satellite. Four different measurements are being made on two Al and

two Au surfaces. The surfaces are plated by vacuum evaporation

on 15 IVHz, 1.27 cm diameter, optically polished Y-cut quartz

crystal sensors shown in Fig. 2. The sensors are located in

gold plated probes 3.3 cm in diameter and 9 cm long, and view

space through a 1.8 cm diameter window. The field of view of

the surface is 0.13 steradians and the probes effectively collimate the contaminate flux reaching the surfaces.

The probes are designated A through D. Probe A is instrumented with an Al plated sensor mounted normal to the probe

axis. Probe B and C have Au plated sensors mounted noimal to the probe axis. Probe D has an Al plated sensor mounted at 14 degrees to the probe axis. Probes A, C, and D are pointed into

the molecular stream and both contamination and energy transfer

measurements are made with these probes. Probe B is pointed

away from the stream by 13 degrees and the molecular stream

loses its energ>^ by multiple impacts on the wall of the probe,

and only contamination measurements are made with this probe.

An eight position shutter, shoxm in Fig. 1, is mounted in

front of the probes. Two of the 1.8 cm diameter aperatures in

the shutter have a 0.5 mm thick BaF2 blank mounted over it for

chopping the beam. BaF2 is used for the shutter because its

wide IR band minimized temperature changes during shutter op- eration. The shutter can be periodically stepped open to permit the upper atmospheric molecular stream to simultaneously

impact on the Al and Au surfaces to measure gas -surface energy

transfer. The shutter is programmed to remain open a total of

671 of the time and closed for 33% of the time and 67% of the

outgassing flux reaches the surfaces. On every apogee pass

(100 min) the frequency outputs of the probes are read out to determine the mass loading caused by contamination. The conta- mination measurement is made at this point in the orbit because

the upper atmospheric kinetic energy flux falls below 100 mW/

m2, the energy transfer threshold of the probe, and energy

transfer does not interfer with contamination measurements.

The probe output frequency is the beat frequency between

the sensor crystal and a matched reference crystal. The beat

frequency is used because it effectively cancels out ambient

temperature changes and any variation in the power supply voltage driving the probe. The reference crystal is mounted in a chamber just behind the sensor crystal. The chamber front wall

is 0.05 mm thick gold plated stainless steel with a 1.27 cm hole cut in it and is attached to the circumference of the sen- sor crystal. This does not affect the crystal oscillation be- cause an amplitude node exists at its edge. The chamber keeps

115
any contamination from reaching the back surface of the sensor

crystal or either surface of the reference crystal which would

effect the accuracy of the measurements. The Q or quality fac- tor of the crystals is 200,000 which assures high frequency

stability from aging and any changes in the transistors used to drive the crystal oscillators. The general equation for the

frequency change, AF, of the sensor crystal caused by mass loading (3) is

AF = (AmF^2/k pA) + F^(T) - F^(T) (1)

where Am is the contaminant loading in g , F^ , the fundamental

frequency of the crystal in Hz, k, constant of quartz = 1.66 x

10^ mm, A, the motional area of the crystal in cm^ , and p, the

density of quartz in g/cm^, and F^CT) the frequency dependence

of the sensor and reference crystal to ambient temperature

changes. A = 0.1 cm^ and letting AF^ = F^(T) - F^(T)

Am = (2.5 X 10" Vcm^Hz) (AF - AF^) (2)

The ambient temperature of the probes is monitored to

t 1°C and AF-j = 20 Hz/°C for Probes A, B, and C and these

probes can measure contaminant mass loading to an accuracy of

t 5 X 10-8g/cm2 (t 0.5 mg/m2) . The sensor in Probe D is affected to a greater degree by temperature because it is mounted

at a 14° angle to the axis of the probe. Its accuracy is t 2 x 10-7g/cm2 (1 2 mg/m2) . SURFACE CONTAMINATION MEASUREMENTS: The surfaces were prepared

in the laboratory by vacuum evaporation and sputtered atomically clean before instrumentation in the experiment. From the

laboratory until insertion into orbit the Al surface collected

32 mg/m2 and the Au surface 14 mg/m^ of mass from exposure to atmospheric contamination on the ground (Los Angeles smog)

Results of contamination of Al and Au surfaces on OGO-6

are shown in Fig. 3. Data given from launch until 31 July are production data from the continuous readout of the experiment.

From 1 August on, the data is "quick look" received every few

days for several minutes at a time which accounts for the gaps

in the measurement points.

The Al surface placed normal to the direction of the con- tamination flux has adsorbed approximately four times as much

contamination as an Al surface mounted 14° to the flux and

shows that the outgassing is external to the experiment and not internally generated. The Al surfaces have adsorbed about

twice as much mass as the gold surfaces. The Al is probably

gaining more mass because the outgassing material has a larger

sticking coefficient for Al than Au. The Au surface in the air

stream has gained a little more mass than the Au surface

116
shielded from the stream. One explanation for this is that the

contamination layer on the surface in the molecular stream^ is picking up a small amount of atomic oxygen by chemisorption

while the control surface shielded from direct atmospheric impacts is not reacting with 0. The contamination flux reaches a maximum when the surfaces

are pointed towards either one of the two solar panels on the

satellite in June and drops to near zero in August when pointed

into deep space because of rotation of satellite orbit plane to

the earth-sun line, and shows that the panels are the primary

source of contamination.

The satellite outgassing flux consists of two main components, volatile gases such as O2, COo, and and relatively low-volatile materials such as oil and paint that require a long period to outgas.

The volatile materials boil off the satellite and reach

very low levels in the first few days after launch. The cloud

that enveloped the satellite apparently shielded it from direct

atmospheric iirpacts for many hours and incoming atmospheric mo- lecules lost a significant portion of their kinetic energy by

particle-particle interactions before impacting on surfaces.

A'e were unable to detect any gas -surface energy transfer from

the upper atmosphere until 72 hours after launch. Because these

volatile gases are not adsorbed to any significant amount by

the surfaces used in the experiment, no quantitative measure- ment on the flux of these particles was made.

The low volatile materials outgassing from the satellite

stick to the experiment surfaces and an estimate of the flux of

the low- volatile material was possible. Assuming a sticking

coefficient of the flux on aluminum of 0.7 at the start of adsorption (4) and taking into account the time the shutter is

closed and the view angle of the surface, the mass flow to the

experiment at zero revolutions from Fig. 3 is

[dM/dt]= 9.2 X 10"-^Vcm^s (3) t=0

^ 2 In the International System units the flow is 9.2 x 10 ^ g/m s. Assuming the molecular weight averages 200 g, because low- volatile oils have large molecular weights, the outgassing flux in

the vicinity of the experiment is

F = 2.7 X lO-^-^part/cm^s (4) In SI units the flux is 2.7 "x lO^^part/m^s . Tliis flux rate de- creases exponentially with time as the panels outgas. Fitting

an exponential curve to the Al curve for the first three days

after launch gives a time constant of 1,000 hours for the exponential decay and

117
F = 2.7 X 10^\~^^^^ part/cm^s (5) The decay in the outgassing flux is plotted in Fig. 4 from

launch through January 1970.

The flux plotted in Fig. 4 is for the case of the experiment remaining fixed relative to the solar panel. Since the

experiment is constantly turning to remain pointed into the air

stream and the panels are turning to remain pointed at the sun, the flux reaching the face of the experiment is

F = [2.7 X 10-" e'^/-''' ][k + cos(e - tt/4) ] (6) where F is in part/cm^s, t in hours, k a constant < 0.1 to ac- count for outgassing other than from the panels, and 9 the angle

between the experiment face and the solar panels. The rate of

rotation of the orbit plane to the earth- sun line detemines

the angle. 9 , between the experiment and the panels where 9 = 4.5 X 10"^ nt where t is in hours.

DESORPTION: The desorption or outgassing of materials from

OGO-6, although easily measured by a quartz crystal microbalance, is very small and reflects the care with which materials were chosen in the construction of the satellite. The maximum

amount of the outgassing material adsorbed by the aluminum sur- face was 9.6 lag/cm^ (96 mg/m^) which represents less than 50 monolayers, assuming an average molecular weight of 200 g. It has been pointed out that the solar panels are the primary outgassing source on OGO-6. The front surfaces are solar

cells and the back surfaces of the panels are painted with Zr02

pigmented potassium silicate inorganic paint. Because this ma- terial has a low vapor pressure, epoxies used in assembling the

panels and contamination absorbed in vacuum tests must be the

contaminant sources. This material is now deposited on the Al and Au test surfaces. The preliminary desorption rate of the material can be

found when the experiment was pointed toward deep space in August. The amount of adsorbed gas, S.^, on the Au surface (4) at time t is

= exp(-t/Tp (7) where is the amount on the surface at t = 0, and t-^ the av- erage residence time of the gas.

= exp(AE/RT) (8)

where Tq is the vibrational period of the lattice, AE the desorption activation energy, R the gas constant, and T the abso118
lute temperature.

The desorption activation energy for the contamination is AE = -RT[ln(2.8 x lO'^^'/t) (In ] (9)

The average temperature of the surface was 7°C and from

Fig. 3 during August for Probe B: Sq = 4.5 x 10"^g/cm-, S-^ = 4.1 X 10-6g/cm2, t = 750 h, and

Griffith (4) points out that diffusion purrp oils and epoxies have desorption activation energies of about 25 to 30 kcal/

g mol and begin to vaporize in significant amounts at temperatures above 45°C. In full sun light the OGO-6 solar panels op- erate above 45°C. The contamination outgassing from the panels

is then most probably epoxy asseiribly materials and diffusion

pump oil absorbed during pre-launch vacuum tests. RADIANT ENERGY EXCHANGE: In designing the experiment, a thermal model was made of the instrumentation to minimize the energy radiated from the surfaces generated by gas -surface interactions to maximize the temperature increase of the sensor to at- tain large output signals.

Radiant energy exchange between elements of a system of

nonblackbodies were computed by

where a, is the Stefan-Boltzman constant, ^i^> the ordinary

view factor for blackbodies, Aj^, the area of the i'th element,

T-, the temperature of the i'th element, T^ , the temperature of

the j ' th element.

Fg-Lj is a factor Avhich accounts for the energy transferred

between i and j w^here heat transfer is a function of multiple

reflections between i and j themselves and other elements of

the system, ^eij ^ function of both the geometry of the sys- tem and the omittance of the various elements . The exchange of

energy is directly related to the emissivity of the surfaces

and would be reflected in the exchange rate measurements made

during calibration.

The energy transfer sensors are calibrated periodically at the several power levels. Fig. 5 is an 878 i-iW calibration of

the aluminum plated sensor made on 7 June 1969 soon after launch

and Fig. 6 is the same level calibration made on 9 August 1969

after accumulation of 8 \ig/cm^ of contamination. A constant power was applied to the sensor for 110 sec. The probe frequency output was 1475 Hz in Fig. 5 and 1450 Hz in Fig. 6 and

changed less than 2% which is the accuracy of reading out the

AE = 26 kcal/g mol (10)

4 4 q. - = aF. - F . . A-fT - T.) ^ij 11 eij 1^ 1 1^ (11)

119
probe frequency. The near normal omittance of Al surface was measured in the laboratory and found to be 0.02 and has not changed significantly after being contaminated in space because

the power calibration has not changed significantly.

Figs. 5 and 6 are typical of the output of the probes when

energy transfer is taking place produced by molecular impacts.

An exponential curve fit is made by computer of the frequency

increase of the probe, F(t) , to find the beam energy trans- ferred when the shutter opens.

F(t) = F^(l - e'^/'^l) + F^ (12)

where is a time constant and F^ is the asymptotic frequency

increase of the probe when the shutter opens. The constants

are found by calibrating the probes in space (1)

SPUTTERING: Sputtering is the ejection of surface atoms by en- ergetic impacts of neutral atoms or ions. For impact energies

below 50 keV the interaction between the bombarding particle

and target atom is treated as a hard sphere collision. For nor- mal incidence the bombarding particle undergoes at least one collision to reverse its momentum vector before a second collision can produce sputtering. The bombarding ion strikes a tar- get atom and rebounds in such a direction as to sputter another

atom from the surface. The other possibility is that the bombarding particle displaces from its site the first atom it meets and is subsequently lost in the interstitial spaces of

the lattice. Meanwhile, the displaced atom uses up its excess kinetic energy to collide with other atoms in the lattice. A

certain percent of the time, this atom will strike an atom near

the surface with enough energy to cause sputtering. At oblique

incidence a single collision can lead directly to the ejection

of a surface atom. The sputtering yield is defined as the number of surface

atoms ejected per impacting particle. The yield is dependent

upon the angle of incidence, the energy, and the mass of the

impacting particle and the mass of the target atom, the average

binding energy of an atom in the lattice, the crystal structure

of the solid, the heats of sublimation of the different crystal

planes, the collision cross section of the target atoms, and

the energy dissipated in the lattice as heat. Henschke (5) points out the dependence of the yield below impact energies of

100 eV on the collision radii. This dependence reveals a marked

periodicity of this energy as a function of the atomic number

of the target atom. More specifically, this behavior is seen

to be related to the filling of the d subshells, (electronic

orbital angular momentum Jl = 2) in the electronic structure

of the elements. It is observed that for these atoms with nearly completely filled d shells, the collisions are more elastic

and less energy is lost in the lattice. This results in higher

120
sputtering yields for these solids. Unfortunately no theoretical relationship exists to predict the sputtering yield for any

arbitrary particle- surface combination and yields must be mea- sured experimentally for particular combinations.

Rol et al (6) and Almen and Bruce (7) have proposed that

the dependence of the yield on impact energy and mass is

where \j. is the yield in atoms/particle, 3 a proportionality con- stant to fit the experimental data, N the number density of sur- face atoms, a(E) the energy dependent scattering cross section, m the mass of the impacting particle, M the mass of the surface

atom, and E the particle impact energy.

The impacting particle energy, E, is expressed in eV. From Eq. 13, the following observations can be made. The scat- tering cross section is known to increase with energy so yields

will increase strongly with impacting energy. The yield will

be maximized when the mass of the impact particle equals the mass of the surface atom (m = M)

Experimentally it has been found that for most particle

surface combination yields as a function of energy fall in the

following ranges:

Measurements for Au in noble gas ions in Fig. 7 illustrate the

effect of impacting particles of different masses between 100 eV

and 800 eV. As the beam energies are reduced below 100 eV to 10 eV the

sputtering yields fall exponentially to values of less than

10"3 atom/ion. It is not possible to conduct sputtering experiments at yields below 10"-^ because laboratory beams are not in- tense enough to produce measureable mass losses in the few

hours time required in the best vacuum systems before system

impurities contaminate the surface. In the upper atmosphere,

an orbiting satellite generates a molecular beam in the 10 eV

range about 1,000 times more intense than is possible in the la- boratory. Sputtering can then be measured on a satellite if

the surface is placed in the molecular stream pointed towards

space so that outgassing cannot contaminate it. During August in Fig. 3, it was possible to determine the

sputtering on OGO-6 when the surfaces were pointed into space. The desorbing miass loss rate of the Au surface shielded from

impacts is 1.2 x 10"^ g/m^s. The mass loss rate for the other

surface caused by sputtering and desorbtion at normal incidence

14° incidence = 3.8 x 10"^ g/m2s. Subtracting the desorption

mass loss rate of 1.2 x 10"9 g/m2s from the surfaces gives the

[1 = m a(E) [4mME/(m + M) ^] (13)

0 < ^< 0.5;0 < E < 100 eV

0.5 < ^ < 15; 100 eV < E < 10 keV

15 < ^ < 50; 10 keV < E < 100 keV

121
sputtering rates. The average particle flux over an orbit is

1.1 X 10^' molecules/m^s , and assuming the contamination has a molecular weight of 200 g, the sputtering yields are

la^^^io = 3 X 10 ^ atom/molecule

^90° ^ ^ atom/molecule

These yields are not for Al and Au but the contamination

on the surfaces. The grazing angle of incidence produced high- er yield because the impacting molecule can eject atoms without

having to reverse its momentum direction.

The sputtering yields are in agreement with the Au yield,

|j = 1 X 10" " atom/molecule, measured on Discoverers 26 and 32

(8) with a quartz crystal microbalance mounted on an appendage

pointed continually into space. The lower yield for atomically

clean Au is indicative of the higher heat of vaporization for

Au than for the oil type contamination measured on OGO-6.

DISCUSSION: The measurements on OGO-6 show that outgassing of

low-volatile materials which cause surface contamination is produced primarily by structures baking out in the sun such as so- lar panels. The contaminated surfaces can be cleaned by desorption and sputtering.

Desorption requires that the surfaces be heated to tem- peratures above 45°C to produce acceptable cleaning rates. The

heating should be done electrically or indirectly by solar heating because if the surface is exposed to the sun, UV will polymerize the oil contamination leaving behind a permanent carbon

glaze which in most cases could cause more problems for a sur- face than a coating of oil.

Sputtering is another cleaning method. It has a big advantage in that it can remove carbon contamination which desorption cannot. To sputter carbon, the impacting particles

must transfer at least 5 eV, to the surface to break the carbon

lattice bond. The neutral upper atmosphere impacts between

5 eV and 10 eV and will sputter carbon but is impractical be- cause the yield is too low. A better source of sputtering particles is upper atmospheric O"*" ions. Metallic surfaces would

be biased to a potential of several hundred eV. A biased wire

grid, 95% open, would be moved in front of a non-conductor, to accelerate the ions. Space charge would spread the ion beam

and no grid shadow would fom on the surface. The surface

would be kept from charging up by impacting ions by directing

about a 1 mA/m^ flux of electrons at the surface.

During the day at altitudes below 1,000 km, the ion density is about 10^ ions/m3 (9) . An orbiting satellite sweeps

out 8 x lO-'-^ ions/m^s. At impact energies of 500 eV, the sputtering rate is about a million times greater than the rate mea- sured at 10 eV on OGO-6. The ion sputtering cleaning rate would be about 1 mg/m^s, which would clean the 100 mg/m2 of

122
contamination from the surfaces on OGO-6 in 100 seconds.

Above 1,000 km, ion densities will be too low for sputtering cleaning. At these altitudes an ion gun would be employed

using an onboard gas supply to sputter the surfaces. The electronics already developed for use in ion engines would be a good starting point for the design of such a gun. REFERENCES

(1) McKeown, D. and Dummer, R.S., IEEE Geoscience Electronics,

Vol. GE-7, No. 2, April 1969, p. 98.

(2) McKeown, D. , Rarefied Gas Dynamics, J. A. Laurmann, Ed.,

Academic Press, New York, 1963, p. 315.

(3) McKeown, D. , Review of Scientific Instruments, Vol. 32, No. 2, Feb. 1961, p. 133.

(4) Griffith, J.S., Proceedings 4th Space Simulation Conference, AIAA/ASim/IES, Los Angeles, Sept. 1969.

(5) Henschke, E.B. , Physical Review, Vol. 106, No. 5, March

1961, p. 1286.

(6) Rol, P.K. , Fruit, J.M. , Vielbock, F.P., and de Jong, M. , . Ionization Phenomena in Gases , N.R. Nilsson, Ed., NorthHolland Publication Co., Amsterdam., 1962, p. 257.

(7) Almen, 0. and Bruce, G. , Nuclear Instruments and Methods,

Vol. 11, 1961, p. 279.

(8) McKeown, D. , Fox, M.J. , Schmidt, J.J. , and Hopper, D.

AIAA Journal, Vol. 2, No. 2, Feb. 1964, p. 400.

(9) Hanson, W.B. and Sanatani, S., Proceedings AGU Annual

Meeting, SPA59c, Washington, D.C. , April 1970, p. 32.

123

125
126
127

Paper No, P

SIMULATION OF RADIATION FROM RTG POWER SOURCES

Phillip A. Newman, Jr., NASA, Goddard Space Flight Center, Greenbelt, Md.

ABSTRACT: The recent interest in Radioisotope Thermoelectric

Generator (RTG) power systems for deep space missions and the ever

increasing use of the radiation sensitive MOSFET in spacecraft electronics has made it essential that an assessment of the radiation

effects for this combined application be determined. An RTG with its continuous flux of neutrons and gamma photons provides a long term

hostile environment for electronic components. This paper deals with

a radiation source (Pu^^^ -Be) selected to simulate, for the most part, the neutron effects of the RTG. Preliminary experiments with this source have indicated that for most electronic applications the neutron

contributions of the RTG over a five year mission in present spacecraft configurations will not be a major problem. An analysis of the

data from several experiments indicate that predictions of the radiation effects can be made for "hotter" configurations with a reasonable

degree of confidence.

KEY WORDS: radiation simulation, neutrons, MOSFETS, RTG, test, evaluation, Pu^^^-Be

INTRODUCTION

Until recently NASA has not been interested in the effects of neutrons on spacecraft electronics since neutrons do not exist to any

substantial degree in the space environment. The proposed use of Radioisotope Thermoelectric Generator (RTG) power systems on deep

space missions has made it essential that an assessment of the effects

of the radiation on the more sensitive electronic components be determined.

129
RADIATION EFFECTS ON SPACECRAFT ELECTRONICS

Most of the advanced spacecraft electronic systems being designed today are making extensive use of MOSFETs (metal-oxidesemiconductor field effect transistor) both as discrete devices and as integrated circuits because of the substantial power and weight savings

and the decrease in circuit complexity. Prior work*^^°^^ has shown

that MOSFETs are the most sensitive of all commonly used electronic

components to ionizing radiation. While some neutron work has been

done on MOSFETs none has been done at the low levels expected from

the RTG.^4 5^ It is well known that radiation can cause permanent damage to semiconductors in two basic ways, ionization damage and displacement damage. Ionization damage, when it occurs, takes place at a much lower dose level than does displacement damage. For the predicted levels here, ionization damage should be the dominant effect.

In discrete MOSFET devices the most prominent ionization effect is the shift of the gate threshold voltage (and consequently the gate

voltage-drain current characteristic) due to a charge build-up in the gate insulator. This shift is strongly dependent on the gate bias

during irradiation.'-^-' While this threshold shift certainly takes place

in integrated circuits it is not always accessible for measurement.

Since gamma radiation effects on MOSFETs have been studied in the past and have been done extensively in this laboratory this preliminary experiment was largely an effort to isolate the effects of the neutron.

SIMULATION REQUIREMENTS

Figure 1 shows a possible spacecraft configuration for a deep

space mission using an RTG power system. The RTG fuel, a special

form of PUO2 (Pu^^^), emits a flux of neutrons and gamma photons

which are impractical to shield against in the spacecraft situation. The location of the RTG, about 2 m from the main part of the spacecraft, indicates that the spacecraft electronics package will be subject

to a long term hostile environment.

An ideal source to simulate this radiation would be an RTG itself. Certain problems would exist in the handling and storage of such a source in the laboratory not to mention its cost. However an RTG was

not available for this work so the problem did not arise. The basic

Numbers in parentheses refer to the list of references appended to this paper.

130
problem was to find a radiation source to do some preliminary work

on the most sensitive electronic components to determine the extent

of the problem.

Both reactor neutrons and neutrons produced indirectly by an

accelerator were considered. However they were both rejected be- cause of the difficulty in controlling the accompanying gamma flux. In addition reactor fluxes are somewhat high for this problem while accelerator produced neutron fluxes are so low that an inordinate amount

of manpower would be required to reach the desired levels. The proposed deep space application using the RTG power system

will generate a total fluence on the order of 4 x 10 neutrons/cm^ at the electronics package during a 5 yr mission. The gamma fluence

depends on a number of factors including the purity of the fuel element

and the gamma build-up. It has been estimated to be 1 x lO^^ photons/

cm^

In view of the nature of the radiation damage and other practical

considerations the source had to satisfy the following criteria:

1. The level be such that the fluence produced during a typical

deep space mission could be achieved in a reasonable time.

2. A gamma level that was low enough so that over the period of irradiation its effects on the MOSFET would not be dominant

over that of the neutron.

3. The neutron energy spectrum be similar to that of the RTG.

4. The activity of the source should be low enough so that handling problems are not excessive.

Criteria 1 and 4 are in some contradiction but can be resolved

with safety in mind. This has the added benefit that although the irradiation will take longer it will be closer to real time and help to eliminate possible rate effects. NEUTRON SOURCE

An available source that seemed to satisfy all of our criteria

pretty well was a cylindrical 9 curie source of Pu^^^-Be (Fig. 2). The reaction that produces the neutron flux is an (a, n) reaction where

the alpha particle emitted by the Pu^^^ reacts in the following wayr^^^

Be^ + He^ —^C^^ + n + 5.708 MeV

The neutron spectrum that results is believed to be (Fig. 3) similar

Numbers in parentheses refer to the list of references appended to this paper.

131
enough to that of the RTG to permit an unambiguous assessment of the radiation effects of the neutron on the MOSFET, The gamma spectrum

has been measured and found to have peaks at 40 KeV and 2,4 MeV.

The major problem in using a source of this size and geometry is in the dosimetry used to determine the neutron and gamma fluences at the point where the devices were irradiated,

EXPERIMENT

For the first experiment a total of eight devices, two each of four

off-the-shelf commercial p-channel enhancement types from three

manufacturers, were used. The devices were mounted on a circuit

board so that the top of the device can was against the source cylinder

(Fig. 4). In this way the maximum neutron flux was encountered. The

devices were all irradiated with a bias of -10 V on the gate during

irradiation. The gate threshold voltage for each device was measured

periodically by an automatic data acquisition system. Four control

samples, one of each type used in the irradiation, were biased and

measured in the same way as the irradiated samples. This experiment was rxm for a total of 230 days, A second experiment was run

for 30 days to give a better idea of the data spread among the device

types,

A third experiment using a Co^° source for a pure gamma flux was used on the same device types to determine the portion of the damage that could be attributed to the neutrons in the original experi- ment.

FLUX DETERMINATION

Although many dosimetry techniques, including calorimetry, neutron long counter, thermoluminescent dosimeters, a nemo counter, ion

chambers, activation analysis, R-chambers and film, were used for various parts of the experiment, in the final analysis, only two of the methods were relied upon. The neutron flux was computed "^^^ from the

total emissivity supplied by the manufacturer. The gamma flux was

determined from film. Assuming that the calculated neutron flux is correct and that the neutron-gamma ratio remains constant away from

the can, the gamma flux computed at the surface of the can from the

ratio agrees fairly well with the measured value. The neutron flux

Numbers in parentheses refer to the list of references appended to this paper.

132
was taken to be 1.7 x 10^° n/cm^-day and the gamma flux at 5 rads/day.

RESULTS

The gate threshold shifts for a typical device type is shown in Figure 5. On the abscissa the calculated neutron fluence and the measured gamma fluence is plotted in addition to the exposure time on

a logarithmic scale. Curves are shown for both the neutron experiment and the Co^^ experiment with the latter being plotted on the same

scale as the measured gamma contribution of the neutron source. The

curves for almost all of the neutron exposed devices fit an equation of the form

AVt = K ((D)-^^ where A vt is the change in the gate threshold voltage and ^ is the fluence. On the other hand all of the Co^° data show that the change in gate threshold voltage is linear with respect to fluence,

DISCUSSION

The results of using this experimental neutron source show that while the total number of gamma photons emitted is somewhat less than the number of neutrons their relative effect is large but not large

enough to mask the effect of the neutrons. From this data one can

assess the radiation damage of the MOSFET due to neutrons and can determine with some degree of confidence the problems that will be

encountered on a deepspace mission. It is clear that at the radiation

levels anticipated for the configuration shown the neutron damage to the spacecraft package will not be significant, while the gamma level

is at the threshold of damage. Another consideration that must be

taken into account for determining the vulnerability of the electronics

is the electronic application to which the MOSFET devices or inte- grated circuits will be put. For most digital applications a shift of one

volt in the characteristic can be tolerated. For some t3rpes of linear

applications lower thresholds must be used.

There is probably some small amount of displacement damage due

to the neutrons in these devices. The fact that the neutrons showed a

slightly non-linear relationship and that the damage could not be an- nealed at temperatures that will ordinarily anneal ionization damage in MOSFETs shows that another type of damage is present.

133
CONCLUSION

While the Pu^^^-Be source served nicely for preliminary experiments the results show that for practical RTG configurations the gamma effects will dominate. From the nature of the curves it seems

reasonable to expect to be able to predict the radiation effects of the RTG for other configurations once the expected neutron and gamma

fluxes are determined.

A number of other radioisotope sources have been considered but they do not appear to be any better than the one used. Short half-life, excess gamma and poor neutron spectrum match seem to be the major

problems.

ACKNOWLEDGEMENTS

The author would like to thank Mr. Kirk King for his extensive

dosimetry efforts, Mr. William Larimer for handling many of the ex- perimental details, especially toward the windup of the experiment,

and Mr. Jule Hirschfield for many useful discussions concerning the simulation of the radiation environment.

134
REFERENCES

(1) J. P, Mitchell, "Radiation-induced spacecharge buildup in MOS

structures," IEEE Trans. Electron Devices, vol. ED-14, pp. 764-

774, November 1967

(2) E. H. Snow, et al., "Effect of ionizing radiation on oxidized silicon surfaces and planar devices," Proc. IEEE, vol. 55, pp. 1168-1185,

July 1967

(3) P. A. Newman and H. E, Wannemacher , "Radiation induced peaking

effect in open geometry p-channel MOSFETS," IEEE Trans. Nuclear Science, vol. NS-14, pp. 299-302, December 1967

(4) G. Messenger, et al., "Displacement damage in MOS transistors,"

IEEE Trans. Nuclear Science, vol, NS-12, pp. 78-82, October 1965

(5) A. A. Witteles, "Neutron irradiation effects on MOSFETS," IEEE

Trans. Nuclear Science, vol. NS-15, December 1968

(6) A. G. Stanley, "A model for shifts in the gate turn-on voltage of insulated-gate field effect devices induced by ionizing radiation,"

IEEE Trans. Electron Devices, vol. ED-14, pp. 134-138, March

1967

(7) J. B. Marion and J. L. Fowler, "Fast Neutron Physics," Inter- science Publishers, Inc., New York, 1960, p. 6.

(8) M. E. Anderson and W. H. Bond, Jr., "Neutron spectrum of a Plutonium-Beryllium source," AEC R&D Report, MLM-1131,

UC-34 Physics, TID-4500 (18th Ed.), September 1964

(9) T. Rockwell, "Reactor Shielding Design Manual," D. Van Nostrand

Company, Inc., Princeton, N.J., 1956, p. 360.

135
136
137
SNOdlDBN JO d3aiAinN 3AllV13d

138
Fig. 4—Neutron- MOSFET Experiment

139
Pu-Be EXPOSURE (days)

1 10 100 1000

1000 -

100 —

10 100 1000 10000

GAMMA FLUENCE (PADS)

I L \ L_

10^1 10^2 I0l3 10^^

NEUTRON FLUENCE (n/cm2)

FLUENCE EXPECTED AFTER 5 YR RTG MISSION

Fig. 5—Typical Neutron-MOSFET Results

140
Paper Ho. 9

IN SITU VACUUM GAUGE CALIBRATION BY THE REFERENCE

TRANSFER METHOD

L.D. Blado^ and R.H. LilienkampREFERENCE:

Blado. L.D. and Lilienkamp. R.H.. ""In Situ Vacuum Gauge Calibration by the Reference Transfer Method." ASTM/IES/AIAA Space Simulation Conference.

14-16 September 1970. ABSTRACT:

Application of the Reference Transfer Method to in situ calibration of

ionization gauges was evaluated in order to ascertain the effectiveness of the method when applied in a manner determined both by the commercial avail- ability of equipment and b\- the restraints of an operating space environment

simulation facility. A prototype Reference Transfer Calibrator (RTC) was

assembled from standard vacuum components. Transfer ratio determinations were made in the range 10—4 to 10^8 torr. These experimental measuremients were then compared with theoretical calculations, thereby verifying the feasi- bility of the technique.

KEY WORDS: Vacuum Gauge Calibration, Reference Transfer Method

The problem of calibrating vacuum gauges for accurate measurement in pressure ranges below 10-4 torr has long been recognized. A technique for

calibration in this range, proposed by Charles Morrison''-^ ^ in 1966, is the Refer- ence Transfer Method. This method, by virtue of its simplicity and the fact that it relies upon a single pump, is idealh" suited for in situ calibrations. The

purpose of the work described here was to assemble a prototype Reference

Transfer Calibrator (RTC) with commercially available hardware and to evaluate the efficacy of this technique for in situ operations.

^Test Engineer, Engineering Laboratories, McDonnell Aircraft Company,

St. Louis, Mo.

-Senior Group Engineer, Engineering Laboratories, McDonnell Aircraft Company, St. Louis, Mo.

Superscript numbers in parentheses refer to the list of references appended

to this paper.

141
The lack of an acceptable pressure standard below 10—4 torr requires

that a means be found to relate lower pressures in some known manner to pressures in a region where an acceptable standard does exist. Reference

transfer, like other calibration techniques for this pressure range (2,-^,4,5) takes advantage of the well-known fact that a molecular flow of gas through an

orifice will produce a pressure ratio across that orifice. A hypothetical two-orifice system (Figure 1) can be used to determine

the pressure ratio in the general case. The pressure ratio across the orifice OaB ^'sn be obtained for equilibrium conditions with flow through the chamber. The flow into B is (Pa - Pr) FaB^ where Pa and Pb are the pressures,

in torr, of Chambers A and B, respectively, and FaB is the conductance

(liters per second) of the orifice OaB- The flow out of B is (Pb - Pc) FBC^

where Pc is the pressure of Chamber C and Fbc is the conductance of the

orifice ObC- Thus,

(Pa - Pb) Fab = (Pb - Pc) Fbc (D

Rearranging Equation 1 gives

Pa/pb = (Fab + fbc)/fab - (Pc/Pb)(Fbc/fab) (2) From Equation 2, it follows that if Pc is small with respect to Pb, the second term is negligible, and

Pa/Pb = (Fab + fbc)/Fab = u (3) This ratio can now be used to generate a series of pressures having a known

relationship. Since we know the interrelationship of all series members, the comparison of one element to a standard will yield a standard value for all the elements. If we let Pr be the standard and Pn be the Nth pressure in the series, we can write PN=PrUN-R (4) The foregoing development does not consider background gas loads which

may be neglected in a well-baked calibration system but which could become

significant in a typical space simulator. With a gas load Qb in Chamber B, Equation 1 becomes

(Pa - Pb) Fab + Qb = (Pb - pc) fbc (5) Rearranging Equation 5 gives

Pa/pb = (Fab + fbc)/Fab - Qb/pbFab - (Pc/pb)(Fbc/fab) (6)

Similarly, for Chamber C with gas load Qc,

(Pb - Pc) Fbc + Qc = pcs (V) where S is the pumping speed in liters per second.

Superscript numbers in parentheses refer to the list of references appended

to this paper. 142
Rearranging Equation 7 gives

pc/pb = fbc/(s + fbc) + Qc/pb (s + fbc) (8) Combining Equations 6 and 8 gives

Pa/Pb = Ki -K2/PB (9) where

Ki = (Fab + fbcVFab - F2bc/(SFaB + FaB FbC) (10)

and

K2 = [Qb (s + Fbc) + fbcQc]/fab(s + fbc) dD

The first term of Equation 10 corresponds to U and the second term

is neghgible when the pumping speed is large. (The pumps used in our

experimental work had effective pumping ranging from 400 to 25,000 liters per second, giving errors due to this term of 0.5 to 0.03 percent)

The first term of Equation 1 1 corresponds to the ultimate pressure of Chamber B and the second term is approximately equal to Fbc/FaB multiplied by the ultimate pressure of Chamber C. Thus, if Pb remains large compared to K2, Pa/PB remains constant, and if the pumping speed is large, Pa/PB equals U. The relative error in the

pressure ratio due to outgassing can now be expressed as dU _ K2

U " PbKi ^^^^

This relation is shown in generalized form in Figure 2. The relative error in a final calibration can be determined from Equation 4. Letting Pn be the Nth pressure in the series, we had

pn = pr un-r

Therefore,

^ = + (N-R) M (13)

The first term in Equation 13 gives the error contribution due to the stand- ard chosen. Bromberg'''^^' has shown that a capacitance manometer is accurate

to within 0.5 percent in the 10—4 to 1 torr range. From Figure 2 and Equation 13, it follows that the closer we get to the chamber's ultimate pressure and the more steps we need to get there, the

larger the error in the calibration will be. In the process of an experimental

Superscript numbers in parentheses refer to the list of references appended

to this paper. 143
determination of U, an approximation of the constant K2 can be made by

observing the pressure at which a drastic change in the pressure ratio occurs. The prototype RTC shown in Figure 3 was assembled from standard

1.5-inch vacuum components, such as crosses, valves, and flexible couplings,

employing sexless flanges with copper seals. The orifices OaB and ObC

shown in Figure 4 are 0.32 and 0.95 cm in diameter, and have conductances

of approximately 1 and 9 liters per second, respectively. The orifices were

given a 100*^ chamfer to approximate a zero thickness condition. The transfer gauge is a Bayard Alpert type ionization gauge operated at an emission

setting of 0.9 ma for a rated sensitivity of 10 amps per torr. Before each

data gathering attempt, the entire system was baked at 400*^F for a mini- mum of 24 hours. Figure 5 shows a schematic of the RTC attached to a space simulation chamber designated as Chamber C. The two chambers (A and B) of the

calibrator are connected through orifice OaB and can both bt valved in- dependently into a transfer gauge Gt. Chamber B is connected to Chamber C

through a second orifice ObC- Chamber C can also be valved independently

into the transfer gauge. The gas source was a regulated supply of room

temperature dry nitrogen admitted into Chambers A and C through separate variable leak valves attached to each of those chambers. The Baratron

capacitance manometer used as a pressure standard attaches to the remaining flange on the cross designated as Chamber A. A typical calibration sequence proceeds in the following manner:

( 1 ) With Valves A, B, and C open and variable leak Valves La and

Lc closed, evacuate the entire system until a stable equilibrium

is reached.

(2) Close Valves A and C. Open Valve La to admit gas to Chamber A. Monitor Chamber B pressure until it is at least 1 decade higher than

its original value. Record the transfer gauge reading.

(3) Close Valve B and open Valve C. (4) Using Valve Lc, admit gas to Chamber C until the pressure as seen by the transfer gauge is the same as that recorded in Step 2. Re- cord the reading of Gauge Gc; this is the first calibration pressure. Close Valves C and Lc-

(5) Open Valve A and record the transfer gauge reading. Close Valve

A.

(6) Open Valve B. Increase the gas flow through La until the pres- sure in Chamber B is the same as the value recorded in Step 5. Record the transfer gauge reading.

Steps 3 through 6 in the above procedure are repeated until the pres- sure generated in Chamber A is high enough to be compared to the available standard. Once this comparison is made and an absolute value for this pres- sure is known, we can extrapolate to the other members of the series, and

the calibration is complete.

Obtaining the desired pressures in Chamber B by the control of the gas flow through the leak valve La was very difficult. To obtain the fine control needed, the supply pressure to the leak valve was reduced to 1 torr. The majority of our effort has been in the experimental determination of the pressure ratio U. For this work the transfer gauge is used to determine

144
the pressure in Chambers A and B and from these data the pressure ratio

is calculated. The transfer gauge need not be calibrated for these measure- ments since the determination of U requires that the measured pressures be

correct relative to each other. The symmetry of the valving system that connects the transfer gauge to the chambers, coupled with the well established linearity of the Bayard-Alpert type ionization gauge over the pressure range under consideration, ensures a high degree of confidence in the in- dividual ratio determinations. The results of two determinations of the pres- sure ratio U, with the RTC mounted on two different space simulators, are given in Table 1 . The first set of data was obtained on a system having a pumping speed of 25 x 10^ liters per second during the space conditioning of spacecraft components. The second set of data was obtained on a system

having a pumping speed of 400 liters per second with the chamber empty.

From both sets of data we can see the change in the pressure ratio as the

pressure in Chamber B approaches the ultimate pressure of Chamber C. For these chambers the ultimate pressures were 1 x 10-7 and 1 x 10-8

torr. K2 for these chambers was approximately 10-6 and 10-7 torr, respectively. Figure 6 presents the variation of the pressure ratio with Chamber B

pressure. The data from Table 1 is also shown on this figure. We have operated the reference transfer system somewhat differently >• than Morrison. He used a flow measurement to establish the reference pressure, but we are using the capacitance manometer for this purpose. We believe that this instrument is accurate enough for most work and is far simpler

to operate than a flow measuring device or a McLeod gauge. For our particular system a pressure ratio of 10 was selected. With this ratio the number of steps in each calibration sequence is kept reasonably

small. The ease of obtaining the reference pressure allows rerunning the se- quence several times to obtain intermediate points. Because the experimental determination of the pressure ratio is limited by our ability to measure the ion current in the transfer gauge and our control of the operating voltages and currents supplied to the transfer gauge, it may become necessary to use the calculated values of the pressure ratio obtained from calculated values of conductance based on measurements of the

orifices to obtain the accuracy expected.

The following main conclusions can be drawn from this work:

1) Reference transfer is well suited to an in situ calibration. Compactness and design simplicity coupled with the use of a single pumping

system are a near-ideal set of characteristics for in situ calibration. 2) Relative calibrations in a facility with several vacuum chambers can be achieved quite easily. An RTC such as the one used in this work can be

readily assembled from available hardware. Its operation does not require the high level of vacuum technique sophistication characteristic of even most re- lative calibrations. Our prototype RTC was mounted on two operating vacuum chambers without making any attempt to bring those chambers to

their maximum operating efficiency. In one case, the RTC was mounted and

the pressure ratio determined while the chamber was being used to degas

several spacecraft components. The fact that calibrations can be performed

to within 1 decade of the chamber ultimate with a reasonable degree of

145
accuracy, regardless of what that ultimate is, can prove to be a valuable

capability. In addition, the compactness of the RTC makes transport re- latively easy. Thus, it is feasible to calibrate all the vacuum systems of a company relative to each other, regardless of plant location, by shipping an RTC from one plant to another.

3) The absolute accuracy of a reference transfer calibration can be improved dramatically by careful operations, using accepted high vacuum

techniques. Equations 13 and 14 show that minimizing the error in the

transfer ratio by reducing the background gas load (i.e., lowering the system

ultimate pressure) will permit the accurate calibration range to be extended

into the ultra-high vacuum region. Morrison has estimated that the error

in the transfer ratio can be kept below 1 percent in the 10^9 torr range, thus allowing the operator to fit the degree of effort involved in a cali- bration to the degree of accuracy he desires. 4) The flexibility of the Reference Transfer Method is a definite advantage. Any adequate standard that is valid in the 10~4 torr range may be used.

References

1 Morrison, Charles, F. "Reference Transfer Method of Vacuum

Gauge Calibration," Journal of Vacuum Science and Technology,

Volume 4, Number 5, September/October 1967.

2 Choumoff, P.S. et. al, "Absolute Calibration of Ion Gauges by Means of a Bakeable Adjustable Conductance, " Journal of Vacuum Science and Technology, Volume 7, Number 1, January/Februai-y 1970.

3 Florescu, N.A., "Reproducible Low Pressures and Their Application to Gauge Calibration," Transactions of the 8th Vacuum

Symposium and 2nd International Conference, Volume 1, pp. 504-510. 4 Roehrig, J.R. and Simons, J.C., "Accurate Calibration of Vacuum

Gauges to 10—9 Torr," Transactions of the 8th Vacuum Symposium and 2nd International Conference, Volume 1, pp. 511-518.

5 Normand,C.E., "Use of a Standard Orifice in the Calibration of Vacuum Gauges," Transactions of the 8th Vacuum Symposium and 2nd International Conference, Volume 1, pp. 534-543.

6 Bromberg, J. Phillip, "Accurate Pressure Measurements in the Region 0.1 to 1000 millitorr; The Intercalibration of Two McLeod

Gauges and A Capacitance Manometer," Journal of Vacuum

• Science and Technology, Volume 6, Number 5, September/October

1969.

146
Table 1 - Determination of Pressure Ratio

High Pump Speed; Chamber Ultimate 1 x lO^^ Torr Lnamber A Lnamber b Rat io 1.70 X 10-6 3.45 X 10-7 4.9 2.15 X 10-6 4.15 X 10-7 5.1 2.80 X 10-6 5.15 X 10-7 5.4 4.70 X 10-6 6.95 X 10-7 6.7 5.10 X 10-6 7.20 X 10-7 7.0 5.30 X 10-6 7.40 X 10-7 7.2 9.80 X 10-6 1.15 X 10-6 8.5 1.05 X 10-5 1.20 X 10-6 8.7 1.40 X 10-5 1.50 X 10-6 9.3 2.35 X 10-5 2.45x10-6 9.7 2.40 X 10-5 2.40 x 10-6 10.0 3.95 X 10-5 4.00 x 10-6 9.8 5.40 X 10-5 5.40 x 10-6 10.0 6.10 X 10-5 6.05 X 10-6 10.0 7.10 X 10-5 6.80 X 10-6 10.4 9.50 X 10-5 9.10 X 10-6 10.4 9.70 X 10-5 9.40 X 10-6 10.3 1.40 X 10-4 1.40 X 10-5 10.0 z.OO X 10 ^ T A A 1 A— 5 1 A A

1 (J.U

-3 on V 1 0-4

J . 7U X 1 U

3 XO V 1 0-5 J.OU X 1 u 1

1 0U. R0 4.40 X 10"4 4.30 X 10-5 10.2 5.40 X 10-4 4.95 X 10-5 10.9 9.10 X 10-4 9.00 X 10-5 10.1 Low Pump Speed,' Chamber Ultimate 1 x 10 8 Tr,rr 1 orr Chamber A Lnamber b Ratio

4.60 X 10-^ 1.00 X 10-8 4.6 1.70 X 10-7 3.56 X 10-8 4.7 3.10 X 10-7 Q 5.35 X 10-8 5.7 5.00 X 10-7 7.50 X 10-8 6.6 7 6.60 X 10-7 0 9.20 X 10-8 7.1 1.50 X 10-6 2.00 X 10-7 7.5 4,80 X 10-6 5.00 X 10-7 9.6 7.50 X 10-6 7.55 X 10-7 9.9 3.00 X 10-5 3.00 X 10-6 10.0 6.60 X 10-5 6.50 X 10-6 10.0 9.40 X 10-5 9.20 X 10-6 10.2 4.00 X 10-4 4.00 X 10-5 10.0 8.25 X 10-4 8.20 X 10-5 10.0 147


Pressure

Standard

Variable Leak

Chamber A

Chamber |o

Variable Leak Lq

BC

^AB

Valve A

Valve C

Chamber C

1 r

Figure 5 - Reference Transfer Comparison System Schematic
Paper No. 10

METHODS OF DETERMINING RESIDUAL GAS COMPOSITION FROM

RESIDUAL GAS ANALYZER DATA

R.H. Lilienkampl

REFERENCE: Lilienkamp, R.H., "Methods of Determining Residual Gas

Composition From Residual Gas Analyzer Data," ASTM/IES/AIAA Space

Simulation Conference, 14—16 September 1970. ABSTRACT: Three methods of determining residual gas composition from

Residual Gas Analyzer (RGA) data are compared using a computer simulation of an RGA to generate the data for analysis. Typical vacuum system

gas mixtures and errors from shot noise were simulated. Errors were also introduced into the calibration data used to interpret the RGA data. The

first method is an iteration technique which proved to be inferior to the other methods. The second is a matrix inversion method which worked

well for the gases used in this study but which must be evaluated for each

gas mixture. The third method is a multiple regression analysis giving least squares fit to the data. This method uses all of the data from the RGA

and is usable with any mixture of gases. The effect of the calculation of

the residual gas composition is also shown.

KEY WORDS: Mass Spectrometry, Residual Gas Analysis Data Reduction,

Residual Gas Analysis Computer Simulation, Space Vacuum System Gas

Analysis

Introduction The residual gas analyzer (RGA) is a very useful instrument for monitoring

the performance of the vacuum system of a space simulator, and the reaction of

the test specimen to the environment. In general, these instruments are used for

qualitative analysis because of the complexity involved in analyzing the data to obtain quantitative results. A recent application of the residual gas analyzer

required quantitative results, and a method of computerizing the system was

1 Senior Group Engineer, Engineering Laboratories, McDonnell Aircraft Company, St. Louis, Mo.

151
developed. n) In this application only five gases were monitored and three of

these had unique peaks which simplified the analysis. For the more complex gas mixtures which can be expected in a space

simulator, a more general method of data reduction would be required. The

following methods of analyzing residual gas data have been investigated:

iteration technique, matrix inversion method, and multiple regression analysis. In the iteration technique( an estimate of the partial pressure of each

gas in the mixture is made. These estimates are used to calculate a mass spec- trum which is then compared to the original data. From this comparison, cor- rections are made to the original estimates and the process is repeated until two

successive estimates are sufficiently close. In the matrix inversion method, a subset of N equations in N unknowns

is selected and matrix algebra is used to solve these equations. A logical choice of equations to be used in this subset was made by basing the choice on the cracking patterns for each of the gases expected to be in the mixture.

Multiple regression analysis(3) gives a least square fit to the data, using all the data from the RGA. With this technique it is possible to weight the data so

that the data with larger errors is not as important as the data with smaller

errors. The RGA Model

To evaluate the methods of data reduction, a computer simulation of a mass spectrometer was used to generate mass spectra for analysis. This model

assumes that there is linearity and no interaction between gas species. These are the same assumptions that are used in the data reduction techniques. Gas mixtures typical of vacuum system operations(4) were used to generate

the mass spectra. Eight gases (hydrogen, methane, water vapor, carbon monoxide, nitrogen, oxygen, propane and carbon dioxide) were possible in these mixtures. The proportions of the mixtures, listed in Table 1 , correspond to those present in an unbaked vacuum system, a vacuum system during bakeout,

and a well-baked vacuum system.

Cracking patterns and sensitivities published for mass spectrometers(5) were

used in the model to calculate the peak heights (i.e., ion current) for these mixtures. The cracking patterns used are tabulated in Table 2. Shot noise effects were simulated by adding errors, which were random

both in magnitude and sign, to the spectra. The shot noise was proportional

to the square root of the peak height. A proportionality constant was chosen

to make the shot noise equal to 1 percent of the peak height for the major

peak of a 100 percent concentration of a gas having a sensitivity of 100. Provisions were made to vary the constants to simulate varying amounts of shot

noise.Random errors were also introduced into the calibration data (i.e., sensitivity and cracking pattern) used to reduce the mass spectra data. These errors were directly proportional to the magnitude of the calibration data to which

they were applied.

Superscript numbers in parentheses refer to the list of references appended

to this paper. 152
The calibration errors and the magnitude of the shot noise were varied to evaluate their importance in obtaining a good gas analysis. For each combination of gas mixture, shot noise, and calibration errors, five different sets of noise data were generated. These noise data were both added to and subtracted from

the basic mass spectra, thus producing 10 spectra to be analyzed. For each pair of spectra, a different set of calibration errors was used. Thus, a small statisti- cal sample was obtained to compare the techniques.

The matrix equation for the simulated mass spectrometer is (S X lAl X IP) ± [N = Ih (1) where h, a column matrix of 22 rows, is the peak height. P, a column matrix of eight rows, is the composition of the gas. A, a rectangular matrix of 22 rows

and eight columns, is the true calibration data for the spectrometer.

S is a constant used to convert relative sensitivities to absolute sensitivities; its value was selected so that the major peak of a 100 percent concentration of a gas having a sensitivity of 1 00 equals 1 O^^

The noise matrix N is a column matrix of 22 rows, and is added to or subtracted from the true peak heights. Each element of this matrix is related to the true peak height H by

Nj j = (Random Number) kHj j (2) where the Random Number is random in sign and magnitude between 0 and 1, and k is the proportionality constant. The matrix equation for calibration data with errors is [A] = tC] + [E] (3) where A is the calibration data with the errors, C is the true calibration data, and E is the error. All three matrices are rectangular with 22 rows and eight columns. Each element of the E matrix is related to the corresponding elements of

the C matrix by

Ej j = (Random Number) kjCji (4) where the Random Number is random in sign and magnitude between 0 and

1, and ki is a proportionality constant.

Iteration Technique

The values of the partial pressure are estimated by calculating the value from each peak height and calibration factor in the spectrum for a given gas and using the lowest of these values as the partial pressure. These partial pressures are used with the calibration factors to calculate

a spectrum which is subtracted from the original spectrum to obtain an error spectrum. A correction to the partial pressures is obtained from the error

153
spectrum in a manner similar to that used to make the first estimate. The second approximation is obtained by adding the correction to the

original estimate.

Iterations are made by calculating a new spectrum which is used to cal- culate another new error spectrum from which new corrections are calculated and then added to the current estimate. The iterations are continued until the peaks in the error spectrum become small compared to the corresponding

peaks in the original spectrums or until the error spectrum ceases to change

with successive iterations. In the initial evaluation of this technique the results from each iteration were printed, and based upon a comparison of these results with those of the previous iteration, a decision was made by the operator to continue or stop

the iteration process. From these data it was concluded that it is unlikely that

all the peaks in the error spectmm will become small compared to the corre- sponding peaks in the original spectrum even if "small" is defined as 1 percent. The Matrix Inversion Technique

Matrix algebra is a conventional method for solving a set of N simultaneous

linear equations in N unknowns. Starting with an equation set X [Pi = [hj (5) to solve for the Pj's, the inverse of matrix A is used and

[Aj,i]-1 X [hj = [Pi (6) To have an inverse, the matrix A must be square and must not be singular.

The original matrix A was not square; therefore, to use this technique, a subset of N equations must be chosen to produce a N x N matrix. Out of the complete set of 22 equations, it is possible to choose 341,088

subsets of eight equations. This number can be reduced by following two

simple rules:

( 1 ) The matrix must not be singular. Thus, if any rows are either equal

or proportional to each other, only one can be used in the matrix. Examination of the matrix in Table 2 reveals that the 22 equations are reduced to 14 and the number of possible subsets is reduced to 3003.

(2) The matrix must include the major peak or at least one prominent

peak in the spectrum of each gas. Using the major peaks gives seven equations

corresponding to m/e 2 (hydrogen), 16 (methane), 18 (water vapor), 28 (carbon

monoxide and nitrogen), 29 (propane), 32 (oxygen), and 44 (carbon dioxide).

The remaining equation must further define the carbon monoxide and

nitrogen concentrations, and thus it is limited to m/e 12,13,14 and 30. The m/e 12 and 13 peaks are formed by methane and carbon dioxide in addition to carbon monoxide. The m/e 14 peak is formed by both carbon monoxide and nitrogen as well as methane. The m/e 30 peak is formed by only carbon monoxide, but is quite small. The m/e 14 peak was chosen for this study. 154
A more mathematical approach is presented by Stanton(6). The N rows

of the matrix are vectors in N dimensional space. If the vector is perpendicular

to all other vectors, it is ideally suited to be in the subset. If the angle between two vectors becomes small, then the two equations are ill-conditioned and only one can be used in the subset. The angle between the two vectors is given by

Cos2 0j =

where j and k are the two rows of the matrix being tested. For the 22 row

matrix used in this study, there are 231 values of Cos2 dj^\^. Table 3 presents the values of the Cos2 0j ]^ for the seven major peaks with

respect to each other. From these data, the worst pair are m/e 28 and 29. The

Cos2 0j 1^ is 0.144152, which means is about 68 degrees. Table 4 presents the values of Cos20j for m/e 12,13,14 and 30 with re- spect to the seven major peaks. The largest Cos^O j^k occurs with m/e 13, the next largest with m/e 12, and the third largest with m/e 30. The m/e 14 peak,

while having the smallest values of Cos20j j^, has two values larger than 0.3. To

determine if it is better to have two pairs of equations with some ill-conditioning or one pair more ill-conditioned, the subset of equations used in the matrix in- version was modified by removing the data for the m/e 14 peak and adding the data for the m/e 30 peak.

Multiple Regression Analysis

Multiple regression is used to obtain the best fit of a set of observations of independent and dependent variables to an equation of the form

y = bo + bi Xi + b2 X2 + . . . bM Xn (8) where y is the dependent variable and the X's are the independent variables. The

b's are the constants to be determined for the set of observations. For residual gas analysis, the y's are the peak heights, the b's are the pressures, and the X's

are the calibration constants. Thus, Equation 13 becomes

hj = ho + Pi Aj 1+P2 Aj,2 + . . . Pn Aj,N (9) which, except for the term, is the same as Equation 5. The least squares fit determines the values of the P's in such a way that the error is minimized and

results in N simultaneous linear equations in P. These N simultaneous linear equations can be solved by any appropriate method. The computer technique

described by EfroymsonC^) and used by Tunnicliff and Wadsworth(3) solves these equations by the Gaussian elimination method. At each step in the elimiSiiperscript numbers in parentheses refer to the list of references appended

to this paper.

N

2

i=l

(ai,j'iak,iJ N ^ N 2 ^ ^jti 2 ak,i 1=1 1=1 (7)

155
nation, either a variable is removed or a new variable is added to the regression.

The criteria used to select the variable to be added to or removed from the re- gression are as follows:

1. If the various contributions of a variable in the regression are insignifi- cant, this variable is removed. If no variable is to be removed, then the following

criterion is used.

2. The variable providing the greatest variance reduction is added to the

regression if the variance reduction is significant. If the variance reduction is not significant, the regression is complete and the remaining variables are set equal to zero. The application of these criteria to an actual problem is shown in Table 5, which presents the results for each step in the regression. The gas mixture used

corresponds to that for a baked system, and a 0.1 percent calibration error and

a 0.1 percent shot noise contribution to the maximum signal are assumed. In the first step the hydrogen partial pressure is the most significant. In the second

step carbon monoxide partial pressure is added to the regression but with a value about twice that actually present. At Step 5, when nitrogen partial pressure is added to the regression, the value of the carbon monoxide partial pressure is re- duced to a more reasonable value. The regression is completed in the seventh

step, since adding the oxygen partial pressure will not produce a significant vari- ance reduction. The oxygen partial pressure is then set equal to zero. In ad- dition to the calculation of the partial pressures, the method also calculated the standard error for each of the partial pressures and for the peak height. The

method also includes provisions for including weighting factors, which Tunnicliff and Wadsworth(3) have done. They also have incorporated a method of rounding off the partial pressures based on standard errors. The majority of the work reported in this paper does not use weighting factors.

Results . , I A comparison of the gas composition calculated by analyzing the same data with the three techniques is presented in Table 6. A comparison of the errors

in mass spectra computed from these results is presented in Table 7 along with

the sum of the squares of these errors. The iteration method produced the most

inaccurate data reduction of the three methods. The errors shown in Table 7 for the peak height calculation from the iteration method are zero for nine of

the peaks (m/e 2,12,15,18,29,30,32,34 and 38). The 32 and 34 peaks are re- dundant. Therefore, the method solved eight simultaneous equations in eight

unknowns. Of the eight peaks, six (m/e 2,15,18,30,32 and 38) are unique to one of the gases in the mixture. The Cos2 ^j^^was calculated for these eight

equations. For m/e 1 2 and 30, the value of Cos2 6 was 0.3 1 2474, and for m/e

29 and 38, the value was 0.999741 . All other values were very small or zero. Following the procedure of Sherrell and Mathews^^)^ ^he smaller peaks were deleted and the iteration was repeated. The results are presented in Table

8. The results were unchanged for four of the gases (hydrogen, methane, water

Superscript numbers in parentheses refer to the list of references appended

to this paper. 156
vapor and oxygen). For nitrogen, propane and carbon dioxide, the results im- proved, but for carbon monoxide, the error increased. A comparison of the computed spectra from this composition with the original spectra reveals that the sum of the squares of the differences has increased from 1.61966 x 10—12

to 1 .7599 X 10—12. In this iteration only seven of the peaks, m/e 2, 1 5, 1 8, 29,

32, 34 and 39, produced zero error. Again, 32 and 34 were redundant, so that, effectively, only six peaks produced zero error. However, the m/e 14 and 44

peak errors were very small; these peaks had been controlling the iteration of

nitrogen and carbon dioxide, respectively, until the m/e 29 error went to zero. Again, the Cos2 9j j^was calculated for these eight equations. The value of Cos2 d for m/e 14 and 15 was 0.402002, and for m/e 29 and 39 it was

0.999741, indicating that this choice of equations was worse than the first in this particular example.

Deleting the small peaks from the calculation removed the unique peak

for carbon monoxide. This resulted in the number of iterations increasing from

2 to 28 in this example. In general, removing the small peaks resulted in an in- crease in the number of iterations. In some cases the iterations would not converge for all gases, but would oscillate between two values. In one of these examples the carbon monoxide varied between 6.5487 and 6.66516, and the

nitrogen varied between 9.26273 and 7.42821. Their actual values were 7.6 and 7.5, respectively.

The choice of subsets of equations to be used in the matrix method was

verified by comparing results from two different subsets. Both subsets included m/e 2,16,18,28,29,32 and 44, which are the major peaks in the cracking patterns of the eight gases. One subset included m/e 14, while the other included m/e

30. The results for hydrogen, methane, water vapor and oxygen were not

affected by this equation, but the results for carbon monoxide and nitrogen were the most affected. Results from one set of calculations are presented in Table 9. These results are typical and clearly favor the choice of the subset which includes the m/e 14 peak. These results illustrate the importance of the choice of the subset and the limitations of the matrix inversion to technique. For these mixtures of gases, the matrix technique using the m/e 14 peak

and the multiple regression technique gave about equally valid results. Rounding off the results to a reasonable number of significant figures is required. The

round-off technique of Tunnicliff and Wadsworth(3) was tried on several calculations and appeared to work reasonably well. With this technique the last significant figure should have an accuracy of approximately one to five digits. Some adjustments are required to account for calibration errors. Weighting of the data was investigated by using two weight factors, one

proportional to the square root of the peak height and the other proportional

to the reciprocal of the square root of the peak height, except for the very

small peaks. The first weighting is based on the fact that the percent error

increases as the peaks become smaller, and the second weighting is based on the

Superscript numbers in parentheses refer to the list of references appended

to this paper.

157
fact that the absolute error decreases as the peaks become smaller. Neither of

these weighting factors improves the results enough to warrant their use. The effects of errors in the calibration factor and errors due to shot noise have been investigated by varying the magnitudes of these errors and observing

the results. For this investigation the same set of random numbers was used,

and only the order of magnitude of the errors was changed. The standard

errors of the peak heights are presented in Table 10. The effect of increasing

the noise is shown in the first row of data. A tenfold increase in noise pro- duces a tenfold increase in the standard error. In this first column of data the

effect of increasing the calibration error is shown, and again a tenfold increase

in the calibration error produces a tenfold increase in the standard error. When both errors are present, these relationships begin to break down and now

a tenfold increase in both errors is required to produce a tenfold increase in the standard error. This most likely is the result of some of the errors cancelling the effects of other errors. The errors in the calculation of the hydrogen concentrations are presented

in Table 1 1. The same basic pattern seen in Table 10 is repeated in Table 1 1, with an interesting variation occurring when the noise error and calibration error are of the same order of magnitude. Under these conditions the errors become smaller than any other in that row or column, furhter suggesting the cancellation of errors. A similar effect is noted in the errors in the calculation of the methane concentrations presented in Table 12. In this case, the minimum errors are obtained when the calibration error is ten times that of

noise error.

It would be expected that if both sources of error were eliminated, perfect solutions would be obtained. This is not necessarily the case, since the multiple

regression analysis technique employed uses a significance test for the addition of a variable. By eliminating the calibration and shot noise errors, one of the minor gases in each of the mixtures was excluded from the regression. For the unbaked system the nitrogen (0.79 percent) was dropped, for the system during

bakeout the oxygen (0.4 percent) was dropped, and for the well baked system

the water vapor (0.2 percent) was dropped. By modifying the regression tech- nique to include all variables, the concentrations of all gases were correctly obtained. With the inclusion of very small errors, these minor gases are no longer excluded from the regression and thus should not be of any concern in the

analysis of data taken from a real RGA.

Conclusions

The iteration method gives the illusion that it is using all of the data in the mass spectra, but it is actually picking out a subset of equations which can be

ill-conditioned and therefore produce quite erroneous results.

If a suitable subset of equations can be found, the matrix inversion tech- nique is fast and produces results consistent with the accuracy of the data

used. The technique is well-suited to on-line calculations for quick-look data.

158
The multiple regression analysis is the best of the techniques. This tech- nique does use all of the data in the mass spectra. The large number of gases which can be included in the multiple regression analysis makes possible the detection of unexpected gases in the mixture. This method also provides stand- ard errors for the results. Both the matrix inversion and the multiple regression analysis techniques

require reasonable accuracy in data and calibrations to obtain significant results. Experimental techniques to reduce the influence of noise on the data have been

developed. Employment of these techniques should provide the quality of data

necessary to achieve quantitative results with the residual gas analyzer.

References

1. R.J. Lilienkamp, H.F. McKinney, and D.I. Fiste,

a. "A Computerized Mass Spectrometer System", AIAA/ASTM/IES 4th Space Simulation Conference, Los Angeles, California, Sept. 8-10, 1969

(AIAA Paper No. 69-1016).

b. "A Computerized Mass Spectrometer System for Spacecraft Ground

Test", Journal of Spacecraft and Rockets, Volume 7, 1970, pp.1 10-1 1 1.

2. E.G. Sherrell and A.J. Mathews, "Data Reduction Techniques for Vacuum

Measurements by Mass Spectrometry", AIAA/ASTM/IES 1st Space Simulation Conference, Houston, Texas, Sept. 7-9, 1966, pp. 215-223.

3. D.D. Tunnicliff and P. A. Wadsworth, "A Stepwise Regression Program for Quantitative Interpretation of Mass Spectra", Analytical Chemistry,

Volume 37, 1965, pp. 1082-1085.

4. Anonymous, "Test Your Knowledge", Varian Vacuum Views, March/

April, 1969.

5. A. Cornu and R. Massot, "Compilation of Mass Spectral Data", Hayden

and Son Limited, London, England. 1966.

6. R.G. Stanton, "Numerical Methods for Science and Engineering,"

Prentice-Hall, New Jersey, 1961.

7. M.A. Efroymson, "Multiple Regression Analysis in Mathematical Methods

for Digital Computers". A. Ralston and H.S. Wilf, Eds., Wiley, New York.

1960.

159
Gas

Composition {%) Unbaked During Bakeout Well-Baked

H2 0 21.9 75.7 CH4 0.8 2.63 6.67 H2O 79.0 17.5 0.2 CO 16.0 21.9 7.6 N2 n 7Q 171/ R.J 7 q / .J OoU2 3.1 0.4 0

0 17.15 2.2 CO2 0.3 0.87 0.2

Table 1- Typical Residual Gases in Vacuum Systems

Gas H2 CH4 H20 CO N2 02 ^3^8 CO2

Column 1 2 3 4 5 6 7 8 M/e Row

1 1 1 n 0 0 0 0 0 0 2* 2 1000 0 0 0 0 0 0 12 3 0 8 0 47 0 0 0 69

13 4 0 29 0 1 0 0 0 9 14* 5 nu

71; u

Q0 u u u

15 6 0 754 0 0 0 0 0 16* 7 0 1000 9 17 0 36 0 94

17 8 0 12 211 0 0 0 0 0 18* 9 0 0 1000 0 0 0 0 0 26 10 0 0 0 0 0 0 53 0 27 11 0 0 0 0 0 0 316 0 28* 12 0 0 0 1000 1000 4 615 82 29* 13 0 0 0 12 7 0 1000 1 30 14 0 0 0 2 0 0 0 0 32* 15 0 0 0 0 0 1000 0 0 34 16 0 0 0 0 0 1 0 0 38 17 0 0 0 0 0 0 43 0 39 18 0 0 0 0 0 0 165 0 41 19 0 0 0 0 0 0 149 0 42 20 0 0 0 0 0 0 61 0 43 21 0 0 0 0 0 0 336 0 44* 22 0 0 0 0 0 0 402 1000

* Peaks used in matrix inversion technique. Table 2 - Cracking Patterns

160
m /e 16 18 28 29 32 44

2 0 0 0 0 0 0 16 1.41499x 10-'* 1.36289x 10-2 5.59433 X 10-^ 2.5438x 10-2 3.58113x10-2

18 0 0 0 0 28 0.144152 3.86872x 10-^ 2.88122x 10-2

29 0 9.44378x 10-2

32 0

Table 3 - Values of Cos^ 6 j ^ ^or the Seven Major Peaks

Major Peak

Candidate Peaks We 12 13 14 30

2 0 0 0 0 16 5.22741x 10-2 0.852571 0.39062 1.19731 x 10-2

18 0 0 0 0 28 0.202258 5.82893 x 10-2 0.304761 0.51036

29 8.28862x 10"^ 2.41448x 10-^ 5.40778x 10-^ 2.08387 X IQ-'*

32 0 0 0 0 44 0.621188 0.280373 • 0

Table 4 - Values of Cos^ 9 : ^ to Select the Eighth Peak

Gas Actual Step 1 Step 2 Step 3 Step 4 Step 5 Step 6 Step 7 H2 75.5 71.5293 73.7899 74.5893 75.3757 75.3722 75.3888 75.4246 CH4 6.67 5.85851 6.52495 6.63657 6.61605 6.65286 H2O 0.2 0.199035 CO 7.6 14.4206 14.6568 13.775 7.62461 7.51718 7.51973 N2 7.5 7.41275 7.56472 7.56177 O2 0 0 C3H8 2.2 2.22568 2.24266 2.18402 2.20731 CO2 0.2 0.200304 0.204763

Baked System Calibration Error 0.1 Percent Shot Noise 0.1 Percent for Maximum Signal

Table 5 - Results of Each Step in the Regression
Iteration Method Matrix Method Regression Method bas Actual Calculated Difference Calculated Difference Calculated Difference

H2 75.5 75.4954 4.64835x10"^ 75.4954 4.64835 xl0"3 75.4906 9.4038x10"^ CH4 6.67 6.70745 -3.74526 xl0~2 6.70143 -3.14347 x10"^ 6.70101 -3.10143x10-2 H2O 0.2 0.201330 -1.32975x10"^ 0.20133 -1.32975 xl0"3 0.198251 1.74917x10"^ CO 7.6 7.51186 8.81354x10"^ 7.76979 -0.169791 7.69521 -9.52123x10-2

N2 7.5 9.59221 9.59221 7.30213 0.197865 7.38658 0.113419 02 0 0 0 0 0 0 0 C3H8 2.2 2.17572 2.42754 xl0~2 2.18815 1.18469x10"^ 2.19320 6.80313x10"^ CO2 0.2 0.289488 -8.94883x 10"^ 0.198073 1.92661 x10"^ 0.194518 5.48246 xl0"3

Table 6 - Comparison of Composition Calculated by Three Techniques

m 'q Peak Height

Errors

Iteration Matrix Regression

1 4.18604x 10"^ D.DoUlil X lU

-9 C CQl 09 V 1 fl' b.oolUZ X iU-9 b./UbUb X lU

-9

2 1.96417 x 10"^ n

3 lOQl 1 Y 1 fl" j.jooi J X lU

-21

1 9'?77'? V 1 ri' l.LOlLJ X iU

-9

12 3.07419x10"^ 0 -4.42107 x10'

-9 -1.52820x10'-9

13 7.74480x10"^ -2.32419x10"

-9 -1.79984x10"

-9 -1.71235x10"-9

14 4.7688x10"^ -7.47515x 10"-8 -3.97047 x10'

-23 -2.33766x10"

-9

15 1.87109x10"^ 0 1.67871 x10'

-9 1.79594x10'-9

16 2.59784x10"^ 5.73121x10"

-9 6.35275x10"

-22 1.41230x10'

-9

17 5.11147x10"^ 9.64329x 10"-11 1.23143x10"

-10 4.49942 x10"

-10

18 1.00631 x 10"^ 0 2.64698x10"

-23 1.53895x10'

-9

26 7.62449 xlO"S 2.40678x 10"-9 1.98499x 10"-9 1.81382x10'

-9

27 4.51453x10"^ 1.12183x10"

-8 8.70354x 10"-9 7.68298x10"-9

28 1.15251 x10"^ -1.26839 X 10 -6

-4.23516x10"

-22 1.57104x10'

-9

29 1.50456x10"^ 0 6.35275x10'

-22 -2.91341x10"

-9

30 1.15625x10"^ 0 -3.97009x 10"-10 -2.82215x10"

-10

32 0 0 0 0 34 0 0 0 0 38 5.99302 xl0"S 0 -3.42343x10'

-10 -4.81274x10'

-10

39 2.33961 X 10"^ 4.15767x10"

-7 2.84495x 10"-9 2.31222x10"-9

41 2.13783 x10"^ 6.37356x10"

-9 5.18877 x10"

-9 4.70795x10"

-9

42 8.59945x10"^ 1.04179x 10"-9 5.56515x10'

-10 3.59576x10"

-10

43 4.77394x10"^ 9.56445x10"

-9 6.89204x 10"-9 5.80751 x10"

-9

44 7.21304x10"^ -6.99986x 10"-8 -2.64698x10"

-23 1.55005x10"

-9

Sum of the squares of the error

1

1.61966x 10"-12 2.33036x10" 16 2.02381 x10"

-16

Table 7 - Comparison of Errors in Mass Spectra Computed

From Compositions of Table 6

162
Gas Actual

Iteration

(All Ppak<;l

Iteration (Small Peaks

Deleted)

Difference

from Actual

H2 75.5 75.4954 75.4954 4.64835 X 10"^ CH4 6.67 6.70745 6.7 07 4 5 -3.74526 X 10-2

0.2 0.201330 0201330 -1.32975 X 10-2 CO 7.6 7.51186 5.72381 1.87619

7.5 9.59221 7.67615 -1.176153 02 0 0 0 0 2.2 2.17572 2.21509 -1.50884 X 10-2 CO2 0.2 0289488 0.189422 1.05776 X 10-2

Table 8 Effect of Deleting Small Peaks on Composition Calculations by the Iteration Method

N2 Actual = 0.79 Matrix CO Actual = 16 Matrix Case m e 14 me 30 m e 14 m e 30 lA 0.795487 -0.233733 15.9602 16.8174 IB 0.791744 1.76231 16.0119 152036 2A 0.732669 2.02149 16.0729 14.9983 2B 0.850865 -0.470798 15.9118 17.0137 3A 0.802927 1.15697 15.9586 15.6631 38 0.775903 0.45944 16.0703 16.3344 4A 0.654809 2.02425 16.1389 14.9981 48 0.925431 -0.475565 15.8665 17.0336 5A 0.738217 1.96145 16.0546 15.0344 58 0.847726 -0^ 80303 15.9547 16.9789

Table 9 Effect of Choice of Subset in the Matrix Inversion Technique

163
Noise (,/ Error

0 0.01 0.1 1.0

Calibration

Error %

0 3.38 xlO~^° 3.38 X 10"^ 3.383 X 10-2

0.01 2.468 X 10"^^ 3.40 X 10"^° 3.38 X 10"^ 3.383 X 10-2

0.1 2.578 X 10"^° 4.345 X 10-^° 3.405 X 10-^ 3.385 X 10-2

1.0 3.571 X 10"^ 2.608 X 10"^ 4347 X 10-^ 3.413 X 10-2

10.0 2.50 X 10"^ 2.50 X 10-2 2.541 X 10-2 4.366 X 10-2

Table 10 Standard Errors of Peak Height

Noise

Error %

0 0.01 0.1 1.0

Calibration

Error %

0 -4.019 X 10-3 -4.019 X 10-2 -0.4019

0.01 4.950 X 10-2 9.411 X 10-^ -3.523 X 10-2 -0.396904

0.1 4.957 X 10-2 4.555 X 10-2 9.404 X 10-3 -0.3521

1.0 0.4928 0.4888 0.4528 9.328 X 10-2

10.0 4.655 4.651 4.617 4.276

Table 11 Actual Errors in H2

Reading

Error %

0 0.01 0.1 1.0

Calibration

Error %

0 -3.215 X 10-3 -3.215 X 10-2 -0.3215

0.01 1.137 X 10-'^ -3.102 X 10-3 -3.204 X 10-2 -0.3214

0.1 1.137 X 10-3 -2.078 X 10-3 -3.101 X 10-2 -0.3204

1.0 1.37 X 10-2 8.154 X 10-3 -2.077 X 10-2 -0.3100

10.0 0.1134 0.1102 8.144 X 10-2 -0.2064

Table 12 Actual Errors in CH4

164
Paper No. 11

SPECTRAL ABSOLUTE REFLECTANCE MEASUREMENTS OF CO2 FROSTS IN THE

0.5 to 12.0 MICRON REGION-'^

B. E. Wood, A. M. Smith, and B. A. Seiber, Aerospace Division, von Kaman Gas Dynamics Facility, Arnold Engineering Development Center, Arnold Air Force Station, Tennessee and J. A. Roux, The University of Tennessee Space Institute, Tullahoma,

Tennessee. ABSTRACT: In situ absolute reflectance measurements have been

made for carbon dioxide frosts formed on liquid-nitrogen

cooled substrates. Data were obtained spectrally in the wavelength range from 0.5 to 12.0|i, using an infrared integrating

sphere. CO2 frosts were found to exhibit an anomalous dispersion reflectance peak at 4.3 microns which was shown to be

a very sensitive indication of the presence of solid CO2. Also

CO2 frosts scatter short wavelength radiation (X<l,0jj.)

significantly and is semi-transparent for much of the wavelength range between 2.0 and 12.0 . The application of these

results to problems associated with cryogenically cooled sur- faces is discussed.

KEYWORDS: carbon dioxide, reflectance, infrared, frost, cryo- deposit, simulation, spectral, cryogenics. Mars, integrating

sphere, scattering, absorbing, anomalous dispersion

INTRODUCTION - The thermal radiative properties of cryogenically cooled surfaces are being studied with increased interest

as more and more applications of cryogenic coolants appear. Examples of situations where knowledge of these properties is important are: (1) simulation of cold outer space in ground

test facilities by liquid nitrogen (LN2) cooled black surroundings, (2) reflective and emissive properties of cooled optics

(3) formation of frosts on cold surfaces to approximate the

"The research reported in this paper was sponsored by the

Arnold Engineering Development Center, Air Force Systems

Command, under Contract No. F40600-69-C-0001 with ARO, Inc.

165
formation of planetary frosts (i.e., polar caps on Mars),

(4) variation in transmission of cryogenically cooled optical

windows caused by thin film interference effects of thin

deposited films, (5) problems associated with absolute calibration of low temperature black bodies as reference sources,

and (6) reduction in visibility through spacecraft windows

caused by condensation of water and other contaminants.

Many of the previous cryogenic reflective property

measurements have suffered somewhat due to the deposits being

formed and measured at atmospheric pressure or because the

measurements were obtained by a calorimetric technique or within a narrow wavelength range. This study presents the

reflective properties of LN2 cooled surfaces in vacuum, with

and without the presence of CO2 cryodeposits . The reflectance

measurements were made in situ and covered the wavelength

range from 0.5 to 12.0^.

Absolute reflectance measurements of scattering type

(non-specular) surfaces are ordinarily made using an integrating

sphere. Integrating spheres having MgO or BaS04 coatings are limited to wavelengths less than about 2.6 microns due to the

low reflectance of the coating for longer wavelengths. However,

it was established in Ref . 1 that the usable wavelength range can be extended if powdered sodium chloride (NaCl) is substituted as the sphere coating. By using the NaCl coating,

and developing a new radiation source, the usable wavelength

range was extended out to 12.0|j,. Cryodeposit reflectance

measurements were obtained as functions of wavelength, deposit

thickness, t, substrate material and deposition rate, t. APPARATUS - The infrared integrating sphere (Fig. 1) was com- posed of two stainless steel hemispheres 8 inches in diameter.

The interior of the sphere was coated with powdered reagent

grade NaCl which was applied by first coating the sphere

interior with a thin layer of low vapor pressure vacuum grease

(lO"'''''" torr at 300°K) . This served as a bonding agent and then

the previously ground NaCl was pressed onto the sphere wall

until a thickness of approximately 5 mm had been obtained.

This technique of coating has the advantage of eliminating

highly absorbing water which was present when the water slurry

method described in Ref. 1 was used.

The cryosurface (Fig. 2) was a hollow, 1 by 1-1/2 by 1/2-

in. rectangular, stainless steel block which was cooled by

continuously flowing LN2 through it. The test surface was

one of the 1 by 1-1/2-in. faces and was either polished stainless steel or black epoxy paint. A copper-cons tantan thermocouple was silver soldered to the surface to allow monitoring

of the test surface temperature. To restrict the cryopumping

area to that of only the cryosurface, the LN2 supply lines

inside the chamber were vacuum jacketed. The pumping system

consisted of an ion pump and an LN2-trapped 4-inch oil diffusion

pump backed by a mechanical pump. This system allowed pressures

in the 10"^ torr range to be routinely achieved.

166
One of the problems associated with using an infrared

integrating sphere is that of obtaining an infrared source which emits enough energy at the longer wavelengths . In this

study, a 1000 watt tungsten-halogen lamp was enclosed in a housing which was cooled at both ends (see Fig. 3). This

allowed the middle of the housing to heat up to glowing red

thereby providing an extra source of infrared energy. With

this arrangement the source was usable over the wavelength

range from 0.3 to at least 25'^. Radiation from the source

was chopped at 13 Hz and focused onto the wall of the integrating sphere after passing through a sodium chloride window

(Fig. 1). After striking the sphere wall behind the test surface face, the radiation was reflected diffusely throughout

the sphere so that the portions of the sphere which were not illuminated directly were uniformly irradiated. Radiation

reflected from either the test surface or a portion of the

wall not directly irradiated was collected by an off axis

paraboloidal mirror and focused on the entrance slit of the

monochromator

The monochromator was a prism type single pass system.

For most of the measurements a NaCl prism was employed since

it was usable over the entire wavelength range from 0.5 to

12. Oil. In one set of experiments a CaF-^ prism was used since

it has a much higher dispersion that NaCl in the l-6\i range. The monochromator and paraboloidal mirror were mounted on a table which could be rotated about the vertical centerline

of the viewport, thus allowing radiation from either the test surface or sphere wall to be collected. The detector was a Charles M. ReedervS' thermocouple which had a sensitivity of

25 |Jv/|j,w. From the detector the output was amplified,

synchronously rectified, and displayed on a strip chart recorder.

Formation of the CO2 cryodeposit was accomplished by

flowing CO2 gas into the chamber at a known flow rate and

allowing the gas to be cryopumped by the LN2 cooled cryosurface.

The flow system consisted of two calibrated standard leaks with

a forepressure of one atmosphere. The deposition rate, t, of

the gas depended on the standard leak used,

PROCEDURE - The hemispherical -angular technique was used for

obtaining spectral absolute reflectance measurements. In

this technique the spectral hemispherical-angular reflectance

is determined from the relations (Refs. 1 and 2)

^i,h(x)/7r ^i^^^

where 1^(0, \) is the reflected intensity within the small

collection solid angle Aw (0.02 steradians) inclined at the

angle 6 and ) is the hemispherical ly incident energy

with the associated diffusely incident intensity

Since the same collection solid angle awt- is used for both

167
quantities, the absolute reflectance can be obtained from

Eq. 1. Even though the exact quantities 1^(0, X) and Ij[(X) are not measured, the detector outputs are directly related

to these quantities by multiplicative constants which depend

on the transfer optics and detection system. The measurements were accomplished by first viewing the sample and obtaining a detector output for a given wavelength. This was followed by

viewing a portion of the sphere wall (not directly irradiated)

at the same wavelength and again recording the detector output.

Since the sphere wall is uniformly irradiated by multiple

reflections, the intensity of the radiation reflected from the

sphere wall is equal to the intensity incident on the test surface. From these two detector outputs, the absolute reflectance can be determined from Eq. 1. Prior to cryodeposit reflectance measurements, the

integrating sphere system was evacuated to approximately

5 X 10"^ torr and then reflectance measurements were made on

the bare test surface.

After obtaining the reflectance as a function of wavelength, LN2 was used to cool the test surface to 94°K. Upon

valving off the integrating sphere from the pumping system

the CO2 gas flow into the sphere was started. By allowing

the gas to condense on only the test surface, a given flow

time resulted in a certain thickness, t, of the deposit. The

average thickness over the entire surface could be calculated

by knowing the mass flow rate, cryosurface pumping area, and

^ density of deposit. The cryosurface pumping area was 40.2 cm , and the mass flow rate either 2.25 x 10'-^ g/sec or 2.39 x 10""^ g/sec. The cryodeposit density was taken to be 1.67 (Ref. 3)

since the deposits were formed at pressures approximately the same as those in Ref. 3. In-situ thickness measurements were

made in order to verify the deposit thickness as calculated.

This was done using thin film interference techniques discussed in Refs. 3 through 5. The thicknesses determined by

the two methods were in good agreement.

The chamber pressure during deposition varied between

10""^ and 10"''" torr and depended on the flow rate and cryo- deposit thickness. For the majority of the measurements the

deposition pressure varied between 10"^ and lO"-'- torr. After

each thickness of deposit had been formed, reflectance measurements were made over the complete wavelength range from 0.5

to 12.0^. This procedure was continued until the deposit

had reached the desired thickness (usually 3-4 mm).

RESULTS - Black Epoxy Paint Substrate - The reflectances of

CO2 cryodeposits formed on an LN2-cooled black epoxy paint

substrate are shown in Fig. 4 for wavelengths between 0.5 and

12.0^. The view angle, 9, was 10 degrees for all thicknesses.

These deposits were formed using a deposition rate of 0.335;i,yfeec

which resulted in a deposition pressure of 1 x 10" to 5 x 10"^

torr, for deposits 0.44 ram thick or less. For the 3.38 mm

168
thick deposit the final pressure during deposition was

1.25 X lO"-*- torr. The flow time required to form a deposit

3.4 mm thick was approximately 2-3/4 hours.

Figure 4a shows the reflectance of the bare substrate

prior to deposition. The reflectance is essentially independent

of wavelength with only a slight increase from 5 percent up to 8-9 percent in the vicinity of 9-llp.. Figures 4b to 4d show

the reflectance of cryodeposits ranging from 0.13 mra up to 3.38 mm thick. As can be seen by comparison of the data for

different thicknesses, the reflectance begins to increase

first at the short wavelengths (\ <1.5'^) and continues to increase for the long wavelengths as the deposit thickness

is increased. The increased reflectance at the shorter

wavelengths has been previously reported (Ref. 5). For wavelengths greater than about 2.0^ the reflectance is seen

to be less than the bare substrate reflectance for the 0.13

and 0.23 mm thick deposits. This is also typical of the

reflectance for shorter wavelengths (\ <2.0|_l) but the effect

shows up for deposits thinner than any shown in Fig. 4. As the thickness is increased further, the reflectance increases

for all wavelengths. Figs. 4b and 4d. For the thickest deposit

investigated, t = 3.38 mm, the reflectance has risen to about

77 percent for X = 0.5'^i but drops off rapidly with increasing

wavelength.

In Figs. 4b and 4c the single most prominent spectral

feature is the sharp rise in reflectance in the region of

4.3iJ,. As the thickness of the deposit increases this feature

becomes less prominent due to the increase in bulk scattering

of other wavelengths around the 4.3^. region. The reflectance

peak is attributed to anomalous dispersion (Ref. 7) wherein a strongabsorption band in a liquid or solid has associated with

it a very high index of refraction so that both absorption and

reflection coefficients are quite high. It is well known that

in the gaseous phase, CO2 exhibits its strongest absorption at

4.25|j,. As can be seen in Figs. 4b through 4d the rapid increase

in reflectance in this region results in a sensitive technique

for identifying the presence of solid C02- The peak at 4.3|j.

has been seen for deposits as thin as 1 micron. For deposits less than 3.38 mm thick there is relatively

little spectral structure seen in the reflectance spectra in

Fig. 4 due to the transmission of the deposit and the highly

absorbing black substrate. However, for the 3.38 mm thick

deposit the scattering has increased significantly which

enhances the absorption bands due to the increased reflectance

for the nearby wavelengths. Absorption bands can now be seen

t o occur at 2.0^, 2.85'^ and what appear to be bands at 3.85

Additional experimental data are contained in Ref. 6.

169
and 4.55'4.. However, the latter two are really part of the

strong absorption band centered around 4.25'^ with the reflectance peak occurring in the middle of the absorption band.

Other spectral features to be seen in Fig. 4d are a broad

reflectance increase extending from 7.9 to 10.5|i and a gradual

rise in reflectance from 10.5i_l to 12. Oa. For all of the

deposits formed on the black substrate, it can be seen that

the reflectance will be less than 20 percent for wavelengths

between 2,7 and 12.0^. This could be a result of the deposits

being highly absorbing or the deposits being somewhat trans- parent with most of the energy being absorbed by the black

substrate. As will be shown later, the latter is the most significant factor.

The reflectance dependence on deposit thickness is pre- sented in Fig. 5. The view angle is again 10° and the dependence on thickness is shown for the selected wavelengths of

0.5, 1.0, 2.0, 4.0, 8.0, and 12.0^. For 0.5^ radiation, the

3.38 mm thick deposit effectively behaves as one of infinite

thickness since the reflectance has leveled off. However,

for \ = 1.0 and 2.0,^, the reflectance is still increasing

with thickness at 3.38 mm. The X = 2|j, curve is essentially

linear with thickness above t = 0.4 mm. At the other 3 wavelengths, the deposit reflectance is still dependent on thickness at 3.38 mm but much less so than for the 2.0|i, curve. The increase in reflectance for X = 8.0 and 12.0|j. is only

slight due to the relatively small scattering coefficient at these wavelengths. For the X = 4.0^. cuirve, very little

thickness dependence is present, as expected, since the CO2

deposit is highly absorbing at this wavelength.

To determine the reflective properties of thin deposits

on the black epoxy paint and to determine possible deposition

rate effects, a series of measurements was made using a deposition rate, t, of 0 . 0355iJ./sec which is about a factor of

10 less than the deposition rate used for the previously

discussed measurements. For these investigations, the wavelength range covered was from 0.5 to 5.0^ since it had already

been established that relatively little change was observed

for the longer wavelength radiation for the thinner deposits.

In Fig. 6, the reflectances of thin deposits formed at a deposition rate of 0.0355 'ii,/sec on the black substrate are shown for deposits 2, 27, 54, 108, and 215 microns thick.

Notice that even for the thinnest deposit, 2 microns thick,

that the 4.3ii anomalous dispersion band is already quite

apparent. The only other significant change in reflectance

for all thicknesses and wavelengths is the gradual increase

in reflectance for wavelengths between 0.5 and 1.5iJ, due to internal scattering. In the thickness region of overlap

between Figs. 4 and 6 it appears that the deposits formed at the lower deposition rate are less reflecting. For example,

at 231 microns, for the larger deposition rate, the reflectance

is about 32 percent at X = 0.5\i whereas for 215 microns formed

at the lower deposition rate, the reflectance is only 21 170
percent at the same wavelength. For the 2 micron thick deposit,

the reflectance is lower than the bare substrate reflectance

for wavelengths less than Ijj,. This effect is usually present

for all wavelengths although it is not so noticeable in Fig. 6. The deposition pressure for the lower deposition rate varied

from 10"^ torr for the 2 micron thick deposit up to 2 x 10"^ torr for the 215 micron thick deposit and was a little lower

than for the deposits formed at the higher deposition rate. Stainless Steel Sub strate - To understand more fully the

radiative properties of CO2 deposits, a polished LN^-cooled

stainless steel substrate was employed. This was the same

test surface used in the previous measurements but with the

black paint removed and the surface repolished. In Fig. 7a the reflectance of the bare polished stainless steel surface

is shown with an absorption band occurring around 3.0^. This

band could not be eliminated although various methods were

attempted such as mild bakeout under vacuum, repolishing

with various types of abrasives, and finally, cleaning with

freon in an ultrasonic cleaner. Since the band shows up at

3.Q|j, and increased scatter in the reflectance data is seen in

the 5 to 6^, region, ti would appear that some form of water

was permanently entrenched in the steel. As shown in Fig. 7b-d

reflectances of CO2 deposits of thicknesses, 0.13, 0.44, and

3.38 mil were formed on the stainless steel substrate. The

pressure during deposition varied from 2 to 4 x 10" torr

for the 0.13 to 0.44 mm thick deposits, and was 2 x 10' torr

for the 3.38 mm thick deposit. Notice that for the 0.13 mm

deposit the reflectance at the shorter wavelengths (X = O.Sjj,) has decreased from the bare substrate reflectance of 58 per- cent to 40 percent. Further increase in deposit thickness,

however, causes the reflectance to increase due to internal

scattering until eventually it levels off. For longer wavelengths, X >2.0^, the reflectance is influenced by strong

absorption bands in some regions, and a gradual decrease in

the reflectance in other wavelength regions as the thickness

is increased. It was expected in these non-highly absorbing

regions that the reflectance would gradually decrease with

thickness since only a slight increase in reflectance was seen for similar thicknesses of deposits formed on the black

surface. In contrast to the reflectance curves for the black

epoxy paint substrate, the absorption bands are quite pronounced

even for the smaller deposit thicknesses. In the wavelength

regions near these bands the deposits are somewhat transparent

with the result that the stainless steel substrate reflectance

continues to influence the amount of radiation reflected. The

high reflectance at wavelengths near the absorption bands accentuates the absorption in these regions.

By comparing the reflectance data for deposits formed on

the two substrates, it becomes apparent that the CO2 deposits

transmit appreciably in the wavelength region from 1 to 12ij,.

171
If the deposits were not somewhat transparent, the reflectances

observed for cryodeposits of equal thickness formed on the

two different substrates would be more nearly the same. For

instance, the reflectance at 5.5!i, for a 3.38 mm thick deposit

on stainless steel is 65 percent whereas for an equal thickness

on the black substrate the reflectance is only 19 percent.

This indicates that even at this relatively large thickness

the substrate reflectance still influences the reflectance

of the cryodeposit-substrate complex which can only be

explained by the deposit being transparent for radiation at these wavelengths.

As mentioned previously, the absorption bands of CO2 shows up more clearly in Fig. 7 than in Fig. 4. Absorption bands for

the 3.38 mm thick deposit can be seen to be centered around

2.0, 2.85-3.0, 4.3, 6.1, 6.85, and 7.55^1. The location of the

absorption band in the 2.85-3.0jj, region seemed to vary

with deposits. For instance in Fig. 7b the band is centered

at 2.85[i while thicker deposits exhibited the same absorption

band centered at 3.0,0,. This variation was also seen for other

deposits (not shown). As shown in Fig. 7c and 7d, the

reflectance peak at 4.3|j, is again present but is not so noticeable as in Fig. 4. Notice, however, that the peak

heights are about the same in each case (20-25 percent). For

wavelengths between 10 and 12jj,, there is a significant reduction

in reflectance. For the 3.38 mm thick deposit the reflectance

has dropped to 25 percent at 12. 0,^. The change in reflectance as the deposit thickness is formed on stainless steel is shown in Fig. 8 for a wavelength

of 0.705|i,, The reflectance was monitored continuously as the

deposit was being formed. Figure 8 shows that the experimental

reflectance curve has a sharp decrease initially, followed

by a more gradual decrease until reaching a minimum, and then

increasing until the reflectance levels off at a value much

hgiher than the bare surface. The explanation of this behavior

will be explained in detail in the next section where a theoretical study is made.

Analytical Results - Figure 8 shows that the reflectance

initially undergoes a decrease, then descends to a minimum and

finally increases with deposit thickness until reaching a plateau. Due to this interesting behavior an analytical

investigation was conducted to obtain a fundamental understanding of cryodeposit reflectances and their variation with

thickness, view angle, substrate, and refractive index. The

theoretical study was conducted with the objective of formulating a mathematical model for the cryodeposit reflectance based

on the results of the present experimental work.

In the search for a realistic reflectance model several

combinations of boundary conditions were considered for the

substrate and vacuum deposit interface. Based on the magnitude

of the reflectances and the general trend of the data presented

172
in this report and in Ref. 1, it was decided that the best

model would be to consider the substrate and vacuum deposit

interfaces as Fresnel surfaces. The internal structure of

the deposit was considered as an absorbing and isotropic

scattering medium; that is, the probability of energy being

scattered into every direction is equal.

This model for the cryodeposit reflectance required the

solution of the radiative transport equation:

u 2u J

dl

df

I(u)du (2)

•1 where

(K + fT)dT (3)

is the local monochromatic optical thickness at physical

thickness t. The transport equation is an integro-dif f erent ial

equation and was reduced to a system of ordinary linear

differential equations by the method of discrete ordinates

which employed the Gaussian quadrature;

> m

J) f(x)dx = XI f (Xi)aT

-1 j=l

(4)

where x. is the quadrature point and a^ is the quadrature

weight. Therefore,

in 11(1^1 = zLiL.^ + y~ I(T,u,)a, (5) dT u. 2u. / ^ 1 1 j=l

i = 1, 2, n

The solution of the system of simulataneous differential

equations together with the diffuse irradiance boundary con- ditions is extremely lengthy and complex and will be discussed

in detail in Ref. 8. It should suffice to mention that the

system of equations was solved numerically using the Milne

predictor-corrector method.

The theoretical investigation was conducted with the point

of view of matching the hemispherical-directional reflectance

measured for cryodeposits . This led to a computer program

that employs two experimental data points (the reflectance

at two different thicknesses) to calculate a monochromatic

absorption and scattering coefficient. Once the coefficients were calculated they were used to "predict" reflectance values

at other thicknesses.

173
Figure 8 shows a comparison between the theoretical and

experimental results for a cryodeposit on a stainless

steel substrate at \ = 0.7\i. The arrows indicate the two data points used in the computer program to calculate the mono- chromatic absorption and scattering coefficients. All other

points were predicted using these coefficients. The theoretical

results are in very good agreement with the data which indicates

that the analytical model proposed is realistic.

Since the cryodeposit reflectance model has proven to be

a reasonable one, it can be used to explain the shape of the curve in Fig. 8. For small deposit thicknesses near zero, there is a sharp (almost discontinuous) drop in reflectance

due to the relative refractive index change. This reflectance

decrease is from 0.49 to 0.388 and is denoted as the first

decrease. As the deposit thickness increases there is a second

reflectance decrease from 0.388 to 0.325. In view of the

theoretical model, this second decrease is a result of internal scattering occurring in the deposit. Some of the

internally scattered intensity is incident internally on the vacuum- deposit interface at angles greater than the critical

angle and undergoes total internal reflection. The reflectance

decreases since the thickness increase causes a higher probability that more intensity will be scattered past the critical

angle and eventually will be absorbed by the substrate before

being rescattered. As the thickness increases further, the

total internally reflected intensity has an even higher

probability of being rescattered before reaching the substrate

and being rescattered into directions less than the critical

angle and hence partially escaping the cryodeposit. At larger

thicknesses the internal scattering dominates causing more intensity to escape the deposit due to the rescattering effect

and resulting in a reflectance that now monotonically increases

with thickness. Now each ray is multiply scattered and is less

likely to reach the absorbing substrate. For greater deposit

thicknesses, the curve levels off at a reflectance plateau

of about 0.71. The reflectance should stay at approximately

this value as the thickness increases further since there is

a small but finite amount of absorption.

DISCUSSION - The results presented in Figs. 4 through 8 show

the reflectance of carbon dioxide cryodeposits formed on stainless steel and black epoxy paint substrates and ranging

from 2 microns to 3.38 mm in thickness. These thicknesses cover the range of concern for most of the situations listed

in the introduction. From the results discussed previously,

it is apparent that several aspects should be considered in

determining the reflectance of CO2 frosts. These are: wavelength, deposit thickness, and substrate material.

Possible applications of the data to some of the problems

mentioned earlier are apparent. The presence of thin

deposits (t <100 microns) on black surfaces will have a

174
relatively small effect; the reflectance in the visible and near infrared wavelength range will show only a slight increase

and for infrared radiation the reflectance will be reduced for

these same thicknesses. For even thinner deposits, t<20

microns, the reflectance in the visible and near IR can also

be reduced. For most applications in space simulation chambers

these changes will be negligible. The reflectance changes

for the larger thicknesses (t>100 microns), however, will

not be negligible.

The thin deposits may also have a significant effect in

the calibration of low temperature black body reference

sources. The presence of thin deposits will, in general,

increase the emittance, especially in the wavelength range

from 3.Q,i, to 12.0^. Although no data were taken for X>12.0^

it is believed that emittance values would also be increased

for longer wavelengths. With the current interest in absolute calibration of low temperature black bodies, these thin

film effects become increasingly important.

Cooled optics are often used in situations where the

surrounding pressure is relatively high, e.g., in balloons

or aircraft. Under such conditions it is impossible to keep

cryogenically cooled surfaces free from contamination due to condensation of CO2 and gases. For determination of

effects of the frosts on cooled optics (such as mirrors).

Fig, 7 shows some of the changes one could expect. In addition

to the reflectance considerations, the increase in emittance,

and the reduction in image quality due to scattering by the

cryodeposit may also have to be considered.

Finally, the recent findings of the Mariner 6 and 7 probes

have increased the interest in the reflective properties of

both carbon dioxide and water frosts. According to Herr and

Pimentel (Ref . 9) the reflectance data obtained from Mariners

6 and 7 indicate that the Martian polar caps are carbon dioxide.

They also reported that absorption bands at 3.34 and 3.04^,

which had earlier been designated as methane and ammonia bands,

respectively, were also due to CO2. They verified this by

obtaining the same absorption bands for CO2 frosts formed under

laboratory conditions. This was the first reported observation

of these two bands for CO2. These two bands were not seen

in the measurements presented in Fi^, 4 and 7 and the data in

Fig. 7 were, like the laboratory data for Herr and Pimentel,

for formed on stainless steel.

In order to obtain higher resolution data, the NaCl prism

was replaced by a CaF2 prism. Figure 9 shows the results of

the reflectance of a CO2 deposit 4 ram thick. The pressure

during deposition was 1.80 x lO'-'- torr or less and the view

angle, 0, was 18 degrees. With the higher resolution an absorption band was seen at approximately 3.36jj, which is probably the same band reported by Herr and Pimentel in Ref. 9. The other band reported by Herr and Pimentel at approximately

3.04jj. was not seen. The anomalous dispersion reflectance

175
peak at 4.3iJ. was not observed in the reflectance data of the

Martian polar cap shown by Herr and Pimentel in Ref. 9. This was probably due to the fact that the reflected radiation from

the polar cap was in turn absorbed by the surrounding CO2 gas. However, in a later article (Ref. 10) Herr and Pimentel report

seeing the anomalous dispersion peak at 4.3'^ but not while

viewing the polar caps but when viewing the bright limb of the

planet. This indicated to them that the solid CO2 was somehow

suspended above the planet in some type of particle cloud in a manner not completely understood. This observation establishes

conclusively the presence of solid CO2 either on the surface

or in the atmosphere of Mars. Herr and Pimentel suggest that

since the peak does show up, the cloud must be suspended

high enough up so that the atmosphere can be considered

optically thin as far as gaseous absorption in the 4.3;,l region

is concerned.

Further comparison of the data of Herr and Pimentel of

Ref. 9 with Fig. 9 indicates som.e significant differences.

First, of all, Herr and Pimentel show a strong broad absorption

band in the vicinity of 5.3 microns, which is even stronger

than the 4.3]j, band. This was not observed in any of the

measurements reported herein, nor was this band seen in Ref. 11 where CO^ frosts formed at atmospheric pressure were studied.

This might be due to the difference in the ordinate scale since

this report and Ref. 11 measured the reflectance whereas Herr

and Pimentel in Refs. 9 and 10 use intensity which probably

corresponds to unnormalized detector output. The present

study and Ref. 11 both observed narrow absorption bands near

5.0 and 5.24|X, whereas neither of these two bands were reported

by Herr and Pimentel in Ref. 9 either in the Martian polar cap data or in their laboratory data. Bands located near these

two were seen in the COr, gas in the Martian atmosphere and were found to be at 4.76 and 5.13 \i. As seen in Fig. 9 for the 4 ram thick CO2 deposit, there

is a broad absorption band centered around 3.0 microns which

is probably due to a small trace of water which was present

in the chamber from out gas sing. The band centered around

6.2 microns also indicates the presence of water impurities even though the chamber was initially pumped down to approximately 5 X 10"^ torr. To determine the effect of water

contamination in this wavelength range, a thin film of water

frost (t = 11 microns) was formed on top of the 4 mm thick

CO2 deposit (see Fig. 9). Even for this small thickness, the

reflectance was reduced for all wavelengths with the most noticeable reductions occurring in the 2.9 to 3.4}i region and

the 5.0 to 6.4jj. range. This means that the 3.04|j, band seen

in Ref. 9 would not have been seen due to the trace of water

which obscured any recognizable band in this wavelength region.

However, it is possible that the slight band at 3.04{j, of

Ref. 9 may also be attributable to water particles sparsely

scattered throughout the polar cap.

176
SUMMARY - In situ reflectance measurements on LN2-cooled

surfaces within a vacuum were carried out in a powdered sodium

chloride coated integrating sphere from 0.5 to 12.0^.

Reflectance measurements of cryodeposits formed on black

epoxy paint coated and stainless steel surfaces were obtained

for deposit thicknesses ranging from 2 microns to 4.0 mm, A

reflectance peak at 4.3}^ was found to be due to anomalous

dispersion. The presence of this peak in a reflection spectra

was shown to be a very sensitive method of detecting solid

CO2 since a deposit only 2 microns thick on black epoxy paint

shows the peak quite clearly.

CO2 deposits were seen to absorb strongly at 2.0, 2.85,

and in the vicinity of 4.3i_l. With a higher resolution

instrument, weaker absorption bands were also seen at 3.36,

5.0 and 5.24]_l. Thick CO2 deposits were highly reflecting at the shorter wavelengths (\ <2.0^) and transmit appreciably at longer wavelengths (excluding absorption bands).

Finally, the applicability of the experimental results to several current problems was discussed. These includeddetermination of the composition of the Martian polar caps, thermal radiative properties of cooled optics, effects on

space simulation chamber cold walls, and effects on emittance

of low temperature black body reference sources.

REFERENCES

(1) McCullough, B. A., Wood, B. E., Smith, A. M. , and

Birkebak, R. C. , "A Vacuum Integrating Sphere for In Situ

Reflectance Measurements at 77°K from 0.5 to 10.0^,"

Progress in Astronautics and Aeronautics , Vol. 20,

Academic Press, Inc., New York, 1967, p. 137.

(2) Sparrow, E. M. and Cess, R. D. Radiation Heat Transfer , Brooks/Cole Publishing Company, Belmont, California,

1966, p. 57.

(3) Muller, P. R. , Frost, W. and Smith, A. M. , "Measurements

of Refractive Index, Density, and Reflected Light Distributions for Carbon Dioxide and Water Cryodeposits,"

AEDC-TR-69-179 (AD692714) September 1969.

(4) Smith, A. M. , Tempelmeyer, K. E. , Muller, P. R. , and

Wood, B. E., "Angular Distribution of Visible and Near

IR Radiation Reflected from CO2 Cryodeposits," AIAA Journal, Vol. 7, No. 12, 1969, p. 2274.

(5) Wood, B. E. and Smith, A. M. , "Spectral Reflectance of

Water and Carbon Dioxide Cryodeposits from 0.36 to 1.15iJ,,"

AIM_i°'i^:^lil' ^^l- 6, no. 7, 1968, p. 1362.

(6) Wood, B. E., Smith, A.M., Seiber, B.A., and Roux, J. A.,

"Infrared Reflectance of CO2 Cryodeposits," AEDC-TR-70-

108.

177
(7) Jenkins, F. A. and White, H. E. Fundamentals of Optics , McGraw-Hill Book Co., Inc., 1957, p. 469.

(8) Roux, J. A., PhD Thesis, (In Preparation), University of

Tennessee Space Institute, Tullahoma, Tennessee.

(9) Herr, K. C. and Pimentel, G. C., "Infrared Absorptions

Near Three Microns Recorded Over the Polar Cap of Mars,"

Science, Vol. 166, No. 3904, 1969, p. 498.

(10) Herr, K. C. and Pimentel, G. C, "Evidence for Solid

Carbon Dioxide in the Upper Atmosphere of Mars,"

Science, Vol. 167, No. 3914, 1970, p. 47.

(11) Keegan, H. J. and Weidner, V. R. , "Infrared Spectral

Reflectance of Frost," J ournal of the Op tical Society

of_A[nerica, Vol. 56, No. 4, 1966, p. "523.

178
179
Substrate - Black Epoxy Paint

e - 10 deg T 0.335 microns/sec

Deposition Pressure 1 x 10"^ - 1. 25 x lO"'^ torr

!0r /-Bare Surface (t 0 mm)

°0 1 2 3 4 5 6 7 8 9 10 11 12

X, microns

a. T " 0 mm

0 1 2 3 4 5 6 7 8 9 10 11 12

X, microns

b. T - 0. 13 and 0. 23 mm

X, microns

c. T - 0.44 mm

°0 1 2 3 4 5 6 7 8 9 10 11 12

X, microns

d. T - 3.38 mm

4 Reflectance of CO^ Cryodeposits Formed

on a Black Epoxy Paint Substrate

180
Substrate - Black Epoxy Paint

e - 10 deg T 0.335 microns/sec

Deposition Pressure » 1 x 10"^ - 1.25 x lO''^ torr

Fig. 5 Reflectance as a Function of Deposit

Tinickness-Black Paint Substrate

Substrate - Black Epoxy Paint

e - 10 deg

- 108 microns t = 0. 355 microns/sec

- 215 microns

Deposition Pressure - 1 x 10'^ - 2 x 10'^ torr W. ./V

27 microns

54 microns

CLB

2 microns

4. \

Fig. 6

1 2 3 4 5 A, microns Reflectance of CO^ Deposits Formed on a Black

Substrate for ^ = .0355^L/sec.

181
9080

7060OL

Bare Stainless Steel

(t - 0 mm)

0 1 2 3 4 5 6 7 8 9 10 11 12

A, microns

a. T 0 mm

Substrate - Stainless Steel

"0 1 2 3 4 5 6 7 8 9 10 11 12 e - 10 deg

^' "^''^^^"^ T- 0.335 microns/sec

Deposition Pressure • 2 x 10'^ - 2 x 10' torr

0 1 2 3 4 5 6 7 8 9 10 11 12

T, microns

c. T 0. 44 mm

9080

<uo70

1_

O)a. 60

oT 50 o

c:

(U 40

tj

a> 30

"oj q; 20

10

0

0 1 2 3 4 5 6 7 8 9 10 11 12

X, microns

d. T 3.38 mm

Fig. 7 Reflectance of CO Cryodeposits Formed on a Stainless Steel Substrate

182
Fig. 8 Comparison of Experimental and Analytical Results

for Deposits Formed on a Stainless Steel Substrate

for \ = 0.7\i

Substrate - Stainless Steel e • 18 deg 4-mm CO2 Deposit - 4-mm CO2 Deposit with Coat of 11-micron-thick Water

1.4 2.0 2.5 3.0 3.5 4.0 4.5 5.0 5.5 6.0 6.4

X, microns ig. 9 Reflectance of a 4 mm Thick CO^ Deposit and a 4 mm

Thick CO^ Deposit with a 11 micron Thick Layer of HO

183

Pacer No. 12

NBS RADIOMETRIC STAM)ARDS — PRESMT MD FUTURE^

Henry J. Kostkowski

REFEPZETCE: Henry J. Kostkovski , "NBS Radiometric Standards —

Present and Future", ASTM/IES/AIAA Space Simulation Conference,

lU-16 September 1970.

ABSTRACT

One of the primary functions of the National Bureau of

Standards is to realize, maintain and disseminate national

standards of measurement. In radiometry, the currently available standards consist of various types of tungsten lamps or hlackhodies that have "been calibrated in terms of spectral radiance, spectral irradiance or total irradiance.

The standards of spectral radiance are either tungsten strip

lamps or blackbodies , have temperatures from 500 K to 3000 K,

and are calibrated, depending on the temperature and source, from 0.2 ym to 15 ym. Their uncertainty varies from about 0.1%

to 5^ depending on the temperature and wavelength (l,2)^. Two additional, useful standards that do not require calibration

and are available commercially are the graphite arc (3) and the

copper-point blackbody {h)

At present, there are two standards of spectral irradiance.

The first is a 1000-watt coiled coil tungsten filament quartzhalogen lamp (5) and the second consists of this lamp mounted

^Heat Division, National Bureau of Standards, Washington, D.C.,

2023U.

^This presentation concerns itself only with incoherent tyve

radiometric standards.

^The numbers in parentheses refer to the list of references

appended to this paper.

185
in a ceramic reflector (6). Their spectral irradiances at a wavelength of 0.9 ym and working distances of 50 and hO cm are about 25 and 105 y"w cm~^ nm"-*- respectively, and they both vary

with wavelength roughly as a 3200 K blackbody. These standards

are uncertain by values ranging from 3% at 2.5 ym to Q% at 0.25

ym. However, a research effort is now under way which is ex- pected to reduce these -uncertainties within a year by a factor

of about 3. The standards of total irradiance consist of the same lamps

used for spectral irradiance standards (6) and in addition 100,

500 and 1000 watt projection lamps (T). They have irradiances

varying from about 0.5 to 135 mw cm~^ at working distances from

100 to ^0 cm. The uncertainty of these sources has been estimated to be about 1%. Radiometric detectors that can be calibrated electrically

are becoming commercially available and appear useful in replacing irradiance source standards for some applications, particularly at the upper levels of approximately 100 mw cm~^ . Such

absolute or calorimetric detectors, as they are sometimes re- ferred to, are being developed and extensively evaluated at NBS. An uncertainty and reproducibility of a few tenths of a percent

is anticipated for some models of these detectors.

A description of the standards, of how they are realized,

of current attempts and limitations in extending their range or reducing their uncertainties, and the present NBS policy for

their dissemination will be included in the verbal presentation.

KEY WORDS: radiometry, radiometric standards, spectral radiance,

spectral irradiance, irradiance.

REFERENCES

1. Kostkowski, H.J., Erminy, D.E., and Hattenburg, A.T. , "HighAccuracy Spectral Radiance Calibration of Tungsten-Strip Lamps",

Advances in Geophysics lU , 111 (l9T0).

2. Yokley, C.R. and Kostkowski, H.J., "High-Accuracy Infrared

Calibrations of Blackbodies" , J. Opt. Soc. Am. 58, 1553A (1968).

3. Hattenburg, A.T., "Spectral Radiance of a Low Current Graphite Arc", Applied Optics 6, 95 (196?).

k. Lee, R.D., "Construction and Operation of a Simple High-Precision Copper-Point Blackbody and Furnace", Nat. Bur. Stand.

(U.S.), Tech. Note U83, 15 pages (May I969).

5. Stair, R., Schneider, W.E. , and Jackson, J.K. , "A New Standard of Spectral Irradiance, Applied Optics 2, 1151 (1963).

186
6. Schneider, W.E. , "A One-Solar-Constant Irradiance Standard"

Applied Optics 9, 1^10 (1970 ).

7. Stair, R. , Schneider, W.E. , and Fussell, W.B. , "The New

Tungsten-Filament Lamp Standards of Total Irradiance", Applied

Optics 6, 101 (1967)

187

Paper No. 13

APPAIUTUS FUR LABOIUTORY SI:IULATI()N OF lONUSPHliRIC FLOW OVER

UN-BOARD IIxiSTRUHENTvS OF AN AES.

E.I'i. Evlanov, V,B. Leonas, S.V. Umansky

Space Research Institute Academy of Sciences USSR, Moscov;;,

ABSTRACT: An apparatus for a simulation of ionospheric plasma

flow over on-board instruments of an artificial earth satellite

(AES) is described. Tlie simulated fluxes of nitrogen, hydroeen

and other ions have energies in the range of 5 -10 eV, current density of I'lO"^ A/cm and provide good simulation of the natural x7orking conditions of on-board instruments. A description

of diagnostic means v;hich are used for analysis of the flow

parameters is given.

KEY WORDS: laboratory space simulation, apparatus, ion source, vacuum svstem, beam diagnostic, ion transportation.

1. INTRODUCTION. The advent of explorations with the help of

AES leads to the necessitv of creating installations for simulation of the space environment parameters and, in particular,

for the construction of apparatus for laboratory simulation of ionospheric plasma.

In this report the apparatus designed for laboratory calibrations of on-board instruments for ionospheric studies is

described. For the solution of such a problem the parameters

of the laboratory plasma must be close to those which accrue in

the ionosphere [1] so one needs to have a weakly divergent monoenergetic (AE=^leV) ion flux Vvrith an energy of 5-100 eV of a certain mass composition and with a current density of

10 -10"-^ A/cm2, The ion beam diameter must be an order of magnitude larger than the entrance apertures of typical on-board

instruments, (D = 100 ram). Neutralization (charge compensation)

of the ion beam charge is desirable. The residual gas pressure

in the apparatus must be sufficiently low (10"'' -10"^ torr) and must correspond to the working conditions of on-board instruments in the upper atmosphere.

189
2. GENERAL DESCRIPTION [2] OF THE APPARATUS . Fig. 1 is a schematic diagram of the apparatus which provides ion flows

v/ith the noted parameters. Its appearance is shown in fig. 2. The apparatus may be schematically divided into three main

parts: I, Ion source 1 (fig. 1), x<rith a neutralizer, for production of the rarefied plasma. The ion source is mounted in

a single tube 1 (fig. 2), which is attached by means of a flange to the main vacuum vessel. II, Vacuum system with mano- metric tools. The main vacuum vessel is connected to the preliminary pumping equipment; inside of it a solenoid coil is

installed which is used for restricting divergence of the simulated flux. III. Test chamber III (fig.l) with movable equipment holders and instruments for the control of simulating

flux parameters. The test chamber 4 (fig. 2) is separated

from the main vessel by a tube 3 (fig. 2) in which diagnostic

instruments are installed. These instruments are removed

during operation from the path of the simulating flux.

3. ION SOURCE . For production of wide ion beams with a weak

divergence, sources with electron bombardment ionization were

used. The use of electron bombardment is conditioned by the

necessity to obtain ion fluxes of low energy (up to 5 eV) with

sufficiently narrow energy spread ("leV). Two modifications of the ion source were used. The first

ion source (fig. 3) consists of an ionizer (Pears system)

which forms the rectangular electron beam. The directly

heated flat tantalum cathode (100 x 1.5 mm) is covered with

yttrium oxide. This electron gun provides a current of 30 mA; the electron beam expansion along the ionization length

(= 100 mm) due to a space charge is - 10 mm, A positive

potential which determines the energy of ions in the beam is

applied at grids 3, 4 (fig, 3) while the accelerating grid 6 is under ground potential. Ions leaving the ionization space

with thermal velocities are accelerated and further move in

the equipotential space of the main vessel.

The ion source provides an ion beam with energy of 5 eV with current density of 1*10 ^ A/cm^ (Current was measured at

15 cm from the source). The source of the second type provides

currents 1 or 2 orders of magnitude higher due to an increase

of the ionizing electron current. The source consists (general

view is shown in fig, 4) of two hollow cylinders; one of which

is anode 1 and the other is an electron reflector 2 (fig. 4).

There are four 8 slots (2x4 cm) in the anode screened by

tungsten grids above which, at a distance of 2 mm, directly

heated tantalum cathodes are installed. The total emission

current is ~2A with electron energy of 100 V, The inside surface of the anode has the shape of a truncated cone (grid190
covered bases are of 40 and 50 mm diameter respectively)

which narrov; in the ion extraction direction. Such a configuration is an inverse ionizer of the V/eiss type [3] in which the

ion expulsion from the ionization region is actualized by the

ions soace charge field (but not electrons as by Iveiss). The source's functioning heavily depends on the pressure and optimal conditions are selected emnirically. The ions coming out from the ionizer are first accelerated and then decelerated by

an ion-optical system [4]. A system of this type effectively

enables one to extract from the ion beam fast background electrons and to reduce substantially widening of the ion beam

during deceleration.

4. VACUUII SYSTEI . A number of special requirements are imposed on the vacuum system of the apparatus.

Under working conditions the residual gas pressure must be

sufficiently low in order that charge-exchange and scattering

of the beam ions by the residual gas will be as small as possible. Keeping in mind the considerable gas load on the vacuum

system during an intense ion beam extraction, one needs a powerful pumping system for maintainance of lox^ pressure. The

presence of vapours of organic matter (oil) , which form films

on the sensitive elements of the instruments, is highly undesirable; i.e. the vacuum system of the apparatus must be

oilless

Titanium sublimation pumping is best for fulfilling the

above mentioned requirements [5] and this kind of pumping is used in the apparatus.

The main vacuum vessel of the apparatus II (fig. 1) has a near spherical shape (D - Im) . Inside of the vessel a near spherical liquid nitrogen cooled copper shield is installed.

In the lower part of the vessel an electron-beam evaporator 10

(fig, 1) is installed (O.lg/min) which sputters out the titanium on to all the inside surface of the shield. The primary

pumping of gas is done with a high-vacuum oil-diffusion pump, (pumping speed of 500 1/sec) provided with a reliable hydrogen

trap. The shield walls on the diffusion pump side are solid

and this sharply limits the oil vapour migration to the working

volume. The measured values of pumping speed for the pressure

of l'10-7 torr are S.IO^ 1/sec (K^) 2*10^ 1/sec (N2),

2»105 1/sec (O2) . The titanium sublimation pumping is not effective for noble gases and their pumping out was done by

the diffusion pump.

_g Ultimate vacuum (1.10 torr) in the apparatus without gas supply is reached in several hours. In the working condition

191
under the total gas load the pressure is 1*10 torr.

5. THE TEST CHAMBER AND DIAGNOSTIC TOOLS . The test chamber

III (fig. 1) is separated from the main vacuum vessel by the

tube 3 (fig. 2) with diagnostic equipment. The test chamber is equipped with a rotating device which provides alignment of the sensor position relative to the beam and facilitates studies of

the effects of the direction of ion incidence on the sensors.

The use of the separate chamber along with the practical con- venience enables one to eliminate sensor contamination with

titanium vapours. Beam diagnostics is carried out by means of the following

tools: Faraday cup 5 (fig. 1), Collimated channel multiplier 6 (fig. 1), and a sectional plate-condenser 7 (fig. 1). The

Faraday cup was used for absolute ion current density measurements. It was mounted on boom 2 (fig. 5), which permits determination of beam profile. An analysis of the ion beam divergence is performed by means of a well-collimated channel electron multioiier 3,

which is mounted on the same boom (fig. 5). Multiplier amplification coefficient is 10'' and the accuracy of divergency

measurement is ~1°. For the ion beam energy and mass composition determination

the time of flight method was used. For successful use of the

method a high-speed shutter for the ion beam break and a quick

sensor is needed. In the given case an open electron multiplier was used for ion registration. Ion beam pulsing was actuated by the sectional plate-condenser. It consists of

parallel connected plate-condensers (the distance between the

plates is 3 mm, width is 100 mm, number of plates is 30 for

first type source) and is mounted directly on the source 5 (fig. A). The composite design of the condenser with trans- parency of 90% permits its size to be decreased significantly

and exclude the effect of a distorting field on ion motion that

is important in diagnostics of low-energy and large diameter

ion beams. A typical TOF-oscillogram for a 20 eV ion beam is presented

in fig. 6. As can be seen the ion beam consists of Ar"*", N^

and H"J ions (the initial gas mixture) . An energy spread for

argon ion (Ar ) is -1,5 eV in the given case.

6. THE ION BEAM TRANSPORTATION . Due to the ion beam divergence for low ( -5 eV) energies which is caused by the effect

of the thermal velocity normal components, defocusing in the

192
accelerating grid cells and mainly by coulonbic repulsion ion current density rapidly decreases with distance. This effect

may be seen from the data presented in fin. 7, which shox^s beam current dependency on the accelerating voltage. Ion current density was measured both in the sources vicinity

(15 cm) and at a distance of 150 cm. In the first case current saturation is already observed for the accelerating voltage at

-5V; i.e. the ion current is not limited by the space charge

effect. At a distance of 150 cm the ion current continuously

increases with an increase of accelerating voltage and only for

300V does it become comparable with that at a distance of 15 cm. Ion beam divergence due to coulomb repulsion can be con- siderably decreased compensating the ion-beam space charge by

electrons. To accomplish this a neutralizing filament assembly

was mounted behind the source 6 (fig. 4) X7hich injected slow

electrons into the beam (electron energy -10 eV) . It is difficult to hold low energy electrons in the ion beam along its

whole track (-2 m) , hence, a solenoid is mounted inside the

main vessel 4 (fig. 1) producing a constant magnetic field

directed along the beam.

Fig, 8A shows the dependence of ion electron currents,

which are measured by the collector, on the magnetic field in

the solenoid. As can be seen, the beam transport optimum

conditions are achieved with a field of 20 G. In this case

the ion current increases by two orders of magnitude as com- pared to the case of a beam without neutralization (fig. 7B and fig, 8B) . Such a magnetic field has a small effect on ion

motion and the operation of the gauges studied.

7, CONCLUSIONS . An apparatus is described that provides

simulation of the conditions of the ionosDheric plasma flow

about satellite on-board gauges,

REFERENCES

1, Y,L. Alpert, A.V. Gurewich and L.P, Pitaevsky,

"Artifical satellites in rerefied plasma",

"Nauka", 1966,

2, E,N, Evlanov, V.B, Leonas and S.V, Umansky,

"Interplanetary medium and physics of magnetosphere"

Chodarev Y,K, ed, "Nauka" in press.

3, R. Weiss. "Rev. Sci. Inst." 32, 397,1961.

4, Von Koch. H. Lindholm E, "Ark Fys," 19.123.1961.

5, V.A. Simonov, G.F. Klamenov and V.A. Kochnev.

"Nucl. Synthes." Part 1, 325, 1962.

193
8

Fig, 1. Schematic of apparatus. I. Ion source chamber.

II. Main vacuum vessel. III. Test chamber, 1, Ion source.

2, diaphragm, 3, neutralizing assembly. 4. solenoid.

5. Faraday cup. 6. channel multiplier. 8,14. movable holder,

9. liquid nitrogen cooled shield, 10, titanium evaporator,

11, vacuum gauge, 12. residual gas analyser, 13. gauge

studied, 15 gas input.

Fig. 2. General view of apparatus. 1. ion source chamber.

2. main vacuum vessel. 3. tube with diagnostic tools.

4, test chamber.

194
Fig. 3. Schematic of ion-source

(fist type). 1, electron gun.

2. cathode. 3,4. ionization

space. 5. repelling grid.

6. accelerating grid.

7. electron collector.

8. reflector.

mi

Fig. 4. General view of sources of second type.

1. anode. 2. reflector.

3. cathode.

4. ion optical systems.

5. TOF-condenser.

6. neutralizer.

Fig. 6. TOF-oscillogram,

1. ion start pulse.

2. TOF-spectrum.

Fig. 5. Ion Beam diagnostic

tools. 1. Faraday cup.

2. Moveable boom.

3. colliraated channel

multiplier.

195
B

too

i'i;?, , 7. Ion ("J^ ) current ver, ion energy for tiie source of first tyr.e. At vvithout !)ean space charge compensation,

l- Source-collector distance 13 cn. 2, source-collector

distance 150 era, I), l.'ith snace cliarge conoensation , Sourcecollector distance 150 cn.

-8 IC

ID 20 SO 30 <o f^iG)

rig, 8. The ion (N,^ ) current and electron current ver, magnetic

field for the source of first tvpe (energy ions 10 eV)

A, 1. and I dependence on solenoid field.

B. it^ denentlence on solenoid field without beain soace charge

compensation.

196
Paper No. ik

THERMAL CONDUCTIVITY MEASUREMENTS OF A CANDIDATE VIKING HEATSHIELD MATERIAL AFTER STERILIZATION, AND DURING EXPOSURE TO

VACUUM, AND TO A SIMULATED MARTIAE ATMOSPHERE

Lawrence R. Greenwood and Roy M. Fleming

REFERENCE: Greenwood, Lawrence R. and Fleming, Roy M., "Thermal

Conductivity Measurements of a Candidate Viking Heat-Shield

Material After Sterilization, and During Exposure to Vacuum,

and to a Simulated Martian Atmosphere," ASTM/IES/AIAA Space

Simulation Conference , September l^-l6, 1970-

ABSTRACT: An experimental program has been conducted to measure the thermal conductivity of the proposed Viking heat-shield

material, Martin SLA-56I, after sterilization (60 hours at 276° F) and during exposure to vacuum and to a simulated

Martian atmosphere (7^*^ percent C02^ 12.8 percent N2^ 12.8 percent Ar) . In situ thermal conductivity measurements

were made at 75° F using the line-source technique. The thermal conductivity of SLA-56I was measured to be O.O298 Btu/fthr-'-'F at atmospheric pressure. In the first environmental

sequence consisting of atmosphere, vacuum, and simulated

Martian atmosphere exposure, a 6o-percent reduction in thermal

conductivity was measured in vacuum. After a 4-hour exposure

to a 7-'borr pressure in the simulated Martian atmosphere, the

thermal conductivity increased 67 percent from the value measured in vacuum. The second environmental sequence consisted of

measurements in the atmosphere, after sterilization and during

exposure to vacuum and the simulated Martian atmosphere. The

results of these measurements showed that sterilization had no

effect on the thermal conductivity measured at atmospheric

pressure nor on the changes measured in vacuum and in the

simulated Martian atmosphere. Thermal conductivity was measured at varying pressures during both environmental sequences

^Head, Space Vacuum Laboratory, Applied Materials and Physics

Division, NASA Langley Research Center, Hampton, Virginia.

^Student, N.C. State University (Employed as a Co-operative

Engineering Student at LRC)

197
and the results showed the thermal conductivity to be essen- tially independent of pressure below 10" torr with the transition from molecular to continuum flow occurring in the 10"^ to

10 torr pressure range.

KEY WORDS: thermal conductivity^ vacuum effect, ablation

material

Recent experimental results have shown that the thermal

and mechanical properties of spacecraft materials can be

altered by exposure to the space environment .-^^^^5 in particular, results have shown that the thermal and mechanical properties of ablative heat-shield materials may change due to sterilization and exposure to vacuum. To determine the thermal

conductivity of a candidate Viking heat-shield material during

the mission environmental sequence, an experimental program

has been conducted wherein in situ thermal conductivity meas- - urements were made after sterilization and during exposure to vacuum and a simulated Martian atmosphere.

EXPERIMENTAL PROGRAM

Thermal Conductivity Measurements

The line-source technique was used for measuring thermal

conductivity."'" In this technique a heater wire and thennocouple wire are placed inside a sample as shown in Fig. 1. Heat is generated in the sample by passing a known heater

current through the heater wire, and the variation of temperature with time is measured with the thermocouple located on the heater wire. The heater circuit consisted of a 6-volt

battery, an ammeter, and a variable resistor for presetting the

heater current prior to initiating sample heating. After

initiating the heater current, the temperature -time history of

the heater wire is recorded using the thermocouple located on the heater wire, a reference junction and a recorder. From the

recorder output, a computer program is used to fit a least

squares straight line to the data and calculate the thermal

conductivity from the slope of the line. The thermal conductivity was calculated from^

where

K = thermal conductivity

q = heat input

198
6 = temperature

t = time after initiation of heat generation

Fig. 2 is a photograph of the thermal conductivity

instmimentation and Fig. 3 is a photograph showing details

of the sample instrumentation. All measurements were made at 200 mA heater current with a heating period of 5 minutes. A

second thermocouple was inserted ±/h inch (see Fig. 3) "below the sample surface and monitored during the tests to determine

if the surface temperature was changing. Also, a plot of the

logarithm time -temperature output was also examined to check

the linearity. All meas^urements reported are the average of

the thermal conductivity measured on three samples at a given

time.

Environmental Chamber

The sterilization_, and exposure to vacuum and a simulated

Martian atmosphere was done in the 20 -cu- ft -ultrahigh vacuum

chamber shown in Fig. 2. This is a stainless steel chamber

capable of achieving pressures of 10"-^^ torr. A complete

description of the chamber is given in Ref . 5v

Material

The candidate Viking heat-shield material studied was

SLA-56l_, made by the Martin Marietta Corporation. The SLA-56I

has a density of 0.201 g/cc and is composed of a silicone

elastomer filled with silica spheres phenolic microballoons

silica fibers, and cork. A complete description of the

material, including its ablative properties, can be found in

Ref. 6. Thermal conductivity samples were ^ by ^ by 6 inches,

and typical samples are shown in Fig. h. The 0.010-inch

manganin heater wires and the 30-gage copper-constantan thermocouple wires were cast into the heat-shield samples during

sample preparation. The material was preconditioned several

weeks at percent relative humidity in a class 100 clean room prior to testing.

Environmental Sequence

Thermal conductivity measurements were made during two

environmental sequences: (l) atmosphere, vacuum, and simulated Martian atmosphere exposure, and (2) atmosphere, sterilization, vacuum, and simulated Martian atmosphere exposure.

The first sequence began by making several thermal conductivity measurements on the three samples in the vacuum chamber

at atmospheric pressure (see Fig. 3)- The chamber was then

evacuated and thermal conductivity measurements were made as a function of time during the 128.5-hour vacuum exposure time.

After 1^6.1 hours, the pressure was increased to 7*0 torr with

a simulated Martian atmosphere consisting of 7^-^ percent CO^,

12.8 percent N2, and 12.8 percent Ar. Thermal conductivity

199
measurements were made at 7*0 torr. The chamber pressure was then reduced to 1.0, and 0.1 torr and thermal conductivity

measurements were made at each of these pressures. The

pressure time history for sequence 1 is shown in Fig. 5« All

environmental exposures and thermal conductivity measurements were at 75° F

At the start of the second sequence^, three different

thermal conductivity samples were sterilized in the chamber by

heating the samples of 276 F for 60 hours at a pressure of

600 torr in a dry nitrogen atmosphere. After sterilization the

thermal conductivity was measured at 75° F and the vacuum

exposure begun. Thermal conductivity measurements were made as

a function of vacuum exposure time for 95*0 hours and after

96.5 hours the pressure was raised to 7«0 torr with the

simulated Martian atmosphere. The pressure time history for

this sequence is also shown in Fig. 5.

EXPERIMENTAL RESULTS

Sequence 1 - Atmosphere, Vacuum, Simulated, Martian Atmosphere

The thermal conductivity measured at atmospheric pressure

was 0.0298 Btu/ft-hr-°F as shown on Fig. 6. After 1.8 hours in

vacuum, the average thermal conductivity had been reduced

k6 percent to O.OI6I Btu/ft-hr-°F . For the first 20 hours in vacuum the thermal conductivity continued to decrease, and after

31.5 hours the conductivity was 0.0120 Btu/ft-hr-°F and

remained approximately constant for the remainder of the vacuum

exposure period. After l46.1 hours the pressure in the chamber

was raised to 7 torr with the simulated Martian atmosphere.

Measurements were made immediately after reaching 7 torr and

after k-, 8, and 12 hours. The thermal conductivity measured

at these times was O.OI86, 0.0200, O.OI9I, and O.OI95,

respectively. The pressure was then reduced to 1.0 torr for

4 hours, after which time the thermal conductivity was 0.01^5 Btu/ft-hr- F. The pressure was then reduced to O.5 torr

for k- hours after which time the thermal conductivity was 0.01^0 Btu/ft-hr- F. Finally, the pressure was reduced to 0.1 torr for h hours after which time the thermal conductivity

was 0.0131 Btu/ft-hr-°F.

Sequence 2 - Sterilization, Vacuum, Simulated Martian Atmosphere

A second environmental sequence on three different SLA-56I

samples was conducted wherein thermal sterilization was

included. The thermal conductivity meas-ured at 6OO torr in a nitrogen atmosphere after sterilization was O.O296 Btu/ft-hr- F

(see Fig. 7). Since the thermal conductivity for the unsterilized samples discussed in sequence 1 was O.O298 at atmospheric

pressure, there was essentially no effect of sterilization on the thermal conductivity. After sterilization the vacuum

200
exposure began, and the measured thermal conductivity values are shown in Fig. 7* The vacuum- induced changes for the sterilized

samples (sequence 2) were similar to the vacuum- induced changes

for the unsterilized samples (sequence l) . For instance, after

2 hours in vacuum, the thermal conductivity of the sterilized

samples (sequence 2) was O.OI6I, whereas the thermal conductivity

of the unsterilized samples (sequence l) was 0.0155 Btu/ft-hr-°F. After 95 hours in vacuum, the thermal conductivity of the sterilized samples (sequence 2) was 0.01^0 Btu/ft-hr-°F compared with

0.0124 Btu/ft-hr-°F for the unsterilized samples (sequence l).

After 96.5 hours the pressure was increased to ^ .0 torr

with a simulated Martian atmosphere. Three hours l^ter the

thermal conductivity for the sterilized samples (sequence 2) was 0.0212 Btu/ft-hr-°F. The thermal conductivity for the

unsterilized samples (sequence l) was 0,0200 Etu/ft-hr-^ after

k hours in a 7-torr atmosphere of the simulated Martian

environment. The sterilization had very little effect on the

thermal conductivity of SLA-56I during vacuum exposure and

exposure to the simulated Martian atmosphere at 7 torr. Effects of Pressure

The effect of pressure for both the unsterilized (sequence

1) and sterilized (sequence 2) samples is shown in Fig. 8. First, it should be noted that there are no discernible

differences in the thermal conductivity measured at the same pressure for either of the sequences studied. The measurements

do indicate that at pressures below 10~^ torr the thermal

conductivity may be independent of pressure. In the pressure

region 10'^ to 10 torr, the characteristic knee in the thermal

conductivity curve (due to transition from molecular to continuum flow) was observed. At about 10" torr, there appear

to be differences in the thermal conductivity measured during

the vacuum chamber pumpdown and those measured in the simulated

Martian atmosphere. This is difficult to confirm because the vacuum pumpdown measurements were made while the pressure was changing as a function of time (see Fig. 5) and. thus it cannot

be assumed that the samples were allowed sufficient time to establish equilibrium with their environment.

CONCLUSIONS

The results of making in situ thermal conductivity meas- urements on SLA-561 during sequence 1 (atmosphere, vacuum, simulated Martian atmosphere) and sequence 2 (atmosphere,

sterilization, vacuum, simulated Martian atmosphere) resulted

in the following conclusions:

1. The thermal conductivity of the unsterilized samples

(sequence l) was reduced by k6 percent after 1.8 hours in vacuum and by 60 percent after 51.5 hours. The thermal

201
conductivity after 12 hours in a Y-torr gressure of a simulated

Martian atmosphere was 0,0193 Btu/ft-hr- F. The thennal

conductivity did not appear to he a function of time in vacuum

after 31*5 hours of exposure..

2. There was essentially no effect of sterilization on either the thermal conductivity measured at atmospheric

pressure or on the vacuum -induced changes in thermal conductivity. Measurements at 7 torr for both sterilized and

unsterilized samples were approximately the same.

3, The thermal conductivity measured for both environmental sequences was essentially independent of pressure below

10"^ torr. In the 10"^ to 10 pressure region the thermal

conductivity was a function of pressure.

B.IBLIOGMPHY

1. G-reenwood, Lawrence R,; and Comparin, Robert A.: Application of the Line-Source Technique for Vacuum

Thermal Conductivity Measurements . Presented at the

AIAA/ASTM/IES Fourth Space Simulation Conference^ Los Angeles^ California, September 8-10 , I969.

2. Pyron, C. M.; and Pears, C. D.: Effects of Sterilization

and Vacuum Exposure on Potential Heat Shield Materials

for Unmanned Mars Mission. NASA CR-66557, I968.

3. McCrary, Leon; and McCormick, Carl: Effects of Sterilization and Vacuum Environment on Ablation Materials.

Presented at the lES ikth Annual Technical Meeting,

St. Louis, Missouri, April 29-May 1, I968.

k-. Carslaw, H. S.; and Jaeger, J. C: Conduction of Heat in

Solids. Oxford, Clarendon Press, 19^7.

5. Bradford, James M. : Operating Parameters of the Omegatron

Mass Spectrometer in Ultrahigh Vacuum. NASA TN D-3129.

6. Strauss, Eric L. : Superlight Ablative Systems for Mars

Lander Thermal Protection. Journal of Spacecraft and

Rockets, Vol. No. 10, pp. 130^-1309.

202
HEATER WIRE —vWA'

BATTERYAMMETER

THERMOCOUPLE WIRE

TEST SAMPLE

THERMOCOUPLE

JUNCTION

V

/ REFERENCE JUNCTION

RECORDER

Figure 1.- Experimental technique used for measuring thermal

conductivity.

203
Figure 2.- Experimental apparatus used in maimig the thermal

conductivity measurements.
;liieli sazriples used in the theiTrial )nduGt;i\n.t7 study.

205
10'

10

10^

10

10

S 10-3

^

10"

10

10"° -

10"

O Sequence 1 (atmosphere, vacuum, Martian atmosphere)

Sequence 2 (atmosphere, sterilization, vacuum, Martian atmosphere)

Shaded symbols denote simulated Martian atmosphere.

° Oo oo

oo ?)0

20 40 60 80 100 120 140 160 180 200 Time, hours Figure 5«- Pressure time histoiy during thermal conductivity

study

206
Shaded symbols denote simulated Martian atmosphere.

o o oooo o o

€) A.O torr yy-0.5 torr //^/O.l torr

0 20 40 60 80 100 120 140 160 180 200 Time, hours Figure 6.- Thermal conductivity of SLA-561 measured for

sequence 1 (atmosphere, vacuum, Martian atmosphere).

000 o o 0000

O Sequence 1 (atmosphere, vacuum, Martian atmosphere) Sequence 2 (atmosphere, sterilization, vacuum, Martian atmosphere) Shaded symbols denote simulated Martian atmosphere.

Figure 7.- Thermal conductivity of SLA.-56I measured for

sequence 1 and sequence 2. 207
/

O Sequence 1 (atmosfdiere, vacuum, Martian atmosphere)

Sequence 2 (atmosjiiere, sterilization, vacuum, Martian atmosjiiere) Shaded symbols denote simulated Martian atmosphere.

oo

-5 -4 -3 -2 -1 10 10 10 10 10 1 Pressure, torr Figure 8.- Effect of pressure on the thermal conductivity of

SIA-561.

208
Farer Nc . 1

METHOD FOR ROCK PROPERTY DETERMINATION IN ULTRAHIGH VACUUM

by Egons R. Podnieks and Peter G. Chamberlain

REFERENCE: Podnieks, Egons R. and Chamberlain, Peter G.

"Method for Rock Property Determination in Ultrahigh Vacuum",

ASTM/IES/AIAA 5th Proceedings of Space Simulation Conference,

Gaithersberg, Maryland, Sept. 14-16, 1970.

ABSTRACT: A special ultrahigh vacuum system was designed for

determining deformat iona 1 and strength properties of simulated

lunar rocks. The system provides a vacuum of 10~- to 10"-- torr (10"'^ to 10~- N/m^) depending on the specimen size and

rock type. The vacuum was measured by various types of gages

and also by ion pump current measurements. The partial pressure measurements of the various gas components were obtained

by a quadrupole residual gas analyzer. The vacuum chamber

has two stainless steel bellow-type feedthroughs for loading

and a spring mechanism to compensate for the atmospheric pressure effect on the rock specimen. The uniaxial load was applied to the specimen by a servo-controlled hydraulic

testing machine. Specimens of tholeiitic basalt, dacite, and

semiwelded tuff were used. Special procedures in specimen

preparation and preconditioning were developed. During the

pumpdown period, prolonged roughing and an initial bakeout

produced significantly low final pressures in the chamber. Test data presented include a typical load-deformation curve from rock specimen in ultrahigh vacuum, ' outgassing characteristics during loading in terms of pressure variation, and

changes in the composition of gases being released by the

specimen.

KEY WORDS: rock mechanics, space environment simulation,

ultrahigh vacuum, pressure measurements, outgassing, mass

spectroscopy, mechanical properties, anisotropy, compression

tests

''Supervisory Mechanical Engineer and Geophysicist , respectively, Twin Cities Mining Research Center, Bureau of Mines,

U. S. Department of the Interior, Minneapolis, Minn.

209
Since 1965 the Bureau of Mines has been engaged in NASAsponsored multidisciplinary research leading to the use of

extraterrestrial resources . As a part of this broad program, rock failure processes and the strength and elastic properties of rock in simulated lunar pressure environment have been

investigated. To facilitate these studies a special ultrahigh vacuum system has been designed that can be readily moved

in and out of a hydraulic testing machine, to permit axial

loading of a rock specimen through a set of bellow f eedthroughs

This vacuum system is equipped with a variety of pumps incorporating both momentum transfer and capture types. With this

system an ultrahigh vacuum of 10~® to 10~^° torr (10"'^ to 10~® N/m^) can be obtained in 2 days depending on the rock type and

specimen size. During the axial loading of a specimen, the

rock deformation behavior is determined and compared with

simultaneously obtained outgassing data. Total pressure rise

in the chamber and the partial pressure measurements of the

outgassing components, made by a quadrupole residual gas analyzer, are correlated with the deformationa 1 and fracture

characteristics of the test specimen.

EXPERIMENTAL EQUIPMENT ' \ The vacuum chamber consists of a 14-in. (35.5 cm) stainless steel cylinder 26-in. (66 cm) long equipped with ports

for attaching pumps, gages and other accesories (Fig. 1).

The purpose of this chamber is to provide an ultrahigh vac- uum environment for rock testing in simulated lunar pressure

environment. Since the first phase of this research covers only the effect of vacuum on rock strength, a relatively

compact test chamber is quite satisfactory.

The pumping system (Fig. 2) consists of a 400 1/sec

(0.4m^/sec) ion pump supplemented by titanium sublimation and

cryopumping. The ion pump is directly attached to the chamber

by a 6-in. (15.8 cm) port. The sublimator deposits titanium

film directly on the wall of the test chamber and onto a Chilled copper coil with the test specimen and the gaging

protected by proper baffling. The cryopumping is accomplished

by flowing liquid nitrogen (LNg) through a coil of copper

tubing lining the chamber near the test specimen. The initial

pumpdown of the system is performed by three cryosorption

pumps supplemented by a mechanical carbon vane pump. The vacuum conditions in the chamber are first monitored

by thermocouple and alphatron gages mounted on the roughing

manifold. After the roughing cycle the vacuum measurements

are obtained directly from inside the chamber by a nude ionization gage and/or a cold cathode (Kreisman) gage and also

by ion pump current measurements. A quadrupole residual gas

210
analyzer is used intermittently first for leak detection and

second for giving partial pressure measurements of the various

gas components originally present in the system or being re- leased by the rock specimen during the test. The vacuum chamber is specially adapted for compression

testing of rock specimens in vacuum (Figs. 3 and 4) . Both of

the axial loading platens are connencted with the chamber by

stainless steel bellows to allow platen movement during the

test and also to provide a positive bakeable vacuum seal. In addition the platens are spring mounted to compensate for the

force applied to them by the pressure differential during the

pumpdown of the system. The rock sample is surrounded by a metal screen to prevent rock particles that spall away during

the compression test from damaging other components in the

chamber. The load applied to the specimen is measured by a load cell located underneath the specimen and by a second load

cell exterior to the vacuum chamber. The specimen deformation

is measured by a strain gage - cantilever beam type extenso- meter which indicates the deflection between the platens.

The vacuum chamber assembly is mounted on wheels and is moved into position in the compression testing machine when a test is to be made (Fig. 5) . A servo-controlled hydraulic

testing machine is used to apply the axial load to the speci- men at a controlled load or deformation rate (Fig. 6) . During

the compression tests an X-Y recorder handles load-deformation

data, a strip chart recorder monitors pressure changes in the vacuum chamber, and an oscilloscope camera documents mass spectrometer traces at various stages of specimen deformation.

EXPERIMENTAL TECHNIQUES

Rock Selection

The rock materials chosen for this program were earthformed rocks assumed to simulate lunar rocks. In establishing

the rock standards for lunar research the following criteria

were used for rock selection (J^)^: 1) they should represent

lunar rocks on the basis of lithology, texture, and composition, 2) they should meet specific requirements for the particular research likely to be performed on the rocks, 3) they

should be uniform in composition and texture and 4) they

should be convenient to collect in sufficient quantity.

Following these criteria the Bureau of Mines selected fourteen

simulated lunar rocks covering a broad range of possible lunar

rocks (2) . From this suite of fourteen rocks three were first

tested in this program: 1) a tholeiitic basalt (from N. E. of

2 "The numbers in parentheses refer to the list of references

appended to this paper."

211
Madras, Oregon) to simulate rock from large basaltic flows;

2) a dacite (from west of Bend, Oregon) to simulate an extrusive rock with a composition intermediate between rhyolite and

basalt; and 3) a semiwelded tuff (from Bend, Oregon) to simulate possible ash flows on the Moon. These three rocks represent a wide range of properties: bulk density - 1.15 to 2.84

g/cc (1150 to 2840 kg/raP) , Young's modulus - 0.3 to 10.3 x 10® psi (2.1 to 71 GN/m^) and compressive strength - 850 to 53,000

psi (5.9 to 370 MN/m^)

Rock Specimen Preparation and Preconditioning

Preparing rock specimens for compression testing requires

special attention because of the time consuming preparation

and preconditioning processes involved. Since most rocks are anisotropic and nonhomogeneous , i. e., their oroperties vary with

direction and location within a rock mass (_3) , their properties

have to be measured and analyzed with respect to fixed reference directions. At the Twin Cities Mining Research Center a standard procedure has been developed for obtaining oriented

rock specimens. If possible, rock blocks are oriented with

respect to in situ principal geologic and geographic directions. From each block a 3- or 4-in. (7.6- or 10.1 cm) spherical specimen is prepared maintaining reference directions

with respect to the source block throughout the process (4)

Longitudinal pulse velocity measurements are performed on these

spheres in many diametrical directions. These measurements are plotted on equal area nets and contoured, resulting in velocity

S3rmmetry patterns that provide information on the orientation

of rock fabric features , e.g., elongated vesicles, which control

other property anisotropy as well (3) . From the velocity

syinmetry patterns meaningful directions in which to measure the

compression properties are selected, and since the sphere

orientation is carried throughout the process, it is possible

to prepare oriented specimens from the source block. Cores were drilled from the source block in the selected directions

with a diamond impregnated bit, and cut to size with a diamond-blade saw. For these tests cylindrical specimens were used with a length to diameter ratio of 2:1. The tholeiitic

basalt and dacite specimens were 1-in. dia. (2.54 cm), whereas the semiwelded tuff specimens were 2-in. (5.08 cm) in

diameter because of large inclusions in the rock. The dimensional tolerances being formalized into ASTM Standards were used in the preparation of rock specimens (_5) . The ends were

cut and surface ground perpendicular to the axis of the speci- men within 0.25°, the diametrical variations were held within

0.005 in. (.1270 mm), and the end roughness within 0.001 in. (.0254 mm). During the cutting, drilling and grinding operations the only coolant was water. Before being placed in the vacuum chamber, the rock specimens were dried in a vacuum oven

for one week at 135 °C and about 50 x 10"^ torr (6.66 N/m^)

212
pressure. After being cooled to room temperature, the speci- mens were backfilled with dry air so that they could be trans- ferred to desiccator for storage under moderate vacuum. Test Procedure

The preconditioned rock sample was placed in the vacuum

chamber while the chamber was in the testing machine to assure

proper alignment, preloading and checkout of interior instrumentation. While the chamber is open to the atmosphere, dry

nitrogen gas was forced through it to minimize moisture con- tamination. After the specimen was properly mounted on the

test platens, the system was closed and the rough pumpdown

started by a carbon vane vacuum pump. At that time the pressure differential across the platens forced the bellows to extend until both platens contacted the specimen and created

a preload on the specimen. This preload was held to the

desired value by the compensating springs. With the specimen

secured between the platens, the vacuum chamber was moved out of the testing machine for the pumpdown cycle.

The initial rough pumpdown of about 15 minutes by the

carbon vane pump reduced the chamber pressure to approx. 25 torr (3340 N/m^) . Thermostatically controlled electrical band

heaters encircling the exterior of the chamber were then

turned on and final roughing by cryo-sorption pump was initiated and continued for 4-hr. During this pumpdown phase the

chamber pressure was monitored with an alphatron gage located

in the manifold section of the vacuum system. After the 4-hr.

roughing, when the pressure in the chamber is in the middle of

10""^ torr (high 10"^ N/m^) region, the manifold was closed and

the ion and the sublimation pumps were started. During the

following pumpdown ion and cold cathode (Kreisman) gages were

used to monitor the chamber pressure. A 24-hr. bakeout was initiated when the ion pump was turned on. This bakeout was thermostatically controlled at 135°C and was automatically

interrupted if pressure in the chamber rose above the 10"^ torr

level. During the bakeout the system was checked for leaks

with a mass spectrometer using helim gas. After bakeout the

system was cooled for 24-hr. period before the rock specimen

was tested. Figure 7 shows typical pumpdown curves for the

empty chamber and for chamber containing specimen of basalt,

tuff and dacite.

After the desired vacuum has been reached, usually in the

low 10-s or 10-^° torr (low 10"'^ or 10"^ N/m^) range, the vacuum chamber was positioned in the testing machine and the

pretesting alignments were made. Shortly before the start of

the compression test, the cryoshroud was cooled off with LN^ resulting in a 0.5 - 1.0 decade pressure drop within the test chamber. The rock specimen was compression-tested to complete

failure at a constant deformation rate of 1 x 10~^ cm/cm/sec.

213
During the tests the load applied to the specimen and the deformation of the specimen were recorded on an X-Y plotter

(Fig. 8) . The chamber pressure variation (Fig. 9) as well as partial pressure changes indicated by the mass spectrometer

traces (Fig. 10) were recorded simultaneously with the loaddeformation data.

RESULTS AND CONCLUSIONS

Applying vacuum technology to rocks requires a modification of the conventional methods of simulating a space environment. The problems of space technology usually deal with

man-made materials that are specifically designed to meet the

extraterrestrial environment requirements. In this program,

however, natural rock material, being studied under simulated

lunar vacuum conditions, cannot be chosen by its adaptability

to vacuum technology. This factor requires special attention

to establish valid preconditioning methods in order to ap- proach conditions in the rock specimen (that have been se- lected because of the likelihood of their occurrence on the

Moon) that are similar to those in actual lunar rocks.

Previous studies at the Twin Cities Mining Research

Center of vacuum and moisture effects on rocks (_5 , 6_, 7^) have

helped to establish specimen preconditioning requirements for

various rocks and specimen sizes. These studies have shown

that is is nearly impossible to obtain complete equilibrium

between the vacuum environment surrounding the rock specimen

and the entire volume of the specimen. Equally evident from

the experiment is the effect of preconditioning treatment on

the pumpdown and bakeout procedures. Since terrestrial rocks

normally contain gases and moisture either in pores or interstitially, the preconditioning treatment and the initial rough

pumping have to be prolonged enough to remove the most of the gas load before the chamber is sealed for high vacuum pumping

with ion and sublimation pumps. Because of the rock structure,

the bakeout process requires caution. To avoid causing irreversible structural changes in the rock, bakeout temperatures

should be kept to the minimum required to discharge water

vapor.

As the rock undergoes deformation during the compression

test, cracks are formed and existing cracks either opened or closed. This formation of new surfaces and opening of interi- or cavities causes a very distinct increase in pressure (Fig.

9) which continues until the specimen fails (compressive

strength) . During the post-failure loading when mostly redistribution of broken rock particles takes place, few new frac- ture surfaces are formed and the pressure tends to drop

because the outgassing rate falls below the pumping capacity

214
of the system. The outgassing during a compression test depends greatly on the rock type. Rocks v/ith interconnected

pores tend to release a smaller amount of gasses because of

the ease with which they are removed during preconditioning

and pumpdo^-m. Whereas relatively dense rocks with interstitially held gas loads tend to release larger amounts during

the deformation and failure process (_5) . Mass spectrometer

analyses indicate that for the three rock types used the

principal outgassing components were: water vapor, nitrogen,

hydrogen, and carbon dioxide (Fig. 10) . The principal increase

in outgassing during the compressive loading cycles was con- tributed by water vapor and nitrogen.

The prime objective of this presentation is to describe

the problems associated with research in rock properties in

simulated space vacuum environment that were encountered 4-n our study on the compression properties of simulated lunar

rock in ultrahigh vacuum. The main emphasis is directed

toward obtaining the deformation and strength properties of

rock under carefully controlled conditions of vacuum. These

studies also have shown that the analysis of outgassing characteristics of rock in vacuum gives highly significant data on

the basic mechanism of rock deformation and failure.

ACKNOWLEDGEMENTS

The work presented in this paper was performed at the

Twin Cities Mining Research Center under a NASA contract on Multidisciplinary Research Leading to Utilization of Extra- terrestrial Resources.

The authors wish to express their appreciation for the

valuable consultation provided by W. W. Roepke during establishment of pumpdown procedures and for the efforts of

R. C. Rosenquist and K. G. Pung in standardizing and conducting

the test program. They also wish to thank L. Sunda for his

contributions towards the design of certain vacuum chamber

modifications

BIBLIOGRAPHY

1. Green, Jack. "Selection of Rock Standards for Lunar Research," Ann. N. Y. Acad. Sci. , v. 123, Art. 2, July 15,

1965, pp. 1123-1147.

2. Fogelson, David E. "Simulated Lunar Rocks," Proceedings , Working Group on Extraterrestria 1 Resources , Brooks Air

Force Base, Feb. 1968, NASA SP-177, pp. 75-95.

3. Thill, R. E., R. J. Willard, and T. R. Bur. "Correlation

of Longitudinal Velocity Variation with Rock Fabric,"

215
Journal of Geophysical Research , v. 74, No. 20, Sept. 15, 1969.

4. Vickers, B. L. , and R. E. Thill. "A New Technique for Pre- paring Rock Spheres," J. Sci. Instr. (J. of Physics E) , 1969, Series 2, v. 2, pp. 901-902.

5. Podnieks, E. R. , P. G. Chamberlain, and R. E. Thill.

"Environmental Effects on Rock Properties," Proceedings , Tenth Symposium on Rock Mechanics, May 20-22, 1968.

6. Roepke, W. W. , and C. W. Schultz. "Mass Spectrometer

Studies of Outgassing From Simulated Lunar Materials in

Ultrahigh Vacuum," Proceedings , 14th National Vacuum

Symposium of the American Vacuum Society in Kansas City, Mo. , October 4-27, 1967.

7. Roepke, W. W. "Friction Tests in Simulated Lunar Vacuum,"

presented at the Seventh Annual Meeting of the Working

Group on Extraterrestrial Resources (WGER) in Denver,

Colorado, June 17-20, 1969. To be published as a NASA - SP.

216
Fig. 1--Ultrahigh Vacuum System.

Mechanical

roughing pump Alphotron gage

Moss

spectrometer head

T.C. gage [_|

Vacuum

chamber

Cryosorptior pumps Moss

spectrometer

controls

I Camera

Sublimation head

Kreisman gage

Sublimation

controls

Bokeout

controls

"I Transformer

[

gage

Ion pump

Recorder

Kreisman gage

controls

gage

controls

Ion pump

controls Fig. 2--Ultrahigh Vacuum System.

217

Fig. 4--Specimen in Position Within Vacuum Chamber.

219
Fig. 5--Vacuum Chamber Positioned in Compression Testing Machine

Automatic

load programmer

Hydraulic

actuator

Hydraulic power

supply

Servo

valve

Fig. 6--Compression Testing System.

220
Fig. 7 --Typical Pumpdown Curves.

221
AXIAL DEFORMATION, in

AXIAL DEFORMATION, mm

Fig. 8--Typical Load - Deformation Curve for Tuff.

0 2 4 6 8 10 TIME, minute

Fig. 9--Chamber Pressure During Compression Test on Tuff.

222
CLS 0

<

(a)

a:

H20

N2

02 18 28 44

ATOMIC WEIGHT,nnole

At Initiation of Compressive Loading.

Q.<

20 -

10

0

1 1 1

I

N2

L

CO2

j<—

1

1

1 02 18 28 44

ATOMIC WEIGHT^mole

(b) - At Specimen Fracture.

Fig. 10--Mass Spectrometer Traces for Tuff During Compres

Test.

223

Paper No. l6

FRICTION BETWEEN SOLIDS AND SIMULATED LUNAR SOILS IN

ULTRAHIGH VACUUM AND ITS SIGNIFICANCE FOR THE DESIGN

OF LUNAR ROVING VEHICLES

Leslie L. Karafiath

REFERENCE: Karafiath, Leslie L., "Friction between

Solids and Simulated Lunar Soils in Ultrahigh Vacuum

and Its Significance for the Design of Lunar Roving

Vehicles/' ASTM/IES/AIAA Space Simulation Conference,

14-16 September 1970.

ABSTRACT: The initial and kinetic friction between

various solids and ground basalt, simulating the

lunar soil, was measured in ultrahigh vacuum with a rotating disk-type apparatus designed for this purpose. Pressures below p = 10 torr have been

routinely reached with the vacuum system consisting

of a turbomolecular , a getter -ion and a liquid-N2~

cooled titanium sublimation pump. The testing program included tests with steel, titanium, and fiberglass disks and ground basalt in two particle size

ranges, coarse (250-500^i) and fine (38-62^1). The

experimental results show that both the initial and

kinetic friction invariably increase in ultrahigh

vacuum. For the range of normal stresses applied in

the experiments, the total kinetic friction was highest for fiberglass followed by titanium and

steel; the total initial friction was highest for

fiberglass followed by steel and titanium.

Research Scientist, Grumman Aerospace Corporation,

Bethpage, New York 11714

225
KEY WORDS: adhesion, friction, ultrahigh vacuum,

basalt, fiberglass, steel, titanium

^ INTRODUCTION

On earth it is possible to perform full scale

performance tests of off -the -road vehicles to determine traction, slip, and other wheel characteristics

under various soil conditions. For the design of

lunar roving vehicles these characteristics have to

be estimated on the basis of wheel geometry, soil

properties, and the friction that develops between

the soil and the contacting solid surfaces of the

wheel. On earth the layers adsorbed on both the

surface of the soil particles and of the solid govern the mechanism of friction; in the ultrahigh vacuum environment of the Moon these layers are absent allowing direct contact between the solid and

mineral constituents of the soil, and thus a change

in the frictional properties of these materials.

The purpose of the experimental investigations re- ported in the paper is to determine the friction between various solids and simulated lunar soils in

ultrahigh vacuum environment.

METHOD OF FRICTION MEASUREMENT

The frictional resistance which develops between solids and granular materials upon mutual displacement has the following characteristic features.

At the beginning of the relative displacement the

frictional resistance rises sharply and reaches its

peak value after a small displacement. Upon con- tinuing displacement between the solid and the granular material the frictional resistance may stay at

this peak level (curve a in Fig. 1), or, as is more frequently the case, drop somewhat and stabilize at a lower level (curve b in Fig. 1).

In this paper the term initial friction is used

to designate the peak value of frictional resistance.

226
Sometimes this initial friction is considered as static friction, since very little mutual displacement is necessary for its mobilization. The fric"

tional resistance which develops when continuous

motion of the sliding solid is maintained is termed

kinetic friction. If the frictional resistance

shows a linear relationship with the normal load,

and uniform distribution of stresses is assumed,

then

T = a + i-La , (1)

where t is unit frictional resistance, a is adhesion, a is normal stress, and is coefficient

of friction (initial friction is indicated below by

subscript i and kinetic friction by subscript c)

For the purpose of estimating shear stresses at

the soil-^heel interface at various performance

levels it is necessary to know the magnitude of both

the initial and kinetic friction. The testing apparatus designed to determine both of these values

under both atmospheric and ultrahigh vacuum conditions is shown in Fig. 2. It consists of two sections: a rotating section with an annular metal

disk 4 cm in diameter and a surface area of

10 cm2^ and a strain gauge balance carrying a stainless steel cup 6 cm in diameter and 1.8 cm

deep, containing the soil specimen. The rotating

section is free to move up and down so that the test

surfaces can be separated during baking of the vac- uum system. Loads were varied by changing weights.

For transmission of motion into the vacuum chamber,

a flexible coupling connects the drive shaft of a variable speed motor to the input shaft of a NRCrotary harmonic feedthrough.

A diaphragm section about the center of the

strain gauge balance is instrumented to measure

axial loads; a thin -walled cylindrical member is instrumented to measure torsional forces. The important feature of the testing apparatus shown in

Fig. 2 is that the geometry of the contacting surfaces does not change upon mutual displacement allowing the determination of both the initial and kinetic

227
friction. The kinetic friction observed in the

tests corresponds to the friction at 100 percent

slip in locomotion theory.

Assuming uniform distribution of the stresses

beneath the disk the following relations hold for

the disks used in these experiments:

T(g/cm^) = 0.066 T(g cm)

a(g/cm^) = 0.103 N(g) , where T = torque and N = normal load. It is

noted that any error arising from the nonuniform

distribution of stresses is negligible as was shown

for triangular distribution in (1)

The determination of the friction between

solids and granular materials by the above method is

valid as long as the shear failure induced by the

rotation of the disk occurs at solid-granular material interface. If the frictional resistance at

this interface and the normal load is relatively

high, shear failure may occur laterally in the granular material before the frictional resistance

reached its limit at the interface. Figure 3 shows

the principal stress planes for these conditions.

Lateral failure in the granular material is accompanied by excessive sinkage of the disk; in the ex- periments reported herein both the normal load and

the frictional resistance was relatively low and no lateral failure occurred.

ULTRAHIGH VACUUM SYSTEM

The ultrahigh vacuum system includes a working

chamber equipped with several different pressure

gauges, including a residual gas analyzer, electrical and instrumentation feedthroughs , internal

heaters, a gas handling system, and various pumps.

For baking the whole system, a combination of ovens,

heating strips and tapes, and baking mantles are available. Normally, a turbomolecular pump with a

228
nominal pumping speed of 260 liter/s is used for

roughpumping and pumping during part of the baking

cycle. With very fine soils, the start of the

turbomolecular pump is preceded by the activation of

two sorption pumps in order to avoid disturbing the

specimen. Ultrahigh vacuums are obtained by the op"

eration of a getter -ion pump with a nominal pumping

speed of 270 liter/s and a liquid-N2 "cooled titanium sublimation pump with an estimated pumping

speed of 700 liter/s. Pressures below 1 X 109

torr were routinely reached within 2 or 3 hours

after baking.

TESTING PROCEDURES

a . Preparation and Properties of Soil Samples

Chunks of basalt from Somerset County, New

Jersey were comminuted in air and sifted to obtain

the desired grain sizes. The container was filled

with 66 g of loose soil and carefully evened to

the same level in each test with the help of an Airite vibrator. Table 1 shows data pertinent to

the preparation of samples.

To compare the frictional resistance of solids

on granular soil specimens with the internal friction of the soil, triaxial tests were performed on both coarse and fine basalt in air-dry condition.

The shear strength envelope for both materials was

found to be a curve which at pressures lower than

1 kg/cm could be approximated by a straight line

corresponding to a friction angle of 45 degrees.

Sieve analyses performed on specimens tested

either for solid-on-soil or for internal friction

did not show a discernible change in the grain size

distribution, an indication that the stresses applied in the tests did not break down the individual

grains significantly.

229
b. Preparation and Properties of Solid Disks

The annular metal disks were fabricated by

machining, and were checked for concentricity before

and after polishing. Only disks with a total runout

of less than 0.01 mm were accepted; their degree

of finish was measured with a prof ilometer . Most

tested disks had finishes between 0.050 and 0.1m,

(rms) . The annular fiberglass disks were cut from

^-in. thick laminated fiberglass (cloth fibers)

sheets and fastened by two screws to aluminum disks

fabricated in a previous program. The sides of the

fiberglass sheet differed slightly in texture; in

the friction tests, the side with the coarser texture was used as the sliding surface.

The steel disks were manufactured of steel

AISI 1020, the titanium disks of the titanium alloy

Ti-6A1-4V. Properties of the latter are given in

(2) . The metal disks were submitted to a thorough

cleaning procedure after manufacturing and prior to

testing

The fiberglass sheets were manufactured according to Grumman specifications (3) . Outgassing ex- periments performed with fiberglass material showed

that with the fiberglass inside the chamber the ultimate pressure reached with the vacuum system was about 1^ decades higher than with empty chamber

(4).

c . Pumpd own

To avoid lengthy pumpdowns, the granular

basalts used in the vacuum tests were outgassed in

a special desiccator prior to their transfer to the

experimental chamber. Besides reducing pumpdown

time considerably, the nearly constant gas loads

produced by the processed specimens made the achieve

ment of a certain uniformity of the vacuum history

of each test easier.

230
Baking temperatures varied from section to

section of the vacuum system; their limitations were

set by high temperature sensitive components in the

system; e.g.^ strain gauges and ion pump magnets.

During the entire baking cycle, a temperature gradient ranging from 250° to 300°C at the TSPcryopump combination to 200° to 220 °C at the ion

pump was maintained. The chamber itself was usually

kept at 200° to 210°C. Total baking times ranged

from 22 to 92 hours. When the system pressure

showed a tendency to fall and reached the 10"^ torr

level, the heat was shut off. From this point, it

took approximately 3 to 6 hours to produce a reasonably stable pressure. Although the achieved

ultimate pressure varied from test to test within a decade, they were without exception in the 10 "^^ torr range, or lower. Pressure readings were taken

with a G.E. Model 22 GR214 trigger gauge.

The pressure was normally kept below 1 x 10"^

torr for an exposure time of about 20 hours. In

two instances, however, scheduling necessitated performing a test exposure times between 3 and 4 hours;

the effect, if any, of this shorter exposure time on friction was within the experimental error.

d. Performance of Tests

Once an acceptable pressure level was obtained,

the loaded stationary disk was lowered to rest on

the granular basalt for 5 minutes; then three clockwise revolutions were executed at 3 rpm. The

torque required to rotate the disk and the normal

load were recorded throughout the test on a Brush

Mark 280 dual recorder.

Pressure evolution during the tests was recorded on a G-14 Varian recorder. At the start of

disk rotation, pressure readings increased by a factor of 2 or 3; as soon as motion stopped, the

pressure decreased rapidly until it reached the

original pressure level.

231
RESULTS OF TESTS

a . Repeatability of Tests

The testing procedure as described in the preceding section consists of several operations each

of which may be a source of error in the measured

friction. To determine the accuracy of the whole

system a series consisting of 10 tests in ultrahigh vacuum and 10 parallel tests in air was performed on coarse basalt with a normal load of 682 g. The

average values of the measured torque as well as the

standard deviation of the 10 observations are shown

in Table 1, together with estimates of the percentage error intervals that occur with 95 percent probability when averages of observations are used to

represent the theoretical mean. Because of the time required for bakeout and

pumpdown it was not possible to incorporate 10 repetitions for each measurement in the testing program.

The error interval for a series of 5 tests is 1.7

times that shown in Table 1, and for 3 tests it is

3.5 times greater.

b . Testing Program

Table 3 shows the number of tests performed to

date in ultrahigh vacuum with various solids and

particle sizes of ground basalt. For each test in

ultrahigh vacuum a parallel test was performed in

air.

c. Presentation of Test Results

The results of tests performed on coarse basalt

in both ultrahigh vacuum and air are represented in

Figs. 4 through 1 , which show the averages of the

measured torque values for each normal load and

their best fitting linear relationships, determined

by the method of least squares.

232
The coefficient of friction, as defined by

Eq. (1) is determined by the slope of the best

fitting straight line and is shown in the figures.

The adhesion can also be evaluated from Eq. (1)

However, in many instances the adhesion determined

by the best fitting straight line was found to be

negative indicating that at low loads the relationship between normal load and frictional resistance

is not linear. For this reason the best fitting

straight lines are shown only in the range of normal

loads where it represents the best approximation to

the test results.

Figure 4 shows the effect of ultrahigh vacuum

on the friction between steel and coarse basalt.

The best fitting straight lines indicate that the

coefficients of friction, as defined by Eq. (1), were not affected by the ultrahigh vacuum. The adhesion due to the ultrahigh vacuum was found to be

aj_ = 18 g/cm2 for the initial friction and

a^ = 9 g/cm^ for the kinetic.

In the tests performed on fine basalt, the

frictional resistance versus displacement curve was

invariably found to be that of type "a" in Fig. 1.

For these tests = i-l^. The results of these

tests are shown in Fig. 5; the coefficient of friction is i-Lj^ = 0.20 for ultrahigh vacuum and

\il = 0.18 for air. The adhesion represented by the

distance between the two straight lines varied from

a = 6 g/cm^ to 9 g/cm^, within the range of

normal loads

The scatter of the measurements experienced

with the fine basalt was greater than that for the coarse basalt. This is attributed to the difficulties of controlling the surface of the soil bed.

Because of the lower gas permeability of the fine

basalt, gas pressures may build up within the soil

bed and loosen the surface during outgassing.

The results of two series of tests performed on coarse basalt with titanium disks in both ultrahigh vacuum and air are shown in Fig. 6. In these two

233
series of tests the difference between the initial

and kinetic friction was in the range of the noise

level of electronic recording and, therefore, was

not evaluated quantitatively.

The results of two series of tests, one in

ultrahigh vacuum, the other in air, on coarse basalt

with fiberglass disks are shown in Fig. 7. In con- trast to the tests with titanium disks, well pronounced initial friction was observed in all fiber"

glass disk tests. The coefficients of initial and

kinetic friction are shown in the figure.

The results of the first two tests with fiberglass disks performed under the same test conditions were found to be almost identical within the ac~ curacy of recording. Further tests at the two other

normal loads showed the same degree of repeatability,

therefore, it was deemed unnecessary to perform more

than two tests at each normal load.

EVALUATION OF TEST RESULTS

The test results show that the frictional resistance between the various solids and soils tested

invariably increases in ultrahigh vacuum but in a different way for the various combinations of materials. A valid comparison of the frictional properties of the various materials in ultrahigh vacuum

can best be made by evaluating total frictional resistance within the range of normal stresses of interest. The normal stresses in experiments varied

from 46 to 91 g/cm^, approximately corresponding

to the expected range of normal stresses under the

wheels of a lunar roving vehicle. Thus comparison

of the total frictional resistance within the range

of normal stresses applied in the experiments is

also meaningful for the applications.

Figure 8 plots average shear stresses representing the total frictional resistances developed

by various materials sliding on coarse basalt in

ultrahigh vacuum against the normal stresses. The

upper graph shows the kinetic, the lower the initial

234
frictional resistances. As can be seen from Fig. 8^ fiberglass develops more initial and kinetic friction than any of the metals tested. This may^ however, be partly due to the texture of the fiberglass.

Comparable results were obtained with the titanium

and steel disks, both having the same surface finish.

The titanium disk exhibited somewhat higher kinetic

friction than the steel. The initial friction developed by the titanium disk, on the other hand, was

so low that it could not be quantitatively evaluated;

the dashed line in Fig. 8 represents a very rough

guess offered for comparison only. The differences

in friction exhibited by these two metals are probably due to their different hardnesses.

From the point of view of lunar locomotion it

is significant that both the initial and kinetic

friction increased in ultrahigh vacuum. Preliminary

tests performed with titanium disks and ground basalt simulating the particle size distribution of

the Apollo 11 samples showed friction values very

close to that found with coarse basalt.

APPLICATION OF TEST RESULTS TO

THE DETERMINATION OF TRACTION

In the application of experimentally obtained

frictional relationships to the determination of

traction that a driven wheel can exert, it is essential to pay attention to the following considerations .

Friction and Slip

The kinetic friction determined by the experiments corresponds to 100 percent slip of the wheel.

There is no direct correlation between slip and the

rotational displacements associated with the development of initial friction in the tests.

235
Lateral Failure

The results of friction experiments apply to

wheel-soil friction as long as sliding occurs only

at the wheel -soil interface. If the stresses transmitted to the soil by the wheel cause lateral fail- ure in the soil mass before the full frictional resistance at the interface is mobilized, the traction

exerted by the wheel can no longer be computed on

the basis of the friction developed at the contact

area. Failure in the soil mass would occur along an oblique plane causing lateral displacement of the

soil; hence the term lateral failure.

An analysis of potential lateral failure beneath a wheel may be made by considerations similar

to that applied to the friction experiments (Fig. 3).

The principal stresses in the plane of the normal

load and shear stress at the interface may be determined and the minor principal stress compared with

that in the lateral direction.

Consideration of Tread or Lugs

Tread or lugs force the sheared surface to pass

through the soil between the tread bars or lugs

mobilizing thereby the internal frictional resistance of the soil which is generally higher than

that between solids and soil. The total traction

which a wheel with tread or lugs can exert is com- posed in such a case of two components, one due to

the frictional resistance on the face of the tread

bars or lugs and the other one due to the internal

frictional resistance of soil between the tread bars

or lugs (Fig . 9) . An estimate of the contribution of the two com- ponents to the total traction may be made by con- sidering the distribution of normal stresses at the

sheared surface. Normal stresses at the face of the

lugs will be generally higher than between them because the soil immediately beneath the face of lugs

is compressed more than between lugs. Experiments

with treaded tires showed high stress concentrations

236
at the face of lugs. A crude approximation of the

distribution of normal stresses may be made by assuming that a firm bottom exists at some depth beneath the wheel and the normal stresses are inversely

proportional to the depth of the compressed layer

(Fig. 9). The total traction can then be computed

by applying the respective relationships between

frictional resistance and normal stress to the areas

in question.

CONCLUSIONS

An experimental apparatus suitable to determine

in ultrahigh vacuum both initial and kinetic friction between solids and granular materials under

various normal loads has been developed. The ex- perimental results clearly show that with every

material tested in UHV there is an increase in frictional resistance between the solid disk and the

granular basalt simulating lunar soil. Since the

magnitude of this increase in relation to the normal

stress varies with each of the solid materials

tested, no general conclusions as to effect of ultra

high vacuum on the friction between solids and soils can be drawn.

The friction that develops between a wheel operating in ultrahigh vacuum and the soil significantly affects wheel performance. In the selection

of materials for the wheels of lunar roving vehicles

consideration should be given to the frictional

properties that these materials exhibit in ultrahigh

vacuum.

ACKNOWLEDGMENT

Acknowledgment is due to G. Mohr who designed

the ultrahigh vacuum system and performed part of

the experiments and to G. Homfeld who assisted in

the operation of the ultrahigh vacuum system and in

the performance of experiments.

237
REFERENCES

1. Karafiath, L. L., Some Aspects of the Determina - t ion of the Coefficient of Friction between

Metals and Soils and Its Application to Lunar

Locomotion Studies , Grumman Research Department

Memorandum RM-379, October 1967.

2 . Properties of Ti-6A1-4V Titanium Engineering

Bulletin No. 1 , Titanium Metals Corporation of

America, New York.

3. Grumman Standard Specification No. IIIOOA,

"Glass Fiber Base-Plastic Laminates for Structural Application," June 1968.

4. Karafiath, L. L. and Mohr, G., The Effect of

Ultrahigh Vacuum on the Friction between Various

Solids and Granular Soils , Grumman Research Department Report RE-376, March 1970.

238
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240
Table 3 TESTING PROGRAM — NUMBER OF TESTS PERFORMED IN UHV

Coarse Basalt:

Disk Normal Load (g)

455 682 910

Steel

Titanium

Fiberglass

10

5

2

Normal load 7gs higher than shown

Fine Basalt:

Disk Normal Load (g)

455 682 910

Steel 3 5 4

241
Fig. 2 Friction Measuring

Apparatus

= Vertical Normal Stress

o- 1 = Lateral Stress T = Shear Stress a ptTj = Principal Stresses a T, <T,, o- in Plane 1 V ' r 0 ^ CT

|, oJ in Plane 2 Fig. 3 Planes of Principal Stresses:

Plane 1 ctq ^ ^1' ^0 ^3

Plane 2 < oq; 03 = a^^

242
300 400 500 600 700 800 900 1000 a Normal Load (N), g Fig. 4 Friction Bet:\^7een Steel and Coarse Basalt

Results of Test

Average of Tests

Best Fitting Straight Line UHV

AIR

10O2OO3OO4CC50O6OO70O80O9OO

Normal Load N , g Fig. 5 Friction Bet\^7een Steel and Fine Basalt.

Results of Tests

700r

a.> Best Fitting Straight Line UHV

AIR 0

^ u =.26

» 500 Normal Load, N

1000

g Fig. 6 Kinetic Friction Bet\v^een Titanium and Compacted Coarse Basalt - Results of Tests

243
Average

of Tests Best Fitting Straight Line

Initial UHV

Friction AIR

Kinetic UHV •

Friction AIR o

0 Oq

10 20 30 oz

6 100 200 300 400 500 600 TOO 800 900 1000 g Normal Load, N Fig. 7 Friction Between Fiberglass and Compacted

Coarse Basalt - Results of Tests

I

40 i 30

Fiberglass

Normal Stress, g/cm

Initial Friction

Fiberglass

60 70

Probable Actual

Distribution of Normal Stresses Fig. 9 Treaded Wheel

Distribution of Normal

Stresses

Normal Stress, g/cm Fig. 8 Comparison of

Friction Between Coarse

Basalt and Various Materials in Ultrahigh

Vacuum

244
Paper Mo. 17

REAL TIME SIMULATION OF ATMOSPHERIC BALLOON

ENVIRONMENT^

A.J. Paillous and M. Roussel

REFERENCE: Paillous, A.J. and Roussel, M. , "Real

time Simulation of Atmospheric Balloon Environment,"

ASTM/IES/AIAA : Space Simulation Conference , 14-16 September 1970

ABSTRACT: A simulation chamber was designed in order to study the behaviour of several balloon materials during long time flights. Solar ultraviolet

radiation, pressure, temperature and atmospheric

composition are the simulated constituents of the

atmospheric balloon environment. Samples are irradiated without mechanical strain. A computer program,

taking the values of different parameters (of the

environment and of the material) against wavelength

into account, gives the information for conducting

the simulation. Experim.ents are performed in realtime. Exact ultraviolet energy doses, in each spectral wavelength interval, between 300 and 410 nm,are

delivered every day to samples. A night-day cycle is

simultaneously achieved with a thermal cycle.

KEY WORDS: Simulation, Balloon, Solar ultraviolet

Radiation, Atmospheric environment, Ozone, Temperature, Plasticmaterial , Computer.

ONERA/CERT - Departement d 'Etudes et de Recherches

en Technologie Spatiale, 51, rue Caraman 31-TOULOUSE

FRANCE

2 This work was supported by a ONES contract

245
1 - INTRODUCTION

The Eole Project (1) includes the flight of se- veral hundred superpressure balloons at a constant

altitude (200 or 300 mbar) for some months (Superpressure balloons are non-extensible balloons which

are sealed to prevent gas release. See reference (2) ) In order to test and select the balloon material

a chamber allowing the simulation of a real long-lasting flight was realized. This chamber can reproduce^

in the laboratory, all the parameters of the balloon

environment at a chosen altitude [i.e. solar ultraviolet radiation, pressure, temperature and atmospheric composition (ozone concentration) ] . All these

components have a role in the degradation of the physical properties of plastic films used in the present

state of superpressure balloon technology.

The selected exposition process enables one to

carry out a real-time study. Exact ultraviolet ener- gy doses, in each spectral wavelength interval bet- ween 300 and 410 nm, are delivered every day to

samples. Samples are alternately illuminated and darkened ; a thermal cycle is simultaneously used.

2 -OVERALL SIMULATION EQUIPMENT

The simulation equipment is shown in Figure 1

It includes a stainless steel irradiation chamber, a light source, a cooling system and an ozone source.

2-1 Irradiation chamber : The stainless

steel irradiation chamber is formed by two portions

of a sphere separated by a cylindrical part (diametor

120 cm) . All internal surfaces are black.

In the chamber, twelve sample-holders (12x21 cm) are regularly laid out around the light source

(Figure 2) . They can be adjusted at a variable distance from the source (32 - 41 cm) . Taking the light

The numbers in parentheses refer to the list of re- ferences appended to this paper.

246
emission diagram of the source into account, the

shape of the sample-holder was found to produce a high uniformity of light on the sample (4%) . Each

sample -holder is connected to the cooling system by

a flexible pipe.

The irradiation chamber is also equipped with

several pass-throughs for thermocouples, four pyrex

windov7s for direct observation of the samples, a quartz v/indow for the determination of the ultraviolet intensity and a vent valve. These elements

are seen in Figure 1

2-2 Ultraviolet source : A 2 500 v/atts

Xenon short arc lamp (OSRAM, XBO 2 500 Watts) is

used as the source of ultraviolet radiation. The

lamp is centered in a quartz cylinder through which

an air flow cools the lamp and carries along the ozone formed in its vicinity.

2-3 Sample positioning and temperature

control : The film samples are fixed

close to the sample-holders by clips and small

U-Shaped springs (Figure 3). Thus, samples are irradiated without mechanical strain.

The sample-holders are cooled by a fluid circulation allowing the temperature control of samples.

The heat exchanger used for this purpose has an output range of from -80°C to + 100°C. The irradiation

temperature is greater than the night temperature.

This thermal cycle is obtained by an automatic change

of the working conditions of the cooler when irradiation stops. Two temperature control devices are employed. Thermocouple leads run from behind each

sample to a tem.perature recorder ; these thermocouples,

placed between two plastic films (Figure 3) , give a fair estimation of the sample temperature. An other

thermocouple is placed in the return pipe of the cooling fluid. It is connected to a safety device which

stops irradiation when the temperature is higher

than a fixed value (See Figure 2)

2-4 Atmospheric composition control : Total pressure and ozone partial pressure, in the

chamber, are controlled every day during long periods

of experimentation.

A mercury manometer measures pressure in the

chamber which is evacuated by a mechanical pump.

247
The ozone concentration in the chamber is determined by a chemical method (3) . A constant volume of

the chamber atmosphere is drawn at a constant rate

through a solution contained in three gas-traps

(Figure 1) . The absorption of the solution is then

read at 412 nm in a spectrophotometer. The loss in

extinction per cm after sampling gives the ozone concentration, vzhich can be adjusted by the introduction of pure ozone, or pure air, and further pumping,

2-5 Measurement of ultraviolet light

intensity : The spectral energy distribution of our light source, in the irradiation device, at the sample location, has not been yet accurately measured, because we are lacking a convenient

spectrometric device. As an approximation of the

truth, we assumed that the Xenon short arc had an output similar to the one given in previously published works (4)

Before a long experiment is carried out, and while the irradiation chamber is open, a calibrated

thermopile is used to measure the total light intensity given by the Xenon source (with the quartz cylinder and pyrex filters) on the sample-holders. If

necessary, the positions of the sample-holders are adjusted in order to achieve the desired value.

During experimentation, solar cells, fixed on

a sample-holder (see Figure 3) , monitor the light

intensity through the solar region. One of these solar cells is behind a wide band filter (300-400 nm)

and gives an estimate of the intensity of ultraviolet

radiation. Another device is used behind the quartz

window to give the variations in ultraviolet intensity of the Xenon arc lamp. This device consists of

a sensor cell attached to a compact metering unit

and an ultraviolet filter. To measure the variation

in ultraviolet intensity of the lamp, two readings

are taken ; one reading with the sensor cell by itself, then a second reading while holding the ultraviolet filter between the source and the sensor cell.

The difference between these two readings gives a proportional value of the ultraviolet intensity of the

Xenon arc lamp.

248
3 -PRINCIPLE OF THE ULTRAVIOLET SOLAR SIMULATION

The solar spectral irradiance for a given wavelength interval, and the solar constant, are well

known (5) . Terrestrial albedo and light attenuation

by the atmosphere are also tabulated for various

altitudes (6)

The ultraviolet energy dose received by the

balloon at a given altitude is a function of solar

spectral irradiance, terrestrial albedo and light

transmission through the atmosphere.

All these terms are used in a computer program

which gives at any latitude, every day in the year,

every hour in the day, at a fixed altitude, the ultraviolet flux density on a horizontal unit surface

(top of the balloon)

The computer v/rites also the energies I^^

received by this unit area ( I are expressed in

2 i J/m ) for different contiguous wavelength intervals

AX^ , during one day ; this calculation takes the

solar radiation reflected by the earth through the

balloon into account.

Then the program assumes that the Xenon arc source provides to the sample, in a given time T and

in the wavelength interval 350 - 370 nm, exactly the

same energy as provided by the sun to the balloon in

24 hours ; it takes also the 1.. values and the

A A . spectral irradiance^ of the Xenon arc lamp into

account. Then the computer writes for each AA^, the ^ duration T,. of the illumination required to deliver

AX . the I energy.

AX.

4 - VALIDITY OF THE SIMULATION

A first experiment was carried out in order to

simulate eauatorial conditions at 300 mbar. Materials

were irradiated 12 hours out of 24, during 6 months.

The spectral irradiance of the Xenon arc lamp was computed from the ratio of lamp energy to Johnson's

solar irradiance values (4) , taking the new values

of solar spectral irradiance into account (5)

249
4-1 Ultraviolet Solar Radiation Simulation

The data used by the computer program for this experiment and the results of the calculation are respectively given in Table 1 and Table 2. In these conditions, three filters are required for the ultraviolet

light simulation.

Spectral transmission plots of these filters

are shown in Figure 4. The pyrex filters fixed in

the chamber (see Figure 2) cut off the wavelength

shorter than 300 nm. The two other filters, fixed on

a polygonal support (Figure 5) can be interposed bet- ween the lamp and the samples, to cut off respectively wavelengths shorter than 315 and 333 nm. The Xenon short arc lamps do not possess a sa- tisfactory solar spectral match in the visible and near infrared regions (800 to 1 100 nm region) (4)

But it is quite unimportant for the sample degradation, which is due to ultraviolet radiation when

samples are very well cooled. So, as a matter of fact our purpose was not the simulation of the whole spectrum at a given altitude, but the simulation of only

the ultraviolet solar radiation at this altitude.

The total light intensity given by the experimental device (having only the pyrex filter) and

measured by the thermopile, is adjusted in order to

obtain on the samples, in the wavelength interval

350 - 370 nm the exact computed value. In the wavelength range 300 - 335 nm an exact agreement between

the I^^ values calculed and those given by the Xenon

source can be obtained by interposition of the mobile

filter during a time in accordance with the computed

T .In the wavelength range 300 - 410 nm, for the

considered altitude (300 mbar) , the agreement is

theoretically better than 20%, provided that there

is no spectral aging of the light source. In order

to be sure that the facility achieves its intended

purpose of adequatly simulating the ultraviolet solar

radiation, an accurate measurement of the spectral

irradiance, at the sample location, ought to be used,

at first at the beginning of irradiation, and then

at even time intervals. We plan to use soon a conve- nient spectrometric device. Meanwhile a preventive

replacement of the lamp must be done every 250 hours

(20 startings)

250
4-2 Atmospheric environment simulation : The other anvironmental parameters are ozone concentration and air temperature. The following values are given in the Handbook of Geophysics (6), at 300 mbar,

at the Equator : Ozone concentration : 0.2 to 0.4 ppm

Air temperature : -37°C (January) to -33**C

(July)

Actual temperatures of the balloon materials in

flight are a function of the thermo-optical properties of these materials.

The experimental device allowed to achieve the

follo>7ing conditions during the long lasting (6 months)

experiment : Pressure : 300 + 10 mbar

Ozone concentration : 0.1 to 0.5 ppm

According to the facts that different materials

are simultaneously tested in the chamber and that

there is only one cooling circuit a same temperature

was maintained for all the samples : for nigths , the

chosen temperature was approximately the average air

temperature (-35°C) ; during the days, the sample

temperature was allowed to be higher and higher as the samples were aging (the difference with the average air temperature was fixed at + 1®C at the beginning of irradiation, and at + 10**C three months later) ; this is in order to simulate the increase in

solar absorptance with time.

5 - CONCLUDING REMARKS

The equipment, described above, allowed a con- venient real-time simulation of a six-months flight

at 300 mbar at the Equator. All the environmental

parameters, which contributed to the balloon material

degradation, were reproduced : ultraviolet energy

dose, pressure, atmospheric composition, sample temperature . This equipment can be adapted for the ultraviolet Solar Radiation Simulation on balloon, satellite

or aircraft materials, in varying atmospheric conditions (pure air, ozone, or others gases) for different pressures from one atmosphere to vacuum (a high

251
vacuum system can be adapted) , with a night and day

cycle, and a thermal cycle (from -120*'C to + 150°C) .

ACKNOWLEDGEMENT

The authors wish to acknowledge the comments

and suggestions offered by Mr. D.W. CONNER of NASA

Langley Research Center.

REFERENCES

(1) P. MOREL - "Preliminary flights of Eole Balloons"

COSPAR Space Research, London 1967

(2) VINCENT E. LALLY - "Superpressure Balloons for

Horizontal Soundings of the atmosphere"

NCAR - TN 28 - June 1967

<3) T. NASH - "Colorometric determination of ozone by diacetyl-dihydrolutidine

Atmospheric Environiaent , 1 , 679 (1967)

(4) M. LILLYWHITE, R. Mc INTOSH, C. DUNCAN and

J.COLONEY - The Journal of environmental Sciences

p. 9, Avril 1968

(5) "The solar constant and the solar spectrum measured from a research aircraft at 38 000 feet"

GSFCX 322-68-304

(6) "Handbook of geophysics and Space Environment"

Scientific editor SHEA L. VALLEY - Air Force

Cambridge Research Laboratories - Mc Graw Hill

Book Company, Inc.

252
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253
Date : 84th Day (equinox)

Altitude : 300 mbar

Latitude : 0 00 degree

Daytime : 12 00 hours

Energies I . received bv a unit area of the balloon during 1 day : i AX.^^^^ ^ I,, ( T/m2)

^ 9 2SO-300 0. S1'^1431S X 10^

300-31'^ 0 noisyoQo X 10^ 31^-33" O.343730QO X 10^

33^-3^0 0.3'^S^1321 X

10'' O.S2^27Q01 X

1.0*"' 3SO-370

370-300 O.S7S7Q303 X

10^^ 300-410 0.7n3f''S44''^ X

10*' 410-400 0.433'^<'^'^'^^ X 10^ Irradiation time T - 12 00 Hours

Necessary light intensity for accurate simulation in the wave- length interval 3^)0-370 nm = 0.1223 x 10 W/m^

AX. Xenon Arc^Irradiance T Simulation Time

^(nm) (V/m') ^1 (Ho irs)

2SO-3OO 0.4';7s X 10^ 0 00

300-31 S 0 9=?7'^ X 10^ 3 00

31S-33^ 0.02S7 X 10^ 10 31 33S-3SO 0 S172 X 10^ 12 OS 3^0-370 0.1223 x 10^ 12 00

370-300 0,1320 X lO'^ 12.12

300-410 0,1400 x 10^ 14.23

410-400 0,n3^'^'^ X

10"^' IS 03

TABLE 2 - Calculated res ilts eiven

by the computer for

one day (equinox)

254
255
256
257
Transmission

100

so

'.0

•!0

'>0

'/ave length )

Figure 4 - Spectral transmission plots

of sharp cut filters

258
259

Paper No. 1R

ANALYSIS OF THE SIMULATION OF THE SOLAR WIND^

D. E. Zuccaro"^

REFERENCE: ZuccarO;, D. E.;, "Analysis of the Simulation of tha

Solar "Wind/' ASTM/IES/AIAA Space Simulation Conference, IU-I6

September 1970.

ABSTRACT: This analysis surveys the properties of the solar

wind, establishes a set of requirements for solar wind simulation, and develops a conceptual design of a simulator system.

The significant features of the design are the following. The

protons are formed in an r-f excited plasma discharge ion

source. A 20° deflection magnetic mass separator is used to purify the proton beam of other ions, energetic charge exchange

neutrals, and Lyman alpha photons. The use of a small diameter

beam permits differential pumping of the ion source and the

sample chamoer. The proton beam either can be expanded to flood the sample, or it can be scanned over the sample. KEY WORDS: solar wind, charge exchange, charge neutralization,

sputtering, proton sources, mass separators, ion optics, ultrahigh vacuum systems.

I. INTRODUCTION

The development of spacecraft thermal control coatings

requires the laboratory evaluation of the coating's resistance

to damage induced by solar photon and particulate radiation.

As the characteristics of the solar wind (or solar particulate

radiation) have been determined rather recently, various types

of simulators, which had different operating and performance

characteristics, were fabricated. The Hughes Research Labora-

"^lon Device Physics Dept., Hughes Research Laboratories,

Malibu, California.

2This work was performed under Contract NAS 2-5585.

261
tories has performed a study of the simulation of the solar

wind for the Ames Research Center of the National Aeronautics

and Space Administration. The goals of this program were to survey the properties of the solar wind, to establish the re- quirements for a solar wind simulator, to analyze the techniques and apparatus to be used in the system, and to establish

an optimized system and its operating techniques.

This publication is one of a pair that report the results

of this study. The companion paper by King(l) analyzes the

ion sources, mass separators, and ion optics systems needed to form a proton beam that uniformly irradiates a 10 cm diameter

sample. This paper will survey the solar wind properties, the

simulator requirements and, using the results of King's analysis, establish a conceptual optimized simulation system.

II. SOLAR WIND COMPOSITION

The particle environments in space that are of principal

interest to the designers of spacecraft are the solar wind and

the radiation that exists at synchronous orbit. The latter has

been summarized in an extensive survey by Stanley and Ryan.'^)

The properties of the solar wind, which have been measured by

a number of satellite probes, will be summarized in this sec- tion of this paper.

The solar wind is the term applied to the streaming plasma

that is evolved from the sun. Because the energy of the par- ticles is much greater than that which can be associated with

the corona temperatures, it is believed to result from a supersonic expansion of the corona's charged particles coupled to the sun's magnetic field. The plasma is neutral, having an equal number of positive charges and electrons per unit volume.

The ions and electrons have much different velocities.

The most abundant type of ion is the hydrogen ion, H^(i.e., a proton) . The second most abundant type is the doubly charged

helium ion, He^^ (i.e., an alpha particle). The He^"^ to ratio ( Hq/Tlp ) has been measured by the Vela 3A and 3B satellites. (3) During a two year period, the ratio varied over the

range of 1 to 8^, with an average value of 3-7%' The variation

is believed to reflect time changes in the plasma composition.

During periods of solar activity, i.e., solar flares, the plas- ma contains a much greater He^^ content. In a recent class 3B flare, an n^^/Hp ratio of 0.22 was observed. Under this con- dition nearly half the charge of the solar wind is carried by

the He^"*" ions.

The numbers in parentheses refer to the list of references

appended to this paper. - ^

262
Although there are spectroscopic data indicating the presence of many other elements in the sun, only multiply charged

oxygen ions O^"*", 0 , and o"^"^ have been identified^ 5) at this

time. The heavy ion component could he resolved only under

quiet sun conditions. During this period it was about 0.5^ (by

number ratio) of the proton flux.' 5; Other ions are believed

to be present, but their low relative abundance and the num- ber of ionic states makes it very difficult to identify them

by means of energy per charge detectors. The foil collector

experiments made during the recent Apollo flights should pro- vide more information about the heavy ion composition of the

solar wind.

The solar wind proton energies were first accurately measured by the Mariner 2 satelliteC?) in the final third of I962.

The average of the daily average proton energies for the period

was 1325 eV. The 3 hour averages ranged from a low energy of

5^0 eV (which is associated with quiet sun conditions) to a high energy of 3100 eV. The Vela 2A and 2B satellites made

about 13, 000 measurements(8) during a period of minimum solar

activity from July V^6k to July I965. The mean value of the

proton energy was 920 eV. The largest number of cases was for

a 550 eV particle energy, which is associated with the quiet

sun condition that was prevalent throughout most of the period.

This distrioution of proton energies is shown in Fig. 1. The energy of the He^"*" ions was k times that of ^the solar

wind protons. The highly charged heavy ions (o'^"'", 0*^, and

0^"^), which could be resolved only during quiet sun periods, (5)

had energies of about 20 times that of the protons.

The solar wind ion flux is about 2 x 10^ cm~^ sec~l during

quiet sun conditions .( This value is relatively constant,

and its variation with solar activity generally less than a factor of 2. The determination of the solar wind electron properties is

difficult because the low energy of the electrons means the

spacecraft potential can perturb the measurements and the solar

induced photoelectrons can cause erroneous measurements. The . most accurate measurements are the recent Vela 4-B observations'^

which showed that electron energy spectrum has a broad maximum

in the energy range of 20 to hO eV.

III. SOIAR WIM) SIMUMTOR REQUIREMENTS

The requirements for the laboratory simulation of the

solar wind are established mainly by the solar wind properties

that have been summarized in the previous section. Additional

factors such as the duration of the tests, the cost and com263
plexity of the equipment, the need for accelerated testing, and

the size and number of samples must be considered.

A. Particle Composition

The most important parameter to be specified is that of

the ion composition. All efforts to date to simulate the solar

vind have used either proton beams that were purified by mass

separation, or mixed (unpurified) ion beams. The latter con- tain molecular hydrogen ions {^2^ and H-^"*") as veil as other

ions.

Although it is possible to rule opt total simulation of

the solar ion plasma (E^, He^+j 07+, 0*-^, and O^"^) because of

the prohibitively expensive and complex apparatus required to produce such a plasma, it is important to note some of the possible effects of these minor constituents. For example, the

sputter yield of oxygen ions is 100 times that of protons. v^^)

Thus, the oxygen ion component would have about the same sputtering effect as the proton component of the solar wind. The

interaction of highly charged ions, such as o'^'*", 0^, and 05+,

with the surface results in the formation of x-ray photons or Auger electrons. Both of these can produce secondary ionization in the target. In this case, the effect of the heavy ions

is much greater than that of the protons.

Based on the present limitations, simulation of the solar

wind is limited to the use of pure proton beams. Nevertheless,

one of the first tasks of this optimized simulator will be to determine if the simulation of the minor constituents of the

solar plasma is necessary. This could be done with a less com- plex 0"^ or 0^+ source in combination with an x-ray source.

B. Particle Flux

The simulation of the solar wind requires a proton flux of

2 X 10^ cm~^ sec"-^, which corresponds to a ion current density

of 3 X 10"-^-^ A cm"^. Because this value is based on satellite

data that were obtained during a period of decreasing and mini- mum solar activity, it may be necessary to modify the value for

the period of maximum solar activity. The change should be a small one. A much higher flux level may be required to accel- erate the testing rate. This will be particularly true when

radiation resistant coatings are developed. At present, accelerated testing of photon (light) induced damage is performed

with 5 to 50 equivalent suns, where limitation is the output

of the sources and the reciprocity failure of the samples. A proton beam of 2 x lO-^-*- cm~^ sec"-^ (ion current density of

3 X 10"^ A cm"^), which corresponds to a level of 1000 times the

normal solar plasma flux, is readily ootained. This is an ac- ceptable design goal as far as the engineering aspects are con264
cerned. The operating limit may be less than 1000 times the

normal plasma flux because of reciprocity failure of the sample.

C. Particle Energy

The simulator should be capable of producing a proton

beam of uniform energy that can be varied over the range of ^00

to 3000 eV, which corresponds to the range of solar proton

energies. The mean energy of the solar protons is about 1000

eV. The simulator also requires an electron emitter which is

capable of providing sufficient electron current to the target

to neutralize the proton beam charge at any proton beam operating level. Based on observations of the electrons in the solar

plasma, these electrons should have energies of about 20 to ^0

eV. D . Beam Purity

A specification for the ion beam, purity must take into ac- count the nature of the impurity. For example, the molecular

hydrogen ion (Hg"^), which is the major impurity, has chemical

and physical effects that are quite similar to those produced

by protons. Therefore, the effect of the molecular ions should

scale approximately as does its concentration. A heavy ion,

such as OH"^, will have a sputter yield over 100 times that of

a proton of the same energy. Thus, in this case, a ifo OH^ con- centration will have an effect equal to the 99^ proton component. Another aspect of beam purity that must be specified is the presence of energetic neutrals formed by charge exchange. As the fast neutrals will have the same mass effects on the surface as ions, and because they are not registered as a current

and are therefore not included in the dose measurements, it is

essential that the number of charge exchange neutrals not ex- ceed 1% of the proton flux. This factor is an important elem.ent in the selection of the mass separator system.

The beam purity requirement should specify a limit for

energetic photons (Lyman a radiation) that are produced in the

ion source. These photons, which are due to the excitation of

hydrogen oy electron bombardment, have energies of 10 to 13 eV. They can cause a very marked effect on the degradation of thermal control coatings .( 11) The production of photons is negligible in the low pressure electron bombarcbnent ion sources and

pronounced in the plasma type ion sources. This factor is con- sidered in the selection of the mass separator.

265
E. System Requirements

The specification of the operating pressure of the system

must be set in terms of the partial pressures of the residual

gases that can be present in the system. This is important be- cause gases containing oxygen are known to cause bleaching of

some of the color centers formed in the degradation of the

thermal control coatings. The tolerance limits for these gases

must be set to ensure that they will not interfere with the ex- periments . Duplication of the vacuum conditions that exist in the interplanetary space (i.e., a random gas with a pressure of the

order of 10' -^3 Torr and a directed pressure of the order of

]_0-ll Torr due to streaming from the sum ^-2) is unrealistically

expensive o However, the results of solar wind simulation performed at 10'^ Torr are different from the effects of operating

the same coatings in space. Experiments have been performed(l3)

that show marked changes in the reflectivity of thermal control

coatings that are exposed to 10*5 Torr partial pressure of oxy- gen (after irradiation). This can also be shown by comparing

the arrival rate of oxygen at 1 x 10-7 Torr to that at 1 xlO"-^^ Torr. The former will accumulate on a surface to form a mono- layer in about 10 sec, while the later condition will require

3 years to form a monolayer.

The significance of this estimate is that it will be nec- essary to perform experiments on the irradiation of identical

samples under a range of vacuum conditions to determine if the

experiments are sensitive to the presence of specific gases. This is a standard procedure when operating an experiment in

the presence of additional factors that can influence the re- sults . Because of the uncertainties in the estimation of residual

gas effects on the experiments, the a priori establishment of

partial pressure limits is not possible. Instead, it is neces- sary to set up guidelines which will result in the design of a system that has the capability to operate at pressures below

the 5 X 10" 7 Torr level. This will permit the establishment of

these basic partial pressure parameters. In view of the fact

that hydrogen vail be the major gas present (due to the operation of the proton source), it is possible to set the total

operating pressure limit at 5 x 10"^ Torr and to set, as a design goal, the limit on the total of all other gases at about

1 X 10"^ Torr. In addition to controlling the sample temperature it is

necessary to duplicate the condition that exists in space where

none of the radiation from the surface and none of the gas desorbed from the surface is returned to the surface. The 3°K

266
black body condition of space must be simulated by cryogenic

shrouds treated to have a high absorptivity.

Another design requirement is that the simulator system

have a continuous operating lifetime that is greater than the test period. This could be of the order of 1000 hours. The reason is that the pressurization of the system to repair a component would cause the loss of the samples due to the bleaching of the irradiation damage. An alternative is to design a system in vhich the beam- forming components can be isolated

from the samples oy means of a high vacuumi valve. This permits

the pressurization of either element while maintaining the rest of the system under vacuum.

The specification of a proton beam area of 10 cm diameter

is based on two considerations. This represents about the

smallest area which can contain about 6 to 8 samples 2.5 cm diameter, which is the size nonnally used. This size oeam is an upper limit for an e>rpanded single proton beam. Thus, this

system makes possible a direct comparison of the continuous

beam and the intermittent rastered beam. This is important be- cause larger samples will require the use of a rastered beam to obtain uniform irradiation.

IV. DESIGN AMLYSIS OF A SOLAE WIKD SIMUIATOR SYSTEI4 This solar wind simulator system is based on the preceding

discussion of the solar wind characteristics, the solar wind

simulator requirements and on King'sv-^) analysis of the techniques and apparatus for proton beam formation. An outline

drawing of the proposed system is shown in Fig. 2. It consists

of an r-f excited plasma ion source, the ion optics system

(which includes the mass separator), and the sample chamber.

The discussion of the components is organized about these three

basic elements.

A. Ion Source

King's analysis of proton sources indicated that the r-f

ion source and the electron bombardment ion sources are the

types best suited for use in a simulator system. The r-f ion source was selected for this design because it is capable of

operating at the maximum ion current ( 1000 times the solar wind

ion density) without causing an excessive gas load in the system. The r-f source also has the advantages of a long operating life and a high proton production efficiency.

If the accelerated testing was limited to less than about

50 times the solar wind level, the electron bombardment ion

source could be used in the simulator system. This source is

267
readily adjustable over a wide range of ion currents, and pro- duces ions with a narrow energy spread.

B. Ion Optics System

The beam extraction system is designed so that only the source is at high voltage. The beam is kept small until it

reaches the last lens, to reduce aberrations in the lenses and

separator and to reduce the size of the components. This also

permits the beam to be rastered. The small size of the beam

also means that it is possible to separate the source chamber

from the main chamber with a small orifice without affecting

the beam. Thus, the two chambers may be pumped separately and

the large gas load from, the source removed without interacting

with the target. This system is shown i:i Fig. 3' The ion extraction system or ion accelerator focuses the

ions from the source into a beam with a divergence angle com- patible with the rest of the ion optical system. It is operated in the accel/decel mode to prevent electron backstreaming

and to permit adjustment of the beam voltage and thus the pro- ton energy without affecting the field gradient at the surface

of the plasma in the source from which the ions are extracted.

The single small aperture provides adequate current while still

acting as an effective flow impedance to reduce neutral hydrogen efflux from the discharge chamber.

The first or focusing einzel lens focuses the ion beam

through the mass separator onto the aperture h cm from the exit

plane of the magnetic sector. By adjusting the beam size at this aperture, the intensity of the beam which arrives at the

target may be adjusted by a factor of 100.

The proton component is separated from the ion beam, which

also contains molecular ions, by means of a 20° sector magnetic

m.ass separator. King(l) has shown that this type of mass sep- arator is superior to the E x B, the Bennett r-f field, and the

magnetic lens types. It fulfills the two basic requirements - separation of the proton beam from both the charged and un- charged (i.e., photons, charge exchange neutrals) contaminants

from the source without seriuosly defocusing the beam. King

has discussed the tradeoff, between adequate separation of

charged species and beam defocusing due to energy dispersion,

that must be made in choosing the separator angle. The 20° sector separates the protons from the beam by 0.6 cm at the

aperture stop located h cm downstream of the exit plane of the

separator. The increase in diameter of the 1 mm diameter,

1000 V oeam due to a 50 eV energy spread when passing through

this 20° separator is less than 0.5 mm. Thus, a 2 mm aperture

will pass all of the proton beam and block other charged spe- cies. This separator will nominally operate at a magnetic

268
field of 300 gauss and consume less than 200 W. It should not require water cooling. The final design of such a separator

that is to be actually constructed should involve careful con- sideration of the shape of the magnet pole entrance and exit

surfaces and of the magnetic gap itself, in order to minimize

aberrations.

This 2 mm diameter aperture also serves to limit the flow

of neutral gas molecules into the chamber. It is thus possible

to differentially pump the gas that is released by the ion

source and thus maintain the sample chs^mber at a better vacuum. The second einzel lens controls the size of the beam at

the target. It is capable of expanding the beam that passes

through the aperture to 10 cm diameter at the target plane or maintaining the beam size at 0.1 cm so that it may be rastered.

As shown in the insert in Fig. 2, a set of deflection

plates may be inserted between the second lens and the target

to raster the beam.

The above four elements will provide a 10 cm diameter beam

at the target in which no proton trajectory has an angle of more than 3*^ with respect to the normal from the target plane. ^

C . Sample Chamber

The proton beam then passes through an all-metal high

vacuum valve that serves to isolate the proton beam forming

part of the simulator system from the sample chamber. This

serves two functions:

1. to permit the samples to remain under vacuum in the event it is necessary to pressurize the ion source

section,

2. to pennit the ion source region to remain under vacuum

and thus protected from exposure to water vapor during

the loading of samples.

The sample chamber shown in Fig. k contains the neutralizer;

the cryogenic shroud; the sample transfer arms; and the sample

mounting plate (which supports the samples); and the proton,

electron, and photon detection and measurement apparatus. The

neutralizer consists of an electron emitter 9perating in an accel-decel mode that produces up to 3 x lO""-" A of 30 eV electrons. The cryogenic shroud is liquid nitrogen- cooled and is

treated so as to have an absorptivity of 0.95 or larger.

The sample mounting plate is a constant temperature plate

to which the samples are mechanically attached. Fluid is cir269
culated from an external reservoir to maintain the plate tem- perature. The plate is moved in the x-y plane by means of

bellovs sealed rotary motion drives in order to make possible

sample transfer and to scan the proton and photon beams.

The sample transfer operates in the following manner. The

plate is positioned above the z-axis transfer arm (see Fig. h)

The arm is extended to engage the sample holder^ and the rotary

motion then is used to unlock the sample. The sample is car- ried upvard by moving the z-axis transfer arm, and the sample

is tranferred to the y-axis sample transport arm. The y-axis

transfer arm moves the sample into the integrating sphere.

D. Beam Measurement Apparatus

The sample mounting plate also supports the apparatus for

determining the flux, energy, and profile of the proton and

electron oeams and the flux and profile of the photon beam.

These detectors, as well as some pressure sensors, are part of

the control loop that regulates the operation of the ion source and the neutralizer.

The determination of the flux, energy, and density profile

of the proton beam is accomplished by use of a Faraday collector located below the sample mounting plate. The collector

should have a length-to-diameter ratio of about ^^i- to 1 to en- sure that the potential within the collector is very small and

thus prevents the loss of secondary electrons or ions. Shielding and operation of the collector at a small positive bias may

be necessary to prevent errors due to photoelectron currents that can result from the solar photon radiation simulation.

The proton energy is determined by measuring the beam po- tential as a function of bias potential on the collector. The

proton beam energy spread is determined by plotting the first

derivative of the beam current as a function of proton beam

energy. The proton beam density profile is determined by mea- suring the beam flux (at a fixed set of operating conditions)

at a number of points by moving the mounting plate in the x-y

plane. Simultaneous measurement of the proton and electron

beams is needed to regulate the operation of both sources, to provide a measurement of the proton dose rate during the run, and to verify the degree of neutralization. This can be accom- plished by passing the combined beam (protons and electrons)

through an aligned electric and magnetic field. The electrons

are swept out by fields to another collector, thus giving an electron and an ion current. Some of the experimental problems that should be anticipated in the design of the measurement apparatus include both

extremely low signal levels and the high noise levels that can

270
result from the r-f generator used in the ion source, the plas- ma in the ion source, stray electron currents in the chamber,

and photoelectrons . While there is little that can be done

about the low signal levels (e.g., signal of the order of

10"-^^ a), there are a number of remedies for the high noise

level. The r-f and plasma noise can be reduced oy the use of

TT or T section filters or by the use of ferrite beads. The

design of the filters can be obtained from any standard text. Carefully designed shields are the most effective way of preventing the stray electron currents from measurement as part of

the signal. Photoelectric currents can be suppressed by operating the collector at a slightly positive bias with respect to the surroundings

The ion-electron converter^ ^ ( -^5) can be used for visual

check of the beam shape and density variations. This consists

of a fine metal mesh mounted near an electron excited phosphor

on a plate. The ions strike the mesh, generating secondary

electrons that are attracted to the phosphor, which is held at

a high positive potential. Use of an electroformed nickel mesh

of hOO wires cm~^ gives resolution on the order of 25 to 50 fj,m.

E. Vacuum Pumps

The vacuum puinps shoi<m in Fig. 2 consist of two pairs of

sputter ion and titanium sublimation pumps. The combination is

necessary because the pumping mechanism for hydrogen in a sput- ter ion pump is limited^to chemical reaction (gettering) with

the cathode material. (16),(1?) The titanium sublimation pump is

added to increase the pumping of the hydrogen (as well as any

other chemically reactive gases). The sputter pump must handle

the gases that are not gettered by the titanium.

The use of liquid helium cooled cryogenic pumps is nou advisable for a system in which the predominant gas load is hydrogen. The reason is that the vapor pressure of hydrogen is of

the order of 10"^ Torr at a cryopump temperature of the order

of k,^ to 4.2°K.(-^^) This value is poorer than that which can

be obtained from the TSP and ion pump.

V. SUMMARY

The design that has been produced satisfies the requirements for simulation of the solar wind. The most significant

features are:

271
a. the magnetic mass separator to purify the beam

of other ions^ energetic charge exchange neutrals,

and Lyman alpha photons,

b. the small ion beam that makes possible differential piimping of the system,

c. the direct comparison of the expanded versus

the scanned beam, and

d. the isolation valve between the beam forming

system and the sample chamber.

VI. REFERENCES

1. H. J. King, "Solar Wind Simulator, to be given at Space

Simulation Conference, 14-15 September 1970, Gaithersburg,

Maryland.

2. A. G. Stanley and J. L. Ryan, "Charged Particle Radiation

Environment in Synchronous Orbit, " Technical Report hh^,

Contract AF 19 (628)-vl'7, MIT Lincoln Laboratory,

15 May 19 68.

3. D. E. Robbins, A. T. Hundhausen and S. J. Bame,

J. Geophys Res. 75 , 1178, (1970).

h. J. Hirshberg, A. Alksne, D. S. Colburn, S. J. Bame, and

A. J. Hundhausen, J. Geophys Res. 7^ > 1 (1970).

5. S. J. Bame, A. J. Hundhausen, J. R. Asb ridge, and I. B. Strong, Phys. Rev. Letters 20 , 393 (19^8).

6. A. J. Hundhausen, Space Sci. Rev. 8 , 69O (I968).

7. M. Neugebauer and C. W. Snyder, J. Geophys. Res. 71 > i+469 (1966) . s- ,;

, 8. I. B. Strong, J. R. Asbridge, S. J. Bame, and A. J. Hundhauser, in "Zodiacal Light and Interplanetary

Medium," J. Winberg, Ed., NASA SP-I50, I967.

9. M. D. Montgomery, A. J. Bame, and A. J. Hundhausen,

J. Geophys. Res. 73 , ^999, (19^8).

10. F. Grjz^nlund and W. J. Moore, J. Chem. Phys. _31 , 1132

(19 ^^9).

272
11. D. D. Swofford, V. L. Mangold, and S. W. Johnson,

Thermal Design Principles of Spacecraft and Entry Bodies , J. T. Bevans, Ed. (Academic Press, New York, 1956), p. 667.

12. L. A. Nelson, Handbook of Solar Simulation for Therma l Vacuum Testing , J. S. Griffith, Ed. (Institute for

Environmental Sciences, I968) . p. 3-1'

13. J. E. Gilligan and G. A. Zerlant, AIAA Paper No. 69-IO25,

AIAA/ASTM/IES 4th Space Simulation Conf., Los Angeles,

8-10, September I969.

Ik. G. Kuskevics, J. Appl. Phys . 39 , ^076 (I968).

15. D. M. Jamba and 0. Husmann, J. Appl. Phys. _38 , 263O

(1967).

16. S. L. Rutherford and R. L. Jepsen, Rev. Sci. Instr. 32 , llkk (1961).

17. S. L. Rutherford, S. L. Mercer, and R. L. Jepsen,

1960 Seventh National Symposium on Vacuum (Pergamon

Press, New York, I96I}

16. E. S. Borovik, S. F. Grishin, and Ya. Grishina, Soviet

Phys. Tech. Phys. 5 , 505 (i960).

273
eV

HRL2I0-42RI 52 208 468 832 1300 1872 2550 VELA 2A AND 2B DATA COMBINED

AREA= 100 CASES _

0 100 200 300 400 500 600 700 800

SOLAR -WIND VELOCITY, km/sec

Fig. 1. Distribution of solar-wind

ion energy and velocity from

July 1964 to July 1965

(from ref . 8 )

274
Fig. 2. Solar wind simulator system.

275
Fig. 3. Layout of solar wind simulator.

276
277

Paper No. 19

SOLAR WIND SIMULATOR^

H. J. King

REFERENCE: King, H. J. , "Solar Wind Simulator, " ASTM/

IES/AIAA Space Simulation Conference, 14- 16 September 1970.

ABSTRACT: It is necessary to expose samples of thermal

control coating materials to a proton flux in the laboratory in

order to estimate the stability of their reflective and absorptive properties when exposed to the solar wind. The choice of the ion source, mass separator, and beam transport system

which make up an experimental apparatus to perform these

tests is discussed. A final system design is presented which

is capable of irradiating 10 standard samples 2 cm in diameter

with a proton beam ranging in energy from 0.5 to 3.0 KeV and

in intensity from 1 to 1000 times that of the nominal solar wind,

KEY WORDS: solar wind simulator, proton source, mass

separator, ion optics, ion beam deflection

INTRODUCTION

Scientific probes traveling through space encounter several

types of particle fluxes. A primary class are those emanating

the sun—the so-called solar wind. A companion paper (1)-^ describes these particles, their intensity and their relation to the activity cycle on the sun. This paper deals with the design

of apparatus to simulate the solar wind in laboratory facilities.

^Head, Propulsion Technology Section, Hughes Aircraft

Research Laboratories, Malibu, California..

This paper is based on NASA CR 73443 prepared for the Ames

Research Center under contract NAS 2-5585. Please see the

final report under this contract for a quantitative discussion

of the material presented here.

^The numbers in parentheses refer to the list of references

appended to this paper.

279
The solar wind simulator discussed herein is intended to be

used for evaluating the synergistic effects of the combined

particulate and photon radiation on thermal control coatings.

Zuccaro (1) has discussed the basic requirements for apparatus

to reliably evaluate these materials and the techniques to perform the experiments. This paper presents the design rationale

behind the selection of the major components of the simulator.

SYSTEM DESIGN

The design of the apparatus such as shown in Fig. 1 requires the selection and design of the ion sources, the mass

separator, and the beam transport system. The system must

be capable of irradiating 10 standard samples 2 cm in diameter

with a proton beam ranging in intensity from 1 to 10 times

that of the nominal solar wind. The nominal sola^ wind is defined as a p^re proton flux of 2 x 10^ protons/cm —sec (3.2 x

lO'-^^A/cm^) ranging in energy from 0. 5 to 3. 0 KeV with a nominal mean energy of 1.0 KeV.

A. ION SOURCES

1 . Physical Characteristics

The characteristics of the ion source affect all the other

major components of the system — the ion optics, the mass

separator, and the vacuum station itself. Fortunately, for

this study, proton sources have played an important role in

nuclear physics for many years, so that a large body of information is available to define and compare the various possible source configurations. Before considering individual ion sources in detail, it is appropriate to interpret the previous

section in terms of the conditions which it imposes on the ion source. , Considering the nominal intensity of the solar wind, the

desire to accelerate the test time by a factor of 1000, the

target size and attenuation in the apparatus, the ion source must generate a proton current of approximately 10" A. The

bulk of the protons in the solar wind range in energy from 0. 5 to 3. 0 kV. In order to provide the necessary experimental

flexibility, the apparatus must be designed so that the proton

energy is variable, at least over this range. While it is well

within the capability of the designer to make the instantaneous

ion energy a variable, this should not be confused witVi the

virtually impossible task of providing an ion beam which contains a controllable mixture of all ion energies at a given time.

Fortunately, the latter is not required for an accurate simulation of solar wind.

280
A related characteristic is the energy spectrum of the ions

inside the ion source. This energy spread appears as an uncertainty in the final proton energy at the target. While this

may not be detrimental to the experiment itself, it does impose

severe limitations on the design of the ion beam transport and

separation systems. Therefore, an ion source with a small

energy spread is desirable.

Mass efficiency n^^ is used to define the percentage of hydrogen atoms that actually leave the source as protons. A

high mass efficiency is desirable to reduce the neutral density

in the region immediately downstream of the source. This

reduces the creation of charge exchange ions which contaminate

the beam. (See Ref. 1) The useful lifetime of the source is important for an apparatus in which experimental simulation may be conducted in

real time. The arbitrary, but realistic, lifetime goal set for

this design was 1000 hours ('-'6 weeks).

The above characteristics provide a scale against which

the various sources may be compared. Where possible this

comparison is made quantitatively in Table 1.

2. Types of Proton Sources

There are alarge number of generic types of ion sources,

each with many individual designs, which will produce a proton

beam. An excellent review of the basic types was presented by

Hoyaux and Dujardin (2). All of these sources dissociate and

ionize the hydrogen by electron bombardment. They differ

radically in power requirements and general operating conditions. Of particular interest for this application are two

requirements mentioned earlier; namely, a minimum kinetic

ion energy spread in the source and the ability to vary the source output current intensity over a wide range to accommodate the required flux variations.

a. Radio Frequency Ion Sources (See Ref, 3)

The radio frequency ion source, sometimes called an

electrodeless discharge source, has been a standard for many

years in high voltage particle accelerators used for nuclear

physics experiments. RF power is coupled into a quartz (or

other insulating material) tube containing hydrogen gas at a pressure ranging from 10 to 100 ^m of mercury. The mean

free path of an electron in the source is on the order of 1 cm.

Thus, electrons will be accelerated by the imposed electromagnetic field and will, under the proper conditions, generate

a stable plasma from which an ion beam may be extracted.

281
The most common method of coupling rf power into the

discharge is to place the discharge chamber inside a solenoid

which is activated with rf power at approximately 10 MHz.

The alternating magnetic field generated in the discharge

chamber accelerates the electrons. Generally, the induced

electric fields are relatively small here and the circulating

currents rather large. Energy spreads ranging upward from

20 eV have typically been reported (4, 5) for this device.

Quite frequently an axial magnetic field is used to increase

the electron path length and to improve stability.

Beam currents up to 1 mA are typically reported for this

type of source. Proton fractions up to 90% of the total ion

current in the beam have been observed under optimum conditions; 50 to 70% is common for well designed sources. Hydrogen consumption varies according to source design but is typically of the order of 4 cm /hour at STP for 1 mA of output

beam. It may not be possible to scale down the gas flow

directly with beam current. Operating frequency varies as

well, ranging from 1 to 400 MHz; 10 MHz is common and rf power requirements are a few hundred watts. Source lifetimes

of 1000 hours at the current levels required for this application are reported

Of the various sources surveyed, the electromagnetic rf ion source most nearly meets all the requirements. It fully

meets the beam requirements and is adequate in terms of mass

efficiency and proton yield. If this type of source is to be

designed or adapted to this application, care should be taken to assure stable operation over long periods by choosing high

quality materials for its construction and designing both the

power supplies and the hydrogen feed system with stability as

a primary consideration,

b. Electron Bombardment Sources

As the name implies, electron bombardment sources

ionize by direct impact of electrons on the hydrogen molecule

or atom. The distinction here is that in contrast toother

types of sources conditions are such that no plasma is created

nor is any arc or discharge struck. While several designs are

possible only the low voltage source is of interest here.

Typical of the type most useful for the purpose at hand is that

described by Redhead ' '. Hydrogen gas is introduced into the

ionization region where it is bombarded by electrons with the

resulting production of protons. An axial magnetic field

significantly increases the electron path length and hence the source efficiency. The electrons may be axially restrained by

holding both ends of the discharge chamber at cathode potential,

thus making it energetically impossible for electrons to reach

them after any energy loss. Redhead has reported that this

type of source is capable of trapping ions in the source for

282
periods as long as 1 sec. This is desirable in a proton source because when a proton collides with the metal wall of the source,

it is re- emitted as part of a hydrogen molecule and must be

both dissociated and ionized again to be useful.

This type of ion source has several advantages. Primary

among these is that it fulfills the two desirable criterial mentioned earlier. It is adjustable in current over several orders

of magnitude and produces ions with a relatively small energy

spread, ranging down to less than 1 eV for specially designed

sources' The primary difficulty with this type of ion source

is that the mass efficiency is of the order of 1 % or less. Thus,

it presents a large gas load to the system when operating at high output currents. A second, and less important, disadvantage is that the source requires a thermionic cathode, which is subject to contamination and has a finite lifetime.

A sirnple source of this type is in use at Hughes Aircraft

Company' It is capable of producing a proton current

equivalent to 2000 times the nominal solar wind value over a

2. 5 cm diameter circle. Input power is about 20 W. The

entire vacuum station, including the source chamber and the

integrating sphere, is pumped by a single 400 liter/sec ion

pump. A single feedback loop that monitors the emitted ion

beam current and adjusts the cathode emission to compensate - for drift in line voltage or hydrogen flowrate serves to stabilize the beam for periods of several days.

c. Other Proton Sources

All other ion sources investigated were considered unsuitable for this application for a variety of reasons. These types

included the high voltage and crossed field electron bombardment sources mentioned above, gas discharge sources such as

the Penning or Finkelstein source, the duoplasmatron and the

capillary arc. The reasons for this selection have been discussed in detail in the final report from which this paper was

taken and are summarized in Table 1.

3. Ion Extraction Systems

The foregoing sections have discussed a number of proton

sources. To be useful, these protons must be extracted from

the source and focused into a beam. The ion extraction system

which accomplishes this forms a critical interface between the source and the ion transport system which controls the beam

shape en route to the target.

The theory of charged particle focusing has been investigated extensively in the development of electron guns used for

many purposes. Because the focal properties of electrostatic

283
lenses are independent of particle mass, the theories derived

for electrons are equally valid for protons. The principal

difference is that the currents that may be carried with a given

set of lenses and accelerating voltages are much less for protons than for electrons because the particle velocities for a given accelerating voltage vary inversely as the square root of the mass

The simple, but important, concept of operating the extraction system in an accel : decel mode is often overlooked by

those designing ion transport systems. The ions are extracted

with a total voltage greater than the desired beam voltage and

then decelerated. This accomplishes two things. First, it prevents electrons from the beam or target from entering the source and damaging components or disturbing the control

system which often operates on total emitted current. Second,

it separates the extraction voltage from the beam voltage and

permits a higher beam current to be extracted than would be

possible at the beam voltage alone. It also permits adjustment

of the beam voltage without disturbing the position of the plasma

sheath because the total voltage may be held constant as the ion

beam voltage is varied.

. B. MASS SEPARATORS

All of the ion sources discussed abo]^e produce an ion beam

output which is a mixture of protons, H2 , H-j''", and many other

ions. The sources may also produce a high flux of fast neutral

particles by charge exchange and a number of ultraviolet

photons which can degrade the test sample. The photons can

create photo electrons if they are allowed to strike the target or the Faraday cup which monitors the beam. The mass separator, which is located between the proton source and the

target, must purify the beam which passes through it so that

only the protons will strike the target.

The physical separation of the protons from the rest of the beam can be accomplished in two basically different manners. An appreciation of the two concepts plays an important role in choosing the most desirable type of mass separator.

All ion sources produce a mixture of protons, other positive

ions, fast and slow neutral particles, photons from the discharge within the source and sputtered material from the ion

extraction system or focusing lenses. The first method is to design a separator which selects only the protons from this

beam. The trajectories of the protons are controlled so that

they alone reach the target, which is shielded from the source

so that none of the contaminants may reach it by line of sight.

The magnetic sector magnet is an example of such a device.

The second system concept is to provide a series of filters to prevent the unwanted particles from reaching the target while

284
permitting the desirable proton beam to pass. The difficulty

here is first to identify all of the contaminants which may exist

in the beam and then to devise filters ^-J; ich will selectively

stop the contaminants but transmit ^ .j.c protons. This is particularly difficult in the case of photons and fast neutrals. The rf separator is an example of this type of system. As will be

shown, a separator which deflects the proton beam is clearly

superior for this application.

1 . Physical Characteristic s The various separators were compared on the basis of their ability to produce a pure proton beam; transmit a large

fraction of the incident photons; provide a continuous rather

than pulsed beam; introduce minimum aberrations into the ion

beam transport system; provide continuous, stable operation;

accommodate a full range of ion velocities; and their ease of design and use.

2. Types of Mass Separators

Three general types of mass separators were considered:

Radio Frequency - which utilizes resonance between particle inertia and an alternating electromagnetic field. Homogeneous Magnetic Field - in which the ions of differ- ent mass are spatially separated due to their different

radius of curvature in a magnetic field. Crossed Field or ExB - where perpendicular static

electric and magnetic fields cause different trajectories

for different masses.

a. Radio Frequency

The basic concept of a UAear rf mass spectrometer was

first described by Bennett. ^ ' The operation had been analyzed by a number of authors, particularly Redhead and

C r owell , who list all the critical performance parameters

for both sinusoidal and square wave operation.

The experimental performance for a system of this design

has been reported recently (Ref. 13). This device has 13 grids,

three rf stages, and two drift spaces. The complete system is 25 cm long and operates at 10 MHz. Assuming that each grid

is 95% transparent but that the individual wires are randomly

oriented with respect to those of the other grids, approximately

50% of the particles from the source will impinge on the grids.

While the beam may be purified by this device to greater than

285
95% , it can be accomplished only at the expense of reducing

the transmission to less than 10% . This in itself may be

quite tolerable if a source with a sufficiently high output current

is used. The difficulty arises because the photon output from

the source and the beam of fast charge exchange neutrals

created at the source aperture are not attenuated by the electric

fields. Thus they are increased in relation to the proton beam

by at least an order of magnitude using this type of separator.

The effects of these contaminants may thus easily be greater

than those of the proton beam for many operating conditions.

A further complication is that the impurity beam may not be

uniform over the target area and may be very difficult to measure accurately.

Therefore, even though this system has many advantages

in terms of minimum energy dispersion, astigmatism, cost,

and size, it is suggested that serious consideration be given

to the problem of beam contamination before it can be used for

any simulation.

b. Homogeneous Magnetic Field

This is a "conventional" type of mass separator which

operates on the principle that the trajectory of a charged

particle will be deflected by a magnetic field. A great deal of developmental work devoted to understanding and improving

this type of mass spectrometer has been reviewed by Inghram

and Hayden (14). Of Primary importance to the overall system performance

are two parameters of the separator design. The first is the

angle through which the beam of interest (protons) is deflected.

It is directly related to the strength of the magnetic field and

the path length in the field. The second factor is the dispersion

of the various masses at the target. In order to preserve the

minimum beam size, the dispersion through the separator

should be minimized. This may be done by establishing the

minimum sector angle for the separator which will spatially

separate the proton beam from the H2 beam.

The magnetic sector mass separator is well suited for

this task, provided that the beam is kept small in diameter

until after separation and then expanded to cover the target area. The power requirements increase rapidly as the beam

size increases, because of the resultant increase in magnetic

gap. The magnetic fringing fields also become very difficult to adjust at large gap spacings, thus introducing unwanted aberrations in the final beam. This type of separator accomplishes

the two necessary functions of bending the proton beam to physically remove it from the efflux of photons and neutrals

from the ion source (which travel line of sight), and of dispersing the charge particles of various masses so that a pure

286
proton beam results. Care must be taken in the design to reduce the ion optical aberrations and to minimuze the dispersion due to the energy distribution in the ion source.

c. Crossed Field or ExB

In a crossed field separator, the magnetic field B, the

electric field E, and the velocity of a charged particle are all mutually perpendicular. If the signs are properly chosen, it is clear that by adjusting the magnitudes of the electric and

magnetic field strengths the electric and magnetic forces on

the particle may be made equal and opposite and the particle

trajectory will be unchanged. Particles with the same kinetic

energies but a different mass will follow a curved trajectory.

The properties of this type of separator have been discussed

by Bleakney and Hippie \^^) . The simplified analysis of Wahlin (16) is recommended as clearly illustrating the important processes. This type of separator will serve equally as

well as the simple magnetic sector in many cases. The extra

variable of the electric field strength provides the opportunity

to design the system to achieve first order focusing.

3. Selection of the Mass Separator

The properties of the three types of mass separators discussed above are summarized in Table 2. Of the three, only

the rf separator is clearly rejected because it does not prevent

uncharged particles and photons from striking the target. The

remaining two —the sector and crossed field separators —are

similar in many respects. The crossed field separator contains aberrations from nonuniformities in both the electric and

magnetic fields, while only a magnetic field is present in the sector. In particular, it is difficult to generate a uniform

electric field when the magnetic and electric pole pieces of the

crossed field separator must of necessity be in close proximity.

This problem has recently been solved analytically and appears

manageable for low energy applications such as that required

for the present program.

The sector magnet was chosen over the crossed field device

for the system design here, principally because it requires one

less power supply and is a well understood, proven device.

Techniques have been derived to produce very sophisticated

instruments of this type, if they are required. Special attention should be devoted to the defocusing of the beam due to energy dispersion in the source. Careful design of both source

and separator can keep this well within tolerable limits.

287
C. BEAM TRANSPORT SYSTEM

,. 1 . Introduction

The beam transport system must accept the ion beam from

the source, focus it through the mass separator, and expand it to 10 cm diameter at the target plane. It must also be capable

of providing adjustment of beam intensity by a factor of 1000.

Finally, it must perform satisfactorily over the full range of proton energies from 0.5 to 3.0 kV. Two general concepts are

possible. The first is to extract from the source a small,

relatively high current density beam which is expanded in size

only after it has passed through the mass separator. This

considerably simplifies the design of the einzel lens, because

all the beam ions pass through the lens near the axis where

aberrations are small and the paraxial ray equation is valid.

The mass separator is simplified as well because the magnet

gap widths can be minimized. The single small aperture in the source serves to minimize the effluent hydrogen gas and the

pumping speed necessary to maintain satisfactory chamber

pressure. Finally, the small beam permits a series of apertures to be installed between the source and the main vacuum

chamber, to reduce the diffusion of hydrogen from the source

into the chamber, A separate pump is provided near the source. The second approach is to extract a large beam from the source and transport this broad beam through the mass separator to the target. This virtually eliminates the need for an ion

optical system, but entails a number of other difficulties which

are obvious from the preceding paragraph. The only mass

separator which readily accommodates a broad beam is the rf separator, although large aperture crossed field and even magnetic sector separators are possible at the low beam

energies used here. Only the first type of system (i. e. , small

beam) will be treated here.

Z. Computer Simulation of Ion Trajectories

As the name implies, each element in the ion optical

system may be treated mathematically by techniques used in

conventional geometrical optics. For instance, the einzel

lenses have an equivalent thick lens analogy, and a sector magnet is equivalent to a cylindrical lens bounded on the entrance

and exit sides by thin lenses. The focal properties of the

lenses are a complicated function of the geometry, beam voltage, and lens voltage for each.

It has been shown (17) that where the paraxial ray equation

is applicable, this problem may be readily treated by a matrix

technique which is easily adaptable to computer calculation.

Assuming circular symmetry, the trajectory of a particle may

288
be defined at any time by its radial distance and radial momentum and represented by a column vector. The effect of each ion

optical element on the trajectory is represented by a two-bytwo matrix so that the total transfer matrix for the entire

system is thus a product of n matrices, where n is the number

of ion optical elements in the system. It is clear that this

formulation is suitable for computer calculation.

Only electrostatic lenses are considered here because

they consume no power and are independent of the mass of the

particle being focused. The most versatile type is the einzel

lens which is composed of three apertures, or cylinders. The

two end elements are connected together electrically so that

the beam leaves the lens with the same kinetic energy with

which it entered. The central element is biased to retard the

beam (i. e. , biased positively for positive ions). The focal

length of the lens is a direct function of the ratio of the voltage on this central element to the beam voltage. While such lenses

are simple in concept, they introduce sizable aberrations if the

beam is not kept small relative to the lens aperture.

Three einzel lenses are used in the beam transport system

for focusing purposes. The first, in conjunction with the weak

focusing properties of the mass separator, focuses the beam

through the aperture following the mass separator. This aperture also provides the pressure differential needed between the

target and source chambers. The second lens produces a magnified image of the aperture, resulting in a 10 cm diameter

image at the target. The third lens (optional) provides the

slight collimation needed to produce an exactly parallel beam

at the target. Without this last einzel lens the long "throw" of the second einzel produces a landing angle of tan"-'- (5/100) —

2. 7 deg.

Because of the large number of elements in the system,

the operation of the beam transport system is too complex to describe quantitatively here. Interested readers are referred

to the final report on NASA Contract NAS 2-5585 (10) for a detailed description.

The total system must be able to change the target current

by three orders of magnitude. Because it is not possible to operate the rf source over such a wide dynamic range, it is necessary that the beam transport system provide approximately

a factor of 100 intensity variation, which together with a variation of source current by a factor of 10 provides the total dynamic range required. The source current will be varied by

changing the neutral flow and/or the rf excitation power.

The variation provided by the beam transport system is accomplished by flooding an aperture and allowing only a small fraction of the current to pass through the aperture by changing the

289
first einzel focusing voltage. This means the beam size at the

aperture may be varied from a radius of less than 0. 1 cm to greater than 1. 1 cm by the first einzel. This is about the maximum intensity ratio possible with this method because the image

size at the target is limited by the maximum magnification

(minimum focal length) of the lens.

3. ^eam Rastering System

The system illustrated in Fig. 2 is based on the requirement that the simulation requires a large dc proton beam which

covers all of the target area uniformly. It has been estimated

(10) that the simulation would not be impaired if the protons

arrived at the target as a series of discrete pulses at rf frequencies. If it can be demonstrated experimentally that this is indeed true, a system employing a small diameter proton beam

which is swept over the target surface is a pattern similar to the electron beam in a TV raster has many advantages. The

system shown in Fig. 2 is easily converted to this design by

changing the operating voltage of diverging einzel lens No. 2 to focus the beam to a spot on the target and introducing two

pairs of orthogonal electrostatic deflecting plates in place of optional collimating einzel lens No. 3. The advantages of a raster ed system are

1. By sweeping a small beam over the large target area,

the beam nonuniformities are smeared out and the

coverage may be made as uniform as desired, even if some lens aberrations exist,

2. By programming the rate at which the beam is swept,

the dose rate to each individual sample may be controlled (as a function of exposure time if desired).

3. The dose rate may be attenuated over several orders

of magnitude while the source operates at the single

fixed point where it performs best. This is achieved

by deflecting the small beam off the target area

entirely for some fraction of each raster period. This

not only extends the operating range of the system, but

simplifies the design of the ion source and its control

system.

4. The accuracy of the dose measurement is improved by

this technique because the total proton beam (which may

be measured accurately) is projected onto each area of the target for a time, which also can be measured

accurately. The result is a far more accurate measure

of dose than sampling a small fraction of the beam and

relying on a prior measurement of beam homogeneity.

290
It is also much easier to design a dynamic control

system for this system than for one that samples^a

fraction (~ 1% ) of the beam current (i. e. , ~ 10~ A).

D. CONCLUSIONS

The discussion in the preceding section illustrates that a system of the general design shown in Fig. 1 can indeed be

designed and that the requirements on each of the various components required is within the current state of the art. The

most critical area requiring experimental verification and

possibly further development is the design of the ion extraction

system which draws protons from the source and focuses them

into a parallel, laminar flow beam. This interface between

proton source and the rest of the system is the key to successfully producing a beam of uniform current density at the target.

Two important design principles were formulated as a result of this study. The first is that the proton beam should

be bent during its traversal of the system so that it is possible

to completely shield the target from photons, fast neutrals and

possibly sputtered material that are inadvertently produced by

the source. The second design philosophy is that if it does

not reduce the experimental accuracy it is more desirable from

a system designer's viewpoint to irradiate the target with a small beam which is rastered over the surface than with a broad, uniform dilute beam.

References

1. Zuccaro, D. E. , companion paper this conference.

2. Hoyaux, M. and Dujardin, L , Nucleonics, 4, 7; 4, 12;

5, 67 (1949).

3. Blanc, D. and Deghilh, A., J. Phys. Radium^, 230 (1961).

4. Hall, R. N. , Rev. Sci. Instr. J^, 905 (1948).

5. Thoneman, P. C. , Moffatt, J. , Roaf, D. , and Sanders,

J. H. , Proc. Roy. Soc. 6J_, 483 (1948).

6. ORTEC Model 320 RF Ion Source Data Sheet.

7. Redhead, P. A., Can. J. Phys. 45, 1790 (1967).

8. Dworetsky, S. Novick, R. Smith, W. W. and Tolk, N. , Rev. Sci. Instr. 39, 1721 (1968).

9. Constructed by H. J. King and D. E. Schnelker; unpublished.

10. Solar Wind Simulation Techniques Final Report Contract

NAS 2-5585 HRL Staff (1970).

291
11. Bennett, W. H. , J. Appl. Phys. 2J., 143 (1950).

12. Redhead, P. A. and Crowell, C. R. , J. Appl. Phys. 24,

331 (1953).

13. Lebduska, R. and Meckel, B. , "Solar Wind Simulator and

Advanced Design with Extended Operational Capabilities, " AIAA Paper No. 70-31, January 1970.

14. Inghram, M. G. and Hayden, R. D. , Handbook on Mass

Spectroscopy (Nuclear Science Series, National Academy

of Science, Washington, D.C., 1954), No. 14. 15. Bleakney, W. and Hippie, J., Phys. Rev. 53_, 521 (1938). 16. Wahlin, L. , Nucl. Instr. Methods 27_, 55 (1967).

17. Cotte, M. , Ann. Phys. (Paris) ^O, 333 (1938).

292
__

,

Disadvantages

i

Narrow

beam

current

range

Wide

energy

spread

Low

mass

utilization

High

power

and

large

energy

spread Low

beam

current

Short

life,

complicated

electronics,

hard

to

stabilize

Current

too

high

Trade

rij-j^g^gg

for

proton

yeild,

cathode

life,

or

energy

spread

Advantages

Long

life

and

high

percent

of

protons

Low

operating

pressure

and

beam

current

Wide

range

of

beam

current

Low

energy

spread

and

wide

range

of

beam

current

Monoenergetic

High

mass

utilization

and

high

proton

yield

o >^

+->

•i-H

d)u

u

3tromagne1 -trostatic

^

Voltage 1

Voltage

lal

Ray)

ssed

Beam

illary

Arc

High

Dens

Low

Dens

:i0 Ele( Ele< ^0 Cro Cap (3^3) pi9i^

passoj^

Type

o

Radio

Frequency

1

Electron

Bombardment

Gas

Discharge

293
rals

ision

Disadvantages

.

«

100 %

purity

.

Low

transmission

of

protons

.

Transmits

fast

charge

exchange

neut

.

Must

trade

beam

purity

for

transmis

.

Photons

from

source

strike

target

.

High

power

for

large

beam

.

Energy

dispersion

.

Two-power

supplies

.

Electric

and

magnetic

aberrations

Advantages

.

Handles

large

beam

.

Small

size

.

Low

total

power

.

Easy

intensity

adjustment

.

Pure

beam

.

High

transmission

.

Compatible

with

scanned

beam

.

Pure

beam

.

Will

take

large

beam

Type

Radio

Frequency

Homogeneous

Magnetic

Field

Crossed

Field

or

ExB

294
Figure 1- -Solar Wind Simulator

295
TARGET

Proton ^1 ^2 ^3 ^4 ^6 ^7 Energy

One Equivalent Sun

500 V 500 -1250 0 400 465 375 0 1000 V 1000 - 750 0 800 930 750 0 3000 V 3000 - 500 0 2 400 2825 2250 0

100 Equivalent Suns^

500 V 500 -1250 0 310 460 375 0 1000 V 1000 - 750 0 605 920 750 0 3000 V 3000 - 500 0 1855 2 80 0 2250 0

^All voltages measured with respect to ground

(target at ground)

^Source provides additional factor of 10 intensity variation.

Figure 2- -System Schematic

296
Paper ^^o. 20

LABOiUTORY KXPl^UME:^! ON SOU\!l \VI:>ID II^TERACTIOlNi ^vITH geo:l\g.;etic field

E.M. Dubinin, 1 G.G. Managadze,^ I.M. Podgomyl

It is inDOssible to reoroduce in a laboratory all

phenoinena './hich exist at the interaction of the solar wind and

the Earth nagnetic field. In previous "jorks it was shown that

choice of exoeririental conditions permits one to simulate sone

of the most interesting effects, for exannle, the collisionless

shock near the Earth maqnetosnhere. .according to the principle

of limited sim.ulation, to reproduce the macnetosphere and the

collisionless shock the parameters of the artificial solar wind

should be:

13 -3

plasma density - n = I'J cm , electron temperature - T = 15 - 20 ev, embedded marrnetic field - B ~ 40 Oe,

c _ ion temperature - T. - d ev, stream velocity - V =

7 ' 1 3*10 cm/ sec.

At these conditions the plasma flux compressed the dinole magnetic field on the dav side and formed a conf irruration of the

magnetic field similar to the magnetic field of the Earth tail. Between the artificial m^agnetosphere and the plasma flow the

collisionless shock is displayed. Tae thickness of the shock

is in agreement with plasma theory and has the same order of magnitude in dim.ensionless expression as in space m.easurements

In the shock high level magnetic field microf luctuation was discovered. The dim.ension of the fluctuations is about that of the Larraour radium of ions. The microf luctuations ;;ere predicted in plasma tlieory as a result of Alfven and magnctosound

instability.

In addition to tlie shock investigations, measurements of the plasma density, velocity and magnetic field strength were

carried out before and behind the shock.

KEY WORDS: earth magnetic field, earth tail, magnetosphere

Dlpsmp theory, shock wave, solar wind

(Paper not available for publication)

^Institute for Space Research M USSR, Moscow, Russia

297

Paper No. 21

EFFECT OF PHOTODEGRADAT I ON ON ORGANIC COATINGS

Peter J . Hearst

Naval Civil Engineering Laboratory

Port Hueneme, California 93041

ABSTRACT: Clear and pigmented coating films were irradiated

with a mercury arc, a xenon arc, and sunlight. Strips of the

free films, consisting chiefly of alkyd, vinyl, and epoxy

coatings, were exposed in a quartz chamber. The major volatile

product, as determined by infrared spectroscopy, was carbon

dioxide, but the products varied depending on the irradiation,

the pigmentation, and the type of coating. Internal reflection

spectra indicated primarily the formation of carbonyl groups,

and the greatest changes were obtained for coatings that did

not have high initial carbonyl absorption.

KEY WORDS: Photodegradat i on , organic coatings, infrared analysis, volatile degradation products, internal reflection spectroscopy, ATR spectroscopy, surface degradation, clear and pigmented

films, mercury-arc irradiation, xenon-arc irradiation, atmospheric exposure.

As coatings become more durable, there is an increasing

need for better methods of predicting their performance. Much

of the deterioration of coatings exposed in the atmosphere is due to the effect of sunlight. Photodegradat i on of coatings

produces volatile products and the formation of these products

should be related to coating performance. Because the deterioration - and especially the photodegradat i on - of coatings

occurs primarily at the surface, it should also be possible to relate changes at the surface to coating performance. Internal

reflection spectra give infrared absorption data for thin surface

layers and should therefore be valuable in following chemical

changes at coating surfaces.

Many studies have been made of the volatile decomposition

products of polymers, especially under ultraviolet irradiation

and at very low pressures. However, comparatively little

299
attention has been given to the photodegradat ion of organic

coatings in a natural atmospheric environment. The volatile

degradation products of linseed oil films exposed to sunlight

in an atmosphere of oxygen was studied in 1927, ElU^ and more recently other studies of the decomposition of oil and a I kyd

films have been reported. [2,3,411 We have investigated the

volatile products from the irradiation of various clear and

pigmented organic coating films, Il5ll and the results are summarized below.

Internal reflection spectroscopy is being used widely to study coatings and polymers, but chiefly to determine the gross

composition. Gedemer Il6ll was the first to use internal reflection spectroscopy to investigate surface changes caused by

photodegradation. He studied po I ycarbona+es and found that,

except for the formation of one new peak, the chief changes in the spectra were a general weakening of the absorption peaks.

In the study of the photodegradation of organic coatings we have found evidence for carbonyl formation and again a general

weakening of other absorption peaks. [711 More recently Chan and

Hawkins in a study of polyethylene and other plastics have

related carbonyl formation, as determined by internal reflection

spectra, to the brittleness temperature of Irradiated samples. [81

Experimental Procedure

The coating films consisted primarily of alkyd, vinyl, and

epoxy films. Both clear films and films pigmented with titanium

dioxide were employed, and the thicknesses of these free films were approximately 40 micrometers.

The films were irradiated with a mercury arc, with a xenon

arc, or with natural sunlight. The approximate spectral distribution of these light sources is shown in Figure I. The

relative intensities for each light source are shown but no direct comparison between sources are available.

For the mercury-arc irradiations and most of the xenon-arc

irradiations, the coating films were placed in an exposure

chamber whose sides were made of quartz to allow the transmission

of the irradiating ultraviolet and visible light, and whose ends were made of I rtran windows so that the chamber could be used

as a cell in a infrared spectrophotometer for the identification

of the volatile products. The cell was filled with air at 50

percent relative humidity at 25°C, and the films were exposed

to 6 hours of mercury-arc Irradiation or 54 hours of xenon-arc

irradiation. In other experiments, without the exposure

^The numbers in brackets refer to the list of references

appended to this paper.

300
chambers, the films were exposed for 500 hours to the xenon arc

or for 5 months to atmospheric sunlight.

The significant infrared absorption peaks of the volatile

degradation products from some of the Irradiated films are presented graphically In Figures 2 to 4. The absorption peaks

of the vapors obtained from the clear vehicle films are indicated by vertical bars. The heights of the corresponding

absorption peaks from similar irradiations of the pigmented

films are shown by short horizontal bars which generally bisect

the vertical bars. The wavenumbers of the significant peaks

and their assignments are given. The yield of any Individual

product was proportional to the percent absorption, except for

the yield of carbon dioxide. (The absorption peak of the latter

was of such magnitude that the percent absorption was no longer

proportional to the absorbance.) Water was also produced, but

no quantitative measurements could be made. Some of the minor

products also have been omitted from the figures.

The effects of mercury-arc irradiation on the internal

reflection spectra of clear alkyd, vinyl, and epoxy films are shown in Figures 5 to 7. Other experimental results are summarized in Table I. Volatile Products from Clear Films

In the photodegradat Ion studies of various clear vehicle

films, ten different products had been Identified and five

functional groups characteristic of these products, or of additional products, were also identified. The products identified included carbon dioxide, carbon monoxide, methane, acetone,

formic acid, methanol, ethylene, acetylene, hydrogen chloride,

and water; the functional groups Included carbon-hydrogen

linkages, carbonyl groups, acetyl groups, vinyl groups, and

vinyl ester groups. The major product was carbon dioxide.

Both the rate of production of volatile products, and the

relative yields of products from the same film, were dependent

on the type of irradiation. Thus, 6 hours of Irradiation with

the mercury arc and 54 hours of irradiation with a xenon arc produced approximately equivalent decomposition of the clear

alkyd film, as judged by the yields of carbon dioxide, carbon

monoxide, and compounds with carbon-hydrogen and carbonyl groups,

and as shown in Figure 2. The same 6-hour and 54-hour exposure

periods were used In the study of other vehicle films even

though the relative yields were somewhat different for different

types of vehicles.

The oxidatlvely drying vehicles gave greater yields of

carbon dioxide than did the vehicles that dried by solvent

301
evaporation or by polymerization and this difference was accentuated in the xenon-arc irradiations. The oxidatively

drying vehicles also gave greater yields of ethylene and of formic acid. The vinyl and epoxy films, on the other hand,

gave greater yields of acetone and compounds with vinyl groups. Methane and acetylene were obtained from the mercury-arc

irradiations of all the clear vehicle films, and this might have

been caused by the shorter wavelength, and higher energy, light

of the mercury arc. Volatile Products from Pigmented Films

The yields of volatile products from the pigmented films were generally much lower than the yields from the clear films,

as might be expected, but the addition of pigment had no uniform

quantitative effects on the yields of individual volatile products. The differences in these yields is shown for alkyd films

in Figure 2, for vinyl films in Figure 3, and for epoxy films

in Figure 4. The predominant product from the irradiation of both the

clear and pigmented films was carbon dioxide. The possibility

was thus raised that the carbon dioxide production might in itself be used as an indication of the degree of photodegradat ion

In the xenon-arc irradiation of several air-drying films, the

reductions in the carbon dioxide yields from the pigmented films

as compared to the clear films paralleled similar reductions in yields of other volatile products. In similar irradiations of

vinyl films the yields of other volatile products were reduced

much more by pigmentation than the yields of carbon dioxide,

and the yields of vinyl groups and acetone were essentially

eliminated by pigmentation. Further work is thus needed to determine whether the carbon dioxide production can in itself

be used as an indication of film deterioration.

Light of sufficiently short wavelength to be strongly

absorbed will penetrate a clear coating film less than light of longer wavelengths. In a pigmented film in which the pigment

is covered by a thin layer of clear vehicle, the pigment therefore would be expected to have comparatively little effect on the photodegradation produced by light of very short wavelength.

This was borne out in the experiments. Acetylene and methane

were produced by the shorter wavelength I ight from the mercuryarc irradiations, and the yields were only slightly affected by

pigmentation. Acetone and vinyl groups apparently were produced

by light of longer wavelength, and the yields were greatly

affected by pigmentation. Generally, pigmentation reduced the

yields of volatile products from the mercury-arc irradiation less than it reduced the products from the xenon-arc irridation.

302
Surface Changes Detected by Internal Reflection Spectrocopy

In internal reflection spectroscopy, or attenuated total

reflectance (ATR) spectroscopy, the light beam is reflected at

the prism-sample interface in such a way that the beam travels

essentially within the prism except for a slight penetration

into the sample. This penetration is dependent on the relative

indices of refraction of the prism and the sample, and on the

angle of incidence of the light; and it is directly proportional

to the wavelength of the light.C9l] For ordinary infrared in- ternal reflection spectroscopy the penetrations may be in the

neighborhood of 2 to 10 micrometers. The spectra obtained are similar to the transmission spectra which would be characteristic of this upper layer.

The chief changes in the spectra, which resulted from the

irradiation of the coatings^ were increases in absorbance in the region of 1600-1800 cm (5.6-6.2 ym) and decreases in the

peak heights in other regions. The increases in absorbance

presumably were due to the formation of various types of carbony I groups in ketones, esters, or organic acids that were produced in the photodegradat i on processes. No specific reaction products could be identified, probably because of the large

variety of products that are formed and because of their

relatively low concentration as compared to the unreacted film

constituents. The decreases in peak heights that are observed

in other regions of the spectra are often due to a broadening

of the peaks and lowering of the baseline. These changes can

be caused by the disappearance of specific functional groups, by

a prol i feration' of the types of chemical linkages, or by poorer contact because of surface roughness.

There is no one specific change common to all the internal

reflection spectra that would be a good indicator of the degree

of photodegradat i on . Different changes were observed in the

spectra of the various generic types of coatings or of the

different polymers. The type of irradiation greatly affected

the relative effects on the spectra, and the presence of pigment

had varying effects.

The mercury-arc irradiations produced large changes on the

surfaces of clear vehicle films, whereas the xenon-arc irradiations produced comparatively small surface changes, as indicated by the spectra. This difference in effect of these

sources of radiation is again due to the difference in penetration of the irradiating light. The clear films are comparatively transparent to most of the light from the xenon arc. On the other hand the same films have a high absorbance for much

of the shorter-wavelength light from the mercury arc. The

mercury arc thus causes much greater surface degradation of the

clear film, even though the same total amounts of volatile

products may be formed in the xenon-arc irradiation.

303
The presence of pigment in the coating films greatly

reduced the changes in the internal reflection spectra that were produced during the same periods of mercury-arc irradiation.

On the other hand, the pigmented films exposed to the xenon-arc

irradiation showed slightly greater changes in the spectra than

those produced for the cleai" films, in spite of the fact that

the yield of volatile products was reduced. It is possible

that this greater change at the surface was due to reflection

of the light within the film.

The results of five-month atmospheric exposure were much more similar to those of xenon-arc irradiation than to those

of mercury-arc irradiation. General reductions in peak heights

were obtained in the spectra of alkyd, vinyl, and eooxy films,

and there was an increase in the carbonyl peaks of the epoxy

films. The 500-hour xenon-arc irradiations had lesser effects

on the alkyd film, approximately the same effects on the vinyl

film, and greater effects on the epoxy film, as compared to the five-month atmospheric exposures.

The results described above, as well as those of other

investigators, show that the chief effect of photodegradat ion on the internal reflection spectra is an Increase in carbonyl

absorption. It thus appears likely that such spectra will be most useful in studying the degradation of those coatings and

plastics which do not have high initial carbonyl absorption.

Changes in Spectra of Various Polymers

The changes in internal reflection spectra that were obtained for coatings of different generic types, under various sources of irradiation, and for clear and pigmented films are discussed below. Also listed are the major changes noted by

other authors in the spectra of other polymeric materials.

A I kyd Fi I ms . Mercury-arc irradiation of clear alkyd films

produced increased carbonyl abso'rbance in the 1600 to 1800 cm"' (5.6-6.2 ym) region which resulted in a widening of the base of

a large absorption peak at 1720 cm"' (5.8 ym) , as shown in Figure 5. This part of the new carbonyl formation was masked

by the carbonyl peak originally present. A new peak, or intensified peak was obtained at 1165 cm"' (8.6 ym) , but there

was a marked decrease in other individual peak intensities at frequencies below 1600 cm"' (above 5.6 ym)

Mercury-arc irradiation had a much reduced effect on the

spectra of pigmented alkyd films, as noted in Table I. Xenonarc irradiation had only a slight effect for either clear or pigmented films. Exposure of clear or pigmented films to the

atmosphere produced small reductions in peak heights.

304
V i ny I F i I ms . Mercury-arc irradiation of clear vinyl films

produced considerably increased carbonyl absorbance in the 1600-

1800 cm"' (5.6-6.2 ym) region, which was less strongly masked

by the original carbonyl peak than was the case for the a I kyd

film. The other peaks were again greatly decreased, as shown

in Fi gure 6

Mercury-arc irradiation of pigmented vinyl films gave a small increase in the carbonyl peak and considerable decreases

in other peaks, as noted in Table 1. Xenon-arc irradiation

produced only slight effects on clear or pigmented films.

Atmospheric exposure also produced only slight effects. Slight

effects after weathering for eighteen months also were reported

for the internal reflection spectra of several PVC formulations

by Chan and Hawkins. CS]

Epoxy F i I ms . Mercury-arc irradiation of clear epoxy films

produced strong carbonyl peaks at 1720 and 1655 cm~l(5.8 and

6.0 ym) . Other peaks were decreased, as shown in Figure 7. Mercury-arc irradiation of pigmented epoxy films produced

somewhat weaker peaks at 1720 and 1655 cm"' as shown in Table I, but had relatively little effect in decreasing other peaks. Xenon-arc irradiation (54 hr) of the clear film had slight

effects but xenon-arc irradiation of pigmented films produced

somewhat greater increases at 1720 and 1655 cm"' without affecting other peaks. Atmospheric exposure (5 mo) of both clear

and pigmented films gave greater increase in the two carbonyl

peaks and did decrease the heights of other peaks; however the

effects were not as great as in the mercury-arc irradiation

(6 hr) of the pigmented film. Longer xenon-arc irradiation

(500 hr) of the pigmented film, produced greater effects than

the atmospheric exposure, even though this irradiation had

comparatively little effect on the alkyd and vinyl films.

Po I ycarbonates . Mercury-arc irradiation of the polycarbonate samples was studied by Gedemer.E6ll He found an increase in

a peak at 1600 cm"' (5.6 ym), a general decrease in other peaks,

and changes in the relative intensities of some of the peaks.

Po I yethy I enes . Exposure in a weathering machine, which

produced higher relative amounts of ultraviolet light than the

xenon-arc, and also outdoor exposure, produced a strong carbonyl

peak at 1710 cm"' (5.9 ym) in the internal reflection spectra

of polyethylene samples, according to the studies of Chan and

Hawkins. HsH There was also increased absorbance in a broad

area centered near 1050 cm- I (9.5 ym)

ABS Resins . Outdoor exposure of aery I on i tr i I e-butad i enestyrene resins, produced hydroxyl peaks near 3300 cm"' (3.0 ym)

305
and carbonyl peaks near 1710 cm ' (5.9 ym) , according to Chan

and Hawkins.

References

1. J. D'Ans and S. Merzbacher, Chem. Umschau Fette, Oele,

Wachse, Harze, CUGFA, Vol. 34, 1927, pp. 283-291, 295-296

[Chem. Abst., CHABA, Vol. 22, 1928, p. I860].

2. E. B. Fitzgerald, Ind. Eng. Chem., lECHA, Vol. 45, 1953,

pp. 2545-2548.

3. E. B. Fitzgerald, ASTM Bull., ASTBA, No. 207, 1959, pp. 65-77.

4. R. S. Yamasaki, Journal of Paint Technology, JPTYA, Vol.

39, No. 506, 1967, pp. 134-143.

5. P. J. Hearst, ACS, Division of Organic Coatings and

Plastics Chemistry, Papers Presented, ACOCA, Vol. 26, No. I, March 1966, pp. 205-215; Vol. 27, No. I, April 1967, pp. 197- 206.

6. T. J. Gedemer, Applied Spectroscopy, APSPA, Vol. 19, No. 5, September 1965, pp. 141-146.

7. P. J. Hearst, ACS, Division of Organic Coatings and

Plastics Chemistry, Papers Presented, ACOCA, Vol. 28, No. I, Apri I 1968, pp. 672-681

8. M. G. Chan and W. L. Hawkins, ACS Polymer Preprints,

ACPPA, Vol. 9, No. 2, September 1968, pp. 1638-1643.

9. N. J. Harrick, Internal Reflection Spectroscopy, Inter- science Publishers, New York, 1967, p. 30.

306
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312
Paper No. 22

RELATION BETWEEN DOSE RATE EFFECTS AND LABORATORY

SIMULATION OF SPACE RADIATION DAMAGE

A. Roizds and R. Schuttler

REFERENCE: Roiz^s, A. and Schuttler R. , "Relation

Between Dose Rate Effects and Laboratory Simulation

of Space Radiation Damage," ASTM/IES/AIAA , Space

Simulation Conference, 14-16 Septembre 1970.

ABSTRACT: The experimental simulation of space ra- diation damage is usually done on a much shorter

time scale than the satellite lifetime. The defect

annealing in the semiconductor may lead to an ove- restimation of the real space degradation ; indeed

the irradiation time may be of the same magnitude

order than the annealing time constant. Cleland's

experiments (1) on germanium show that for a sufficient dose the defect number electron induced may

be half as much in the long irradiation case than in

the shorter one. The higher density of ponctual defects introduced by a greater flux rate enhances the production

of multiple defects as reducing the production of

vacancy-impurity complexes.

Different factors that can influence the defect

center distribution are analysed in the range -lOO^'C

to + lOO^'C.

It is shown that to improve the device hardness

against radiation, we must study the production rate

of defects in the bulk material coming from different growing methods and makers, as we know that the

damage depends on impurity constant.

^ONERA/CERT - Departement d' Etudes et Recherches en Technologie Spatiale 51, rue Caraman 31-TOULOUSE

FRANCE

2 Institut National des Sciences Appliqu§es et

Departement d' Etudes et Recherches en Technologie

Spatiale 31-TOULOUSE - FRANCE.

313
Moreover a second improvement can be reached by

annealing enhancement. Thus we must choose a material

v/hose defect annealing temperature is as near as possible than the device temperature operation. A

combination of these two requirements can help to se- lect an optimal crystal for devices in space environment.

It is shown that with a better knowledge of the

recovery process it will be possible to simulate long

irradiation time with reasonable operation time of a particle accelerator by irradiating devices at a controlled temperature higher than the usual temperature operation.

For example, calculations made on the E center

defect (vacancy - phosphorous association) show that one year irradiation at 300°K can be simulated with

the same dose, in one hour irradiation at 400 ^'K. We deduced than in a room temperature simulation

the number of defects, in this case, is overestimated by about 50% approximately,

KEY WORDS: Silicon, defect annealing, Space simulation , impurities.

1 - INTRODUCTION

Ground simulation of a semiconductor where ra- diation produces degradation in a much shorter interval of time than under real conditions.

In such a way we neglect all defect annealing,

and thus we overestimate the degradation effect of

energetic particles.

The higher density of punctual defects introduced by a greater flux rate enhances the productions

of multiple defects as reducting the production of

vacancy- impurity complexes.

As different defects have different influences

on measured physical properties, the degradation

coefficient becomes to be dependent upon the flux

rate.

The improvement of device properties against

space radiations can be obtained through a careful

study of the material electrical properties. Indeed,

314
the device caracteristics before irradiation depend

much less upon the impurities nature than the bulk

properties after irradiation.

Previous studies by Stein et Gereth (2) , have

shown that one could select a material whose annealing temperature was near the working temperature

of devices, thus lowering the sensibility to a slow

irradiation.

We show that it is possible to take into

account the annealing phenomena during an irradiation

by high flux rates if one maintains the devices at

a higher temperature than the device vzorking temperature.

The study of the major defect annealing enables

us to show that a simple relationship exists bet- ween irradiation duration and irradiation temperature.

Calculations concerning the E center (vacancyphosphorous assocaation) show, for example, that

space irradiation at 300°K can be simulated by a one

hour ground simulation at 400**K.

In classical simulations one overestimates the

defect number by about 50%

2 - INFLUENCE OF IRRADIATION CONDITIONS ON THE

PRODUCTION RATE

2-1 Temperature : The creation defect production rate varies with temperature in the low temperature ranges (3) , (4) , where a correlation exists

between the vacancy and the interstitial of the primary defect. This effect vanishes above 200**K in

Silicon.

2- 2 Flux rates : The experiment performed

shows that the defect production rates does not depend on flux rates up to a limit of 10^^ el/cm2s for

Silicon (5) . 3 - DEFECT ANNEALING

3- 1 Simple defect : For studying the defect

annealing at room temperature one must define exactly

the material caracteristics.

The following equation gives the variation of

the number N of defects with temperature T and time t:

315
is the activation energy

is the jump frequency

very often 1 < Y < 2 Well known defects are of type E center i.e vacancyimpurity associations.

HIRATA et Al.(6)have shown that annealing depended on the kind of impurity : we choose the vacancyphosphorous defect to show the importance of defect

annealing. It is caracterised by ar activation ener- gy E = 0.94 eV and a jump frequency K = 6 10 8 . a o Keeping the crystal at ordinary temperature,

only 35% of created defects remain one year after

the irradiation.

If we work using the real space conditions of

temperature, we have approximated the temperature

cycles of the solar pannels, as following. In shadow, the temperature is constant and equal to -20**C

and in sunlight the temperature is 27®C. The corres- ponding intervals of time for these temperatures

are respectively 30 and 90 mn in one cycle. We assu- me there is a sharp temperature transition between

these states.

We have calculated the ratio — , between the

o number of defects N effectively remaining after one year of irradiation and the number N^, obtained by

irradiation at 300 **K, by delivering the same total

dose over a short period prior to the diffusion times

of the defects.

N We have found ^ = 0.7, for a 3 years irradiation

o ^ = 0.47, for 4 year irradiation |- = 0.18, the

o o effect increases with temperature and the irradiation

time, for a given integrated dose.

The same calculus for a maximum temperature of

60**C and a one year irradiation time gives 5— = 0.01.

316
The E center will disappear for another center

of greater stability, the A center for example. In

this condition the physical meaning of a spatial irradiation is not the same as a classical irradiation

on the ground.

3-2 Complexe defect : Above a threshold

dose, the defect density can be so high that interaction among defects takes place and there is a new

defect non existant in a real irradiation. This result was obtained by FANG (7) who compared solar cell

degradations after successive cycles of irradiation

followed by a recovery.

For the other defects, the effect will be more

important if the annealing temperature is nearer the

device temperature operation : for example this is

the case ot doped lithium solar cells.

4 - SPACE IRRADIATION SIMULATION

The semiconductor devices degradation simulation is not possible during their space exposure

periods.

We propose to accelerate the diffusion phenomena by irradiating the samples, with accelerators

dose rate i.e. about lO^ el/cm^s at a temperature

rather higher than the real temperature of irradiation, for a reasonable period.

We consider a first order kinetics annealing,

the result obtained can be generalized for any kinetic order.

The recovery is governed by a differential equation of the form : f = - N(t) exp -

i.e. Log^ N(o)

= - K

o

t exp - ^

kT

N(t) is the number of defects remaining at the

time t after irradiation

N(o) is the number of defects introduced just

after irradiation

317
If we perform irradiation with the same dose on two identical samples kept at different temperatures

and T^, the number of defects present in the sam- ples will be identical at the end of the respective

time t^ and such that : E E

^o ^1 " kl^ = ^o ^2 - kf^

that is the ratio of the observation times in the

two experiments will be

~

- ^2 ^a 1 1 tT ^ kT ^ tT " t7 ^ 1 1 ^. Example - In the case of the vacancy-phosphorous

defext E =0.94 ev,

a the ratio of the irradiation times for = 300®K

and T2 = 400**K, will be :

.

. ^

....

^ = 1.2 10

^2

It is evident than a continuous irradiation of

one year at 300°K could be simulated by an irradiation of one hour at 400°K with the same dose.

5 - CALCULATION OF DOSES OF PARTICLES RECEIVED : EFFECT OF RECOVERIES

If we used as dose accumulator a device sensitive to minority carrier lifetime , the annealing

effect will induce an error that we evaluate here.

The measured dose <^

' is lesser than the real

dose see figure 1 . The quantity ~ - — is proportional to the number of

defects N.

T is the minority carrier lifetime after irradiation

Tq is the minority carrier lifetime before

irradiation

318
i - i = K = K'*' thus i = |.

In one case after a one year's irradiation — ,= 1.43

a four years irradiation — ,= 5.5

The annealing defect become appreciable during a lesser irradiation time.

6 - PREDICTION OF A SEMICONDUCTOR DEVICE

DEGRADATION

We consider a device sensible to the minority

carrier diffusion length.

A 20% degradation prediction after one year's

irradiation will be 17%.

A 20% degradation after four years will be 5%.

7 - CONCLUSION

The degradation prediction of semiconductor

device working in a space radiation environment re- quires the degradation study of the bulk material.

It is better to select a low annealing temperature

material, nearest to the working temperature : for

example the lithium doped Silicon is an extreme case. Many years of irradiation can be better simulated by accelerating the annealing phenomena : we

have defined a new irradiation procedure which gives

a better prediction of a semiconductor device degradation. We have shown that there is a simple relation between time and temperature irradiation. In a particular case, a one year room temperature degradation can be simulated by a one hour irradiation of

a sample maintained at 400°K with the same dose.

This problem will be important for long duration satellites, this is the case with technological

application satellites.

319
\

Figure 1 - Number of defects as a function

of irradiation integrated flux

320
BIBLIOGRAPHY

(1) A. HIRAKI, J.W. CLELAND, J.H. CRAWFORD - Radiation effects in Semiconductors

Proc. of the Santa Fe Conference (1967)

(2) H.J. STEIN and R. GERETH -

J. A. P. 39f 2890, (1968)

(3) R.E. WHAN and F.L. VOOK - Phys. Rev. 153, 814, (1967)

\A) H.J. STEIN - Bull. Am. Phys. Soc. 1],,194 (1966)

(5) D.E. HILL - Phys. Rev. 11_4, 1414, (1959)

(6) M. HIRATA et al . - J. A. P. 38, 2433, (1967)

(7) P. FANG - Phys. Rev. Letters, 20, (1966)

321

Paper No. 23

SOME PROBLEMS CONCERNING THE SIMULATION OF PLASTIC

FILMS DEGRADATIONS BY IONIZING RADIATIONS

A.J. Paillous ' and Ph. Fayet

REFERENCE: Paillous, A.J. and Fayet, Ph., "Some

Problems Concerning the Simulation of Plastic Films

Degradations by Ionizing Radiations," ASTM/IES/AIAA

Space Simulation Conference , 14-16 Septembre 1970.

ABSTRACT: Simulation in the laboratory, of the ac- tion of polyenergetic proton-and electron-fluxes

encountered in Space is necessary for a convenient

estimate of polymeric film materials. In order to

select the best simulation conditions, the following points were experimentally investigated on Poly (ethylene terephtalate) films, irradiated in vacuum by 1.5 MeV-protons and by 0.35-, 0.6-, 2.0

MeV-electrons : a) dose rate effects b) effects

of interruptions during irradiation (samples being

kept in vacuum) c) recovery during post-irradiation exposures to air or to vacuum d) effect of

the sample temperature during irradiation e) effect

of particles' energy f) effect of the particles'

nature

It appears that accelerated tests in a vacuum can

be used to predict the damage created during missions in Space which are of a longer duration than

To whom correspondance must be adressed.

2ONERA/CERT - Departement d' Etudes et Recherches en Technologie Spatiale 51, rue Caraman 31-Toulouse

FRANCE

^Centre National d' Etudes Spatiales - Direction des

Programmes et du Plan B.P. n" 4 91 - Britigny-sur-Orge - FRANCE.

323
that of the tests themselves. Electron damage in

Space can be simulated by a monoenergetic electron

irradiation which introduces into the sample the same absorbed energy dose than the computed dose

which would be absorbed in Space. For plastic films

there is no practical equivalence in the effects of

proton- and electron-irradiation, except for the

optical transmission properties. The need for an insitu testing of mechanical and optical properties is

outlined.

KEY WORDS: tests, evaluation, plastic films. Poly

(ethylene terephtalate) , radiation effect, space

simulation, proton damage, electron damage, radiation intensity, recovery, temperature effect, electron energy, proton energy, optical properties,

mechanical properties, viscosity, solubility, gel,

electron spin resonance.

I - INTRODUCTION

Simulation in the laboratory, of the action of

ionizing polyenergetic isotropic radiations encountered in Space (i.e. chiefly protons and electrons)

is necessary for a convenient estimate of materials

to be used for a particular mission.

Such a simulation depends on existent irradiation apparatuses (particle accelerators providing a monoenergetic non-isotropic irradiation with a limited range of energies) ; moreover it must be shortlasting and easy to bring into use (simultaneous

irradiations by electrons, protons and ultraviolet

radiations, must be done separately in the laboratory) .In order to select the best simulation conditions (which must be compatible with our own existing

apparatuses) for testing polymeric films, the following problems were studied:

a) recovery of defects during an exposure to the

air, following irradiation

b) dose rate effects, for same absorbed dose deposited in a sample maintained in vacuum (particles

being the same)

c) effect of interruptions during irradiations

324
(samples being kept in a vacuum during breaks)

d) effect of the sample temperature during ir.-

radiation

e) equivalence of effects for particles of same nature but having different energies

f) equivalence of effects for same absorbed

dose delivered by radiations distinct in their nature (electrons, protons)

II - EXPERIMENTAL PART

Material used in this study

The plastic material used in this study was

Poly (ethylene terephtalate) (P.E.T.). The sample of

P.E.T. was Terphanel film of thickness 0.0050 cm. Electron exposure apparatuses

Two 600-KeV electrostatic SAMES accelerators and

a 2 MeV Van de Graaff electron accelerator were used.

A same exposure device was used for these accelerators : The electron beam was diffused by a thin Aluminum foil (25 or 30 ym thick) . The diameter of the

irradiated surface was 150 mm. Flux homogeneity was

better than 10% (measured by five small cups) ; multi-diffused electrons were eliminated by use of a collimating system.

Samples were irradiated in a vacuum better than

10~^ Torr.

Film specimen were generally fixed on a frame

at 4 cm from the irradiation chamber back-part, and

were not externally cooled.

In the experiments reported in Tables 1 and 2,

and in figures 1 and 3, small film specimens were

laid on an Aluminum sample-holder ; the sample-holder temperature was adjusted by the use of a circulating fluid.

In all experiments, the maximum sample-temperature was estimated by the use of temperature-indicating thin plates (Temp-Plates^ , Mylar support)

Soci§t6 La Cellophane, 110 boulevard Haussmann,

Paris 8e - FRANCE

2Williams Wahl Co. Santa Monica - California

325
Proton exposure apparatus

A 2 MeV Van de Graaff proton accelerator was

used. A mechanical scanning system allows an homogeneous irradiation of specimens which are fixed on a rotating drum ; the diameter of the proton beam is

20 mm2. The specimens are irradiated in a vacuum better than 10~4 Torr. The maximum temperature of

samples was estimated by the use of temperature indicating thin-plates.

Optical measurements

Optical transmission measurements, in the 200-

2700 nm wavelength interval, were performed in air

with a Beckman DK2A Spectrophotometer. Except for

the measurements given in Tables 1 and 2 , and in

Figures 1 and 3, spectra were recorded with an integrating sphere attachment. The reproducibility at

any wavelength was within + 1 per cent. Measurements

given in Tables 4, 5 and Figures 3 and 4 are obtained after only a post-irradiation exposure to air at

the end of irradiations (samples being kept at room

temperature). For recovery experiments (Tables 1,2 ; Figure 1) , procedure is given in the text.

Mechanical measurements

Tensile stress-strain curves were recorded on

a Lhomargy testing machine. The speed of testing was

30 mm/mn. At the end of irradiation, samples were exposed to air for less than 5 minutes ; then they were kept in a 10~2 Torr vacuum for one hour ; they were subsequently kept in air (20*^0 + 1**C, 65% + 2%

relative humidity) for 23 hours ; measurements were performed in air at 20**C. For each irradiation, 6 rectangular specimens (105 x 80 mm ? 5 cm between

grips) were tested : the average measurement is given in the text) . For recovery experiments, samples were exposed for less than 5 minutes to air ; they were subsequently kept in a container in a 10"^ Torr

vacuum, at ambient temperature for 20 minutes during

their transport ; then access of air to these samples was allowed, and curves were recorded at intervals.

Viscosity measurements

Viscosity measurements were carried out at

326
40**C + 0,02°C in an Ubbelohde viscometer. Specimens

were dissolved at 60^0, for 15 mn, in a phenol-tetrachlorethane (1:1 vol.) mixture.

Solubility analysis and equilibrium swelling values

We used the experimental methods described by

MARCOTTE and Al. (10) for the measurements of the

insoluble fraction and of the equilibrium swelling

value

Electron spin resonance

Irradiated films were examined in a ThomsonHouston THN 252 Spectrometer (X band) at room temperature.

Spectra were either recorded in air (the irradiated plane film^ 5 x 0.6 cm, was contained in an open Suprasil tube) or recorded in vacuo (the irradiated plane film, 5 x 0.6 cm, was sealed in a Suprasil tube in a 10""" Torr vacuum) . Film was examined

with different orientations in the magnetic field ; measurements reported below were performed with the

irradiated film perpendicular to the magnetic field.

Ill - RESULTS

Post-irradiation recovery

. Variations of optical properties

A yellowish-brown coloration is produced by gamma irradiation (1)^ or electron-irradiation (2) of

Poly (ethylene terephtalate) in a vacuum. This coloration is less intense in samples irradiated in air.

A rapid post-irradiation decrease (3) in optical density of P.E.T. in air, was explained (1) by the reaction between the trapped radicals and the oxygen

which diffused into the polymeric film. The recovery

is limited : a rapid fading of the ultraviolet absorption occurs until a stable level is reached after a about 12 hours ; this stable value may be attributed

to the formation of cross-links between phenylrings.

In order to evaluate the extent of recovery, the

following experiments were carried out. Three samples,

A,B and C, were irradiated by 600 KeV electrons in a vacuum at 15-20*0 j after this period of

The numbers in parentheses refer to the list of

references appended to this paper.

327
irradiation, the samples were in contact with air

for 5 minutes ; then they were kept at 77 °K under vacuum (10"^ Torr) for 10 minutes ; they were subsequently warmed for 5 minutes up to 20'*C ; then access of air to the samples was allowed. Optical

transmission of sample A was recorded at intervals.

Sample B was kept in a vacuum (10~^ Torr) in darkness

at 20**C for 75 hours ; its spectrum was then recor- ded in air.

Table 1 gives optical transmission variations

during the exposure of A to air. Table 2 gives measurements for the sample B. It can be seen that the recovery extent at room temperature after 75 hours under a 10~2 Torr vacuum

is comparable with the recovery extent after 3.5

hours in air.

Therefore it can be inferred that the optical

recovery for sample A in a vacuum is quite weak during the earliest minutes. On account of this observation, it is therefore possible to overvalue

and undervalue the transmission which would be presumably recorded in-situ on sample A at the end of

irradiation without any contact with air. The upper

value is given by extrapolation to zero-time of the

transmission variations ; the lower value is given

by the first measurement in the air ; values deduced

from Table 1 by this method, are shown in Table 3.

Such a method must be used when in-situ recor- ding of spectra is not possible.

It is worthy of note that the recovery we observed for sample B during a 75 hours-exposure in a 10"2 Torr vacuum, at the end of a short exposure to

air (5 minutes) , may be greater than the recovery

which would be actually observed in-situ after

75 hours without any contact with air.

Similar optical changes are recorded during ex- posures to air following in vacuum irradiations (by

1.5 MeV protons, or by 600 KeV electrons) of all the

plastic films we tested (Polyimides, Pyrrones, Polychlorotrifluoroethylene , polyvinyl fluorides...).

The extent of changes and their algebrical sign

(degradation or improvement in optical transmission)

depends on the nature of the plastic, and, in some

cases, of wavelength interval.

. Radical Decay

Biaxially oriented films of Poly (ethylene

828
terephtalate) are reported to give E.S.R. spectra

after exposure to high energy radiations in vacuum

(4) . The spectrum comprises either 6 or 8 lines ac- cording to the orientation of film in the magnetic

field. During a post-irradiation exposure to air, a decay of radicals is observed, owing mainly to a destruction of radicals by oxyg3n (leading to formation

of carboxyl groups) and also to a recombination.

In order to compare the evolution of ultraviolet transmission with ESR signal decay, by exposure

to air, the sample C, which was treated in the same way as sample B (see the paragraph above) , was exa- mined by ESR at 20*C. The decay of radicals was studied by following the decrease in height of the central peaks of the first derivative ESR spectrum,

with time. In air, radical decay is directly related

(see Figure 1) to transmission variations, specially

at 495 nm where there is no permanent optical damage

(cf . Table 4) , and where recently irradiated films

present a weak absorption band (cf . Figure 4)

No signal decay is noted after 20 days for a film which was irradiated in vacuum at 20**C, then

allowed to be in contact with air for 5 minutes,

then sealed in a good vacuum (10"^ Torr) in a Suprasil tube and kept in darkness. This observation leads

us to think that no correlative transmission variation of irradiated films would have been recorded

in-situ for such a long time.

• Variations of mechanical properties

A very fast variation of mechanical properties

was observed during the very earliest minutes following the beginning of exposure to air of an in vacuoirradiated sample . Samples which are, at the end of

irradiation, qualitativaly very brittle and difficult

to handle, become flexible again within a few minutes. Tensile stress-strain curves of Terphane films

irradiated by electrons ( 600 KeV ; flux : 4.3 lO^^

e.cm"2) were recorded after different times of postirradiation exposure to air. The curves in Figure 2 show that an important change is produced quite rapidly : experimental conditions are described above

(see Experimental Part)

We must thus conclude that there is an imperious

need for an in-situ mechanical testing device for

plastic films.

329
Dose rate effect

Table 4 contains optical, mechanical and viscosimetrical measurements of Terphane films irradiated

in vacuum up to 2.16 10^^ electrons. cm-^ , with flux

densities ranging from 2.5 10-^^ e.cm""2,s~l to 2.0

10^2 e . cm-^ , s"~l . Nearly all of the measurements are

the same which show that there is no dose rate effect.

However slightly more damage can be observed for the

film irradiated with the highest intensity ; in this

particular case, we noted an important increase in

sample temperature during the irradiation : 80 or

90*C, which corresponded to a temperature much higher

than the glassy transition of P.E.T. (70-80°C) . It

is significant to note that a conservation of a non- irradiated Terphane film in vacuum at 90°C for 3 hours

does not change its optical transmission.

All of the electron-irradiated samples quoted

in Table 4 were easily solubilized in a phenol-tetrachlorethane (1 : 1 vol.) mixture at 60*^0. Only slight

variations in limiting viscosity niomber {n) (except

for the highest dose rate) were measured in the doserate interval investigated (0.46 - 3.73 Mrad.mn"^).

These resultats are different from those of TURNER,

PEZDIRTS and SANDS (5) who found a dose-rate effect

on electron - and gamma-induced viscosity changes

in Mylar, for dose rates ranging from 0.17 Mrad.

mn^l to 1.4 Mrad.mn"!. This discrepancy may be due to

the different origins of the materials. In any case,

if {r\} is considered as a direct evaluation of the

ratio of crosslinks Nq to fractures Nf, it appears

that a simultaneous increase of N^ and Nf will not

affect (n) ? no quantitative measurement of carboxylic groups was performed on our samples ; however

an important increase observed in the infrared absorption band at 3 000 cm"!, leads us to believe

that there is simultaneously cross-linking and degradation during electron-irradiations.

Intermittent irradiations

When exposure is segmented by interruptions in

irradiation (these interruptions are necessary, for

the longest exposures, during the nights for example)

results are the same as those noted after unsegmented exposures, provided that the samples are kept

in vacuum, in the irradiation chamber, during breaks.

330
In Table 4, irradiations n^S and n** 6 refer to such

unsegmented and segmented irradiations at the same

flux density (10^^ e.cm~2 .s"-'- during exposure) ; in

the latter experiment the total duration of the interruptions is equal to eight times the true irradiation

time.

Temperature effect

In order to confirm temperature effects, during

an electron exposure, two irradiations were performed above and below the glassy transition of Poly

(ethylene terephtalate) , The first experiment was

carried out with a sample holder cooled at 15**C by

a cooling fluid ; the second was carried out at 92 °C,

other conditions being the same (flux density : 1.5 10^^ e.cm~2.s"l; 1.5 MeV ; 1.5 hour). Transmission spectra were recorded in air 30 minutes after

the end of irradiation. Shape differences between

curves are observed (Figure 3 (a) ) : the absorption

band at 495 nm is more intense in the cold-irradiated sample. Moreover spectra recorded 24 hours after

the end of irradiation show (Figure 3 (b) ) a less

intense absorption, in the whole wavelength interval

350 - 700 nm, for the cold-irradiated sample. It can

be inferred that both transient and permanent effects

of irradiations, are different for irradiations performed above and below the glassy transition of

P.E.T. Similar difference may be expected for the mechanical properties of cold- and hot- irradiated samples. We believe that synergistic effect of temperature and irradiation (see the paragraph above) would

be further investigated by E.P.R. measurements. We

intend very soon to develop these studies at various

temperatures and also to investigate the effect of

sample temperature during a post-irradiation storage.

Energy dependence experiments

These experiments were performed only with electrons at 350 and 600 KeV. Computed values of stopping power in P.E.T. at these energies are respectively 3.02 and 2.67 MeV.cm"^ ; corresponding ranges

are 0.78 mm and 1.67 mm (to compare with film thickness 50 pm) . Two samples were irradiated to the same absorbed dose : 670 Mrad. Sample 3 was irradiated by

600 KeV electrons with a flux density equal to

2.5 lOll e. cm~2 .s~l for 24 h 00 ; sample 7 was

331
irradiated by 350 KeV electrons with the same flux

density for 21 h 15 mn. Slightly higher optical and

mechanical degradations are observed in the case of

the low-energy irradiation : see Table 5. Both sam- ples are easily soluble in trifluoracetic acid ; however in a phenol-tetrachlorethane (1:1 vol.)

mixture, microgels are formed when attempt is made

to dissolve the low-energy irradiated sample 7 : these microgels prevent viscosimetric measurements.

A greater number of cross-links is suggested by the

ultraviolet absorption spectrum and by the microgel

formation in the case of the low-energy irradiated

sample. The weak differences in stopping powers can- not satisfactorily explain the greater extent of

damages recorded at the lower energy. It is worthy

of note that degradation differences between electron-irradiated samples at 350 KeV and 600 KeV are slight in comparison with the differences between

irradiated and non-irradiated samples.

No attempt was made to perform such experiments

with protons : observation would be further complicated by differences in range of protons (for ener- gies lower than 2.0 MeV, these ranges are less than

the film thickness ; see below)

Effect of particle nature

Computed range and stopping power for 1 . 5 MeV

protons in P.E.T. are (8) respectively 4.94 10"3

g.cm~2 and 1.89 102 MeV.cm2.g~l. Thus exposure to

1.5 MeV protons irradiates only 35 ym, to compare

with the film thickness 50 ym. The absorbed energy

dose is not homogeneously distributed in these 35 m:

a greater dose is expected at the end of the proton

penetration. In Table 6 are given the average absorbed-energy dose D (expressed in Mrad) , and the ener- gy W absorbed by square centimetre, for three 1.5

MeV-proton irradiation and one 600 KeV-electron irradiation.

Figure 4 shows the transmission spectra of

Terphane films irradiated, in vacuum, under comparable conditions, by 1.5 MeV protons and 600 KeV electrons. It appears that, when comparing proton- and

electron-irradiation, the optical degradation can better be correlated with the energy W absorbed by

unit area, than with the absorbed - energy dose D. In Polymers, defects created by electrons and protons

332
are qualitatively the same : ionized and excited species, and radicals ; but optical transmission properties are more directly related to the total number

of created defects than to their spatial distribution within the solid (The case of optical reflective properties may be quite different)

Mechanical properties depends on the penetration depth of the particle and cannot be directly

connected with energy doses. From only the mechanical measurements reported in Table 6, it is impossible to conclude that a non-equivalence of effects of

different particles, in the same absorbed-energy

scale, exists (owing to the limited penetration of

1.5 MeV protons in P.E.T.).

Network formations during proton -and electron - irradiations can be easily compared. Samples irradiated by electrons at 350 KeV and at 600 KeV, up to

670 Mrad are easily soluble in trifluoracetic acid ; all samples irradiated by protons have a limited so- lubility in trifluoracetic acid, even at the lowest

dose. Measured values of insoluble fractions and

swelling of gels are given in Table 6. These observation lead us to think that molecular weight evolutions are different, for same absorbed energy doses,

in the case of protons and in the case of electrons,

owing to the spatial distribution of the created

defects.

Generally speaking, it must be outlined that, at

low doses (for instance lower than 100 Mrad) , the

great density of ionized and excited species and ra- dicals along the 1.5 MeV protons track, does not ensure a "homogeneous" irradiation in bulk, contrary to

irradiation ensured by 600 KeV electrons at the same

dose. A "homogeneous" irradiation by protons, with a random distribution of intermediate species created

in different tracks, will be only ensured for doses

higher than a D^^ value ; for these high doses, and

for penetration depths much greater than the film

thickness, there will be actually equivalence of effects after electron-and proton-irradiations. Taking

an approximative mean distance (10 A) between ions

or excited species created in a 1.5 MeV-proton track

into account, a value equal to a few hundred

Megarad is plausible for temperatures lower than the

glassy transition Tg (7) (9) ; a lower value of D^^ is expected for T > Tg owing to a limited diffusion

of created defects in the polymer. According to our solubility measurements, D-j^ appears to be higher than

670 Mrad for 1.5 MeV protons at T < Tg.

333
IV - CONCLUSIONS

a) From the great extent of the recoveries observed during a post-irradiation exposure to air of

in-vacuo irradiated polymeric films, we must conclude that all samples should be measured without any

contact with air.

For a convenient estimate of in vacuo-irradiated samples, an in-situ mechanical and optical testing device is absolutely necessary.

b) When in-situ measurements of optical transmission are not possible, an extrapolation method

can be used.

c) In-vacuo recovery is very weak according to

optical and E.S.R. measurements.

d) No dose-rate effect was observed during ex- posures with flux densities ranging from 2.5 lO-'--'- e.cm"2.s-l to 2.0 10^^ e,cm~^ .s"^

e) Intermittent irradiations and unsegmented

exposures lead to the same sample degradations

(even for breaks so long as eight times the true irradiation time) . f) From the three latter statements, it can be

inferred that long-lasting irradiations in Space can

be validly simulated by much shorter experiments

(some dozen hours or so)

g) Optical properties show a clear temperature

effect, near the glassy transition point. A correlative effect is expected for mechanical properties of

the irradiated sample. Thus a strict control of the

sample temperature is necessary during irradiations.

h) Energy dependence experiments show a slightly greater extent of degradations during a 350 KeVelectron irradiation than during a 600 KeV-electron

irradiation for the same absorbed energy dose. Such

an effect should be confirmed at low energy (150 - 300 KeV)

This effect being apparently slight, it seems possible to simulate the polyenergetic electron irradiation which takes place in Space, by a monoenergetic electron irradiation delivering approximately

the same energy dose (Mrad) into the material. However, according to the above abservation, it seems more advisable to simulate the degradation of plastic films by electrons in low orbite (for which 80

or 90 % of the observed energy dose is due to electrons with energies lower than 250 KeV), by a lowenergy electron irradiation (200 KeV for instance)

care must be

334
taken than the range corresponding to the selected

energy will ensure irradiation of the whole film

thickness.

i) For the highest absorbed doses (D > ^j^) ' provided that the range of the used protons is much

greater than the film thickness, proton-and electroninduced degradations are expected to be equivalent.

However, the range of the majority of protons

encountered in Space being short with regard to the

film thickness, it is not practically possible to use this equivalence in order to simulate a protonirradiation by an electron-irradiation, or reciprocally.

ACKNOWLEDGMENTS

The authors v/ish to express their gratitude to

Messrs. LEVI, TALAGRAND, AUROUX, AMBROSI and MAROT

of Laboratoire Central de I'Armement (ARCUEIL - FRANCE) for their assistance in irradiating the samples. They also wish to acknowledge the very valuable assistance of Mrs. HENOUX of L.C.A. (ARCUEIL,

FRANCE) in performing most of the optical tests,

and of Mr. BELLOC of Laboratoire Central des Industries Electriques (FONTENAY-AUX-ROSES , FRANCE) in

performing mechanical tests on samples.

REFERENCES

(1) D. CAMPBELL, D.T. TURNER - "Influence of doserate on Radiation induced network formation in Poly

(ethylene terephtalate) " Polym. Letters 5, 471-475,

(1967)

(2) N.A. SLOVOKHOTOVA , G.K. SADOVSKAYA, V.A. KARGIN

"Effects of high speed electrons on the structure of

P.E.T." J. Polym. Sc. 58^, 1293-1299, (1962)

(3) J.W. BOAG, G.W. DOLPHIN, J. ROTBLAT - "Radiation

dosimetry by transparent plastics" Radiat. Res. 9_, 589-610, (1958)

335
(4) D. CAMPBELL, K.ARAKI, D.T. TURNER - "ESR Study

cf free radicals formed by y "I^J^cidiation of Poly

(ethylene terephtalate) " J. Polym. Sc. 4, 2597-2606,

(1966)

(5) D.T. TURNER, G.F. PEZDIRTZ, G.D. SANDS - "Influence of dose rate on radiation induced network

formation in Poly (ethylene terephtalate)" J. Polym.

Sc. 4, 252-254, (1966)

(6) A. PAILLOUS - "Action of ionizing radiation on Poly (ethylene terephtalate)" A bibliography. 21 ref.

ONERA/CERT/DERTS NP-03-1-68, Oct. 1968.

(7) A. PAILLOUS - "Simulation of ionizing radiation

acting in Space - The case of structural polymeric

materials" ONERA/CERT/DERTS NT-03-3-69, July 1969.

(8) R.P. HENKE, E.V. BENTON - "Charged particle

tracks in polymers : n** 3 - Range and energy loss

tables" USNRDL TR-1102 (1966)

(9) A. CHARLESBY - "Atomic radiation and Polymers "- Pergamon Press Oxford - London - New York (1960)

(10) F.B. MARCOTTE, D. CAI4PBELL, J. A. CLEAVELAND

D.T. TURNER - "Photolysis of Poly (ethylene terephtalate)" J. Polym. Sc. Al, 5, 481-501, (1967).

336
o

•HP

P. O

B

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nra Ln un IN o oc oc oc OC c

B

C ^ oc CM o to Ln oc oc 1

B

CI a- 1—1 o OC o to 1 to

s

coto

CO 1

a; s: op

P Ti

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to oc

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oc to

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ri* CM

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O

•Ha

4) 'H

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o

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337
Optical Transmission (%) at 350 nm

375 nm

400 nm

450 nm

400 nm

550 nm

600 nm

700 nm

Before

irradiation 65.8 71.2 73.5 76.5 77.5 79 79.8 81.2

75 hours

in-vacuim after

the end of

irradiation

16 36.5 52 67 70.5 75 77 80

TABLE 2 - Sample B - Optical transmission

after a post-irradiatign

storage in-vacuum ( 10 Torr)

Wavelength

(nm)

Initial Transmission

i%)

350 7.8 - 9.8

375 22.4 - 26. 400 40.3 - 43. 450 59.5 - 61.5

490 63.9 - 65.9

TABLE 3

- Fresumed transmission

in-vacuum at the end of

irradiation

338
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341
0 100 290 300 TIMS (mn) Fig. 1 - ESR Signal decay and transmission variations

during a post-irradiation exposure to air

at room temperature

A Optical transmission

+ Central Peak, left part of the ESR Spectrum

• Central Peak, right part of the ESR Spectrum

STRAIN CA) Fig. 2 - Tensile Stress-Strain curves of Terp^gne fil^s

irradiated by electrons (dose 4.3 10 e.cm ) The curves are drawn as a result of the following

times of post-irradiation exposure to air

• 1 mn 20 s + 10 mn 30 s X 3 mn SO s 6 mn 30 s Y I'; mn 20 s Time origin corresponds with opening of the container

342
0. /

c» y—

o\—y 1 , 1 1 300 400 500 600 100

(a)

100

(b)

Fig. 3 - Temperature effect during irradiation

Spectra recorded after a 30 mn (a), 24 hours (b),

post-irradiation exposure to air at room

temperature

. . . Irradiation at 92°

Irradiation at 15°C

343
Fig. 4 - Transmission Spectra of Terphane Films

irradiated by protons or electrons (after

a 15 mn exposure to air) A 1 o MeV-protons

B 1 . MeV-protons

D 1 , S MeV-protons

C 600 keV-electrons

= 0 77 10^4

p cm = 1 S4 10^4 p- cm

$ - 3 OS 10^4 p. cm = 2 16 10^^ e. cm

344
Paper No. 24

ULTRAVIOLET AND CHARGED PARTICLE DEGRADATION OF

ALUMINUM AND SILVER COATED FEP TEFLON SECOND SURFACE MIRRORS

F. G. Cimningham B. L. Bean and S. G. Park'

ABSTRACT: An experimental program has been undertaken in whic?

bare FEP Teflon films were subjected to ultraviolet irradiation and

charged particle bombardment, and in which five mil aluminum and

silver coated FEP Teflon films were: i) ultraviolet irradiation in vac- uum and oxygen; ii) irradiated with 5, 10 and 25 or 30KeV electrons

and protons; and iii) exposed to simultaneous ultraviolet-proton and

ultraviolet-electron bombardment

The charged particle irradiations indicate that a measurable

change occurs in the spectral reflectance of both silver and aluminum

backed material at a dose of about 1x10 particles/cm^, and that this degradation increases as the total dose is increased. This spectral reflectance change appears first and is the most pronounced in the ultraviolet for a given total dose. At a total dose of Ix 10 particles/cm^

the surfaces generally look cloudy and oftimes mottled. Upon closer

examination this mottling is found to be due to the presence of Litchenberg

figures within the teflon—especially in the case of electron irradiation—

which are caused by charge storage and subsequent electric discharge.

The calculated changes in integrated solar absorptance generally agree

with those reported by other observers. The ultraviolet results con- firm that silver coated teflon is stable when irradiated in vacuum and

indicate that aluminum coated material undergoes a small decrease in integrated solar absorptance (i.e. , an improvement) when soir-roriiated. KEY WORDS: aluminum coated teflon, electron degradation,

proton degradation, silver coated teflon, teflon film, ultraviolet degradation.

^NASA/Goddard Space Flight Center, Greenbelt, Maryland

^Electromechanical Research, Inc., College Park, Maryland

345
INTRODUCTION

Metalized type A FEP-*- Teflon films, used as second surface mirors , have shown much promise as suitable thermal control coatings for spacecraft. In particular five mil silver backed FEP film has been

chosen, pending the outcome of the experiments now in progress, to be

a part of the thermal control coating on the following spacecraft: IMP

I, scheduled for laimch in October 1970; OAO-B, scheduled for launch

in late 1970, s\ scheduled for laimch 1 May 1971 and OAO-C. This

tentative certification for flight use is based upon findings to date which

indicate that teflon second surface mirrors exhibit exceptional ultraviolet stability, and have shown no change in integrated solar absorptance for low energy electron and proton bombardment at total fluxes

less than 1x10^^ particles/cm^

These data are the results of thermal data interpretation obtained

from limited flight testing as well as from several series of laboratory

studies. Five mil aluminum coated FEP Teflon was flown as the outer layer of the lower thermal shield (sunlit) on Mariner n (Venus), while

1 mil aluminum coated FEP film was used in the same capacity on Mariner V (also Venus). Results from measurements made on this surface from Mariner V indicate an apparent Aa of about 0. 043 after 5000 equivalent sim hours.

On OGO-6, launched in Jxme 1969, two mil silver coated FEP has

been used in conjunction with black paint (75% FEP and 25% black paint) as the control surface of the solar facing side of one of the experimental packages. To date (May 1970) this package has performed satisfactorily, and no gross change in the solar absorptance of the teflon has

been noted. OGO-6 also has a coatings experiment, one sample of which is five mil silver coated teflon. After 5000 sun hours, performance was still quite satisfactory .4 Laboratory investigations with

which we are aware have been carried out by: i) The Boeing Co. ^,

FEP is the acronym for the copolymer fluorinated

polyethylenepropylene

'Teflon is a registered trademark of E. I. du Pont de Nemours

company.

W. F. Carrol; "Mariner V Temperature Control Reference Design, Test and Performance," AIAA paper 68-791 presented at AIAA 3rd

Thermophysics Conference, Los Angeles, June 1968. 'Archie Fitzkee; Private Communication.

*L. B. Fogdall and S. S. Cannaday; "Dependence of Thermal Control

Coating Degradation Upon Electron Energy. " Final report to NASA/

GSFC for contract NAS5-11164, May 1969. (Also AIAA paper 69-

643); and L. B. Fogdall and R. R. Brown; "Degradation Tests on Thermal Control Materials," Contract NAS5-11219, Final report to NASA/GSFC, May 1970.
ii) The General Electric Co.-"- and iii) Electro-Mechanical Research,

Inc.^ It is with this latter work that this paper is concerned.

APPARATUS DESCRIPTION AND EXPERIMENTAL PROCEDURE

A. Irradiation System

A diagram of the irradiation system is given in Figure 1. The sys- tem can be used for three kinds of irradiation: solar simulation, proton or electron beam, and solar simulation plus proton or electron

beam.

The solar simulator consists of an ultraviolet source lamp, focusing optics, and two separate vacuum chambers. The source, a 1000

watt Xenon gas filled arc lamp,^ is focused and collimated by two quartz

lenses and transmitted through a quartz window onto a sample holder in the vacuum chamber. The level of radiation at the sample surface is adjustable from zero to five solar constants (one solar constant = 1350

watts/m^) and monitored with a solar cell detector. For solar simulation work alone, the gate valve to the accelerator is closed.

The proton and electron beam source is an RF ion source,^ electrostatic collimating lens, and magnetic analyzer. Hydrogen gas is admitted into the source through a palladium leak and ionized by an RF

field . Depending upon the applied potential , protons or electrons are drawn from the discharge and formed into a uniform beam. The magnetic analyzer is used to remove neutral atoms from the beam as well

as ions of other than the desired charge-to -mass ratio. The uniform

beam then passes through a four foot long drift tube and is scanned

with a Beam Scanner.^ The beam scanner is mounted in a tee as shown in Figure 2, and it can scan either the verticle or horizontal

dimension of the beam. The beam finally falls on the sample holder,

where flux ranges are available from 1x10^° particles/cm^ -sec to 1x10^^ particles/cm^ -bsc.

B. Linder and R. N. Griffin; "Development of Series Emittance Thermal Control Coatings," NASA CR-66820; Final report to NASA/LRC

for Contract NASl-8603, September 1969.

'M. J. Brown and S. G. Park; NASA/GSFC Contract NAS5-9469, completed August 1969; andR. Thompson, B. L. Bean and S. G. Park,

NASA/GSFC Contract NAS5-9469, completed May 1970. 'Hanovia model

;Ortec, model C-SO-173K

'Physicon, Model 512

347
B. In Situ Measurement System

The In Situ Measurement System consists of the sample holder in- side a quartz finger, a double beam recording spectrophotometer j-'- and

an absolute integrating sphere.^ The sample holder is adjustable so

that the sample can be removed from the sphere for a 100% reflectance

measurement. For this portion of the measurement, both a reference

beam and a sample beam are alternately incident on the barium sulfate coated sphere wall, the sample beam traversing the quartz finger while

the reference is positioned so as to miss it entirely. The two beams

are diffusely reflected and detected with either a lead sulfide cell or a

photomultiplier tube. The spectrophotometer then automatically ratios the two beam energies and plots the result as a fimction of wavelength.

The sample is then lowered into the quartz finger, and the sample beam

is now reflected from the sample surface before it hits the sphere wall. The reference beam is still in the same position, and the spectrophotometer again ratios the energies of these two beams.

The vacuum is maintained by a 25 L/S ion pump at the sample end

and a 4-inch diffusion pump at the accelerator end. Rough pumping is accomplished with several sorption pumps. Pressures in the 10"^ torr range are maintained at all times.

C. Experimental Procedure

The ultraviolet in situ experiments were conducted first. Three

initial and three final reflectance measurements were made on each

sample: in air, in air in the quartz finger, and in the proper environ- ment. The two environments investigated were vacuum and oxygen. The oxygen was allowed to leak slowly through the system so that a

positive pressure of oxygen was always maintained. The samples were

five mil teflon coated with aluminum or silver, and they were clamped

directly to the water-cooled sample holder. The reflectance measurements were made with the wavelength and detector ranges shown in Table 1; and this table was used for all other reflectance measurements as well. The irradiation rate was at five solar constants, and

each sample was irradiated for 120 hours.

The electron beam in situ experiments were conducted using a

different accelerator port. The samples used were five mil teflon coated with silver; and these samples were generally clamped directly

to a sample holder. Further discussion of the various experiments

will follow later in the paper.

'"Beckman Instruments, Inc. , Model DK-2A

2Gier Dunkle, Inst.

348
The proton beam and proton beam plus ultraviolet experiments were done in the same chamber as the earlier ultraviolet experiments.

The samples used were five mil teflon coated with silver; and they were glued to 3/4-inch squares of 0. 060-inch aluminum with an alu- minum filled epoxy putty. These samples were clamped to the sample

holder, with several small squares of aluminum foil placed between

the 60 mil aluminum mount and the sample holder, as in Figure 3. This method was used to obtain better sample cooling. The ultraviolet

irradiation was at a five solar constant rate; and the proton flux was

IxlO^^p/cm^ - sec to 10^^ p/cm^ and then 10^^ p/cm^ - sec to 10^^ p/cm^, with measurements made at 1x10 p/cm^ , 5x10^^ p/cm^,

and Ix 10^^ p/cm^.

EXPERIMENTAL OBJECTIVES AND DIFFICULTIES

The proximity of EMR to GSFC has made it possible for close col- laboration throughout the entire period of this work. Our original objective was to conduct a systematic investigation of bare, aluminum

and silver backed 2, 5 and 10 mil (later changed to 7. 5 mil) Type A FEP

Teflon film. The initial investigations were simply screening tests, involving short term ultraviolet (240-280 ESH) studies and low energy

electron and proton bombardment of the various thicknesses of bare

material and of five mil silverized and aluminized teflon, undertaken

for orientational purposes and possible indication of the correct ap- proach for a more through investigation.

These screening tests showed that bare and metalized Teflon were

stable to short term ultraviolet irradiation, even though a mercury lamp was used instead of the Xenon lamp incorporated with the in situ facility

previously described. The preliminary charged particle irradiations

indicated quite clearly that teflon when so irradiated tends to turn brown. Measurements showed a general decrease in transmission with

the greatest decrease occurring in the ultraviolet region. As the results were similar for the various film thicknesses employed and since five mil film seems to be most desirable from a thermal control viewpoint,

the systematic investigation of charged particle effects was limited to a study of five mil silver and aluminum coated material. No quantitative transmission data from these early irradiations is presented here,

however, since certain nonuniform patterns were present on these samples, particularly on those irradiated with electrons. This manifested

itself, in some cases, by a change in the degree of brownness at various parts of the sample, and in other cases as a browning of only a

portion of the sample. This nonuniform pattern was not a regular geometrical figure but looked rather like a shadow graph of a moving

starfish.

349
A subsequent and systematic series of electron and proton irradiations was begun. Alignment of the beam was made and monitored by

visual means using a phosphor and the irradiations were made as carefully as possible. No beam nonuniformities were apparent and the current as measured by a Faraday cup, about the same size as the sample

and placed adjacent to it, remained steady. Measurements made by

moving the Faraday cup with the sample absent showed no beam irregularities. However, the irradiated samples continued to show nonuniform irradiation patterns, albiet to a lesser extent than previously, and

the electron irradiated samples were still the worse as viewed by eye. The damage as shown by the spectral reflectance curves was quite extensive at a total dose of Ix 10^^ particles/cm^ and the damage as measured by the increase in integrated solar absorptance was greater (sometimes excessively) than that reported by other mvestigators.

This latter fact, in itself, is no particular reason to suspect the

data. But, before this series of irradiations was completed and before we were able to devise a better method for monitoring the beam an electrical storm connected power failure coupled with the failure of a "fail safe" device on the accelerator allowed a backflow of diffusion pump

oil into the system. After the subsequent cleaning, a preoperational

check showed that a beam misalignment located at the exit of the ion source did, in fact, exist. Because of this, it is our belief that the nonuniformities were due to extraneous particles hitting the sample at angles different from the primary beam. This secondary beam could

be reflecting or scattering from a wall of the beam tube and would not necessarily be picked up by the Faraday cup. It is probable that it was

picked up but that the resolving power of the cup was too low to show it. For subsequent measurement of charged particle fluxes the idea of a

separate Faraday cup was abandoned and the sample holder itself, electrically isolated from the sample chamber, was so used. In any case, we believe that the severe damage observed during this series of measurements was excessive and was due to a radiation dose larger than was

intended and what we believed it to be. For this reason the quantitative

data obtained during this period will not be reported here.

After accelerator repair further irradiations were held in abeyance

imtil a beam scanner consisting of a vibrating tungsten wire could be

installed in the beam tube. With this device it was possible to continuously monitor the beam during irradiation and in this way we were as- sured that the beam was properly aligned and that it remained reasonably uniform.

Since the problem of beam nonuniformity was most apparent with

electrons an anticipated complete rerun of the electron irradiation program was begun, using a flux rate of 3x 10^^ e/cvo? - sec. This program was again suspended after two samples had been irradiated with

electrons; the first at lOKeV to a total dose of Ix 10^^ e/cm^ and the

350
second at 25 KeV to a total dose of 2 x 10^^ e/cm^. When the first sample was removed from, the in situ chamber, after completion of the in situ measurement and during the preparation for the air measurement,

a small circular brown spot 2-3 mm in diameter was discovered, lo- cated at the very center of the circular sample. The in situ measurement included this brown spot, while the air measurements were made

both on and off spot. The second sample developed a similarly located

small brown spot. Because of this the regular irradiation procedure was suspended imtil a series of studies pertaining to this brown spot

formation could be conducted. These experimental results will be discussed elsewhere.

The browning phenomenon did not occur for proton irradiation but

time and financial limitations, created by the unexpected brown spot

studies, made it necessary to reduce the planned scope of the proton

irradiation studies. We were, therefore, unable to repeat all of those

irradiations made prior to the discovery of the misalignment of the beam at its source. Studies were limited to five mil teflon and the ir- radiation limited to two beam energies, 5 KeV and 25 KeV.

In toto, with the properly aligned and continuously monitored beam,

four samples were irradiated with 25 KeV protons and two with 5 KeV

protons. Each group of samples was divided in half; one half was ir- radiated with protons only and the other half was irradiated with simultaneous ultraviolet light and protons . In all cases the proton irradiations were the same for each pair. All samples were proton irradiated

at a dose rate of IxlO^^p/cm^ -sec. to a total dose of 1x10^^ p/cm^;

then one pair of the 25 KeV irradiated samples and the pair of 5 KeV ir- radiated samples was further irradiated at a rate of IxlO^^/cm^ -sec.

to a total dose of 1x10^^ p/cm^. The simultaneous ultraviolet irradiation was done with an Xenon lamp at an intensity of five suns.

Internal bureaucratic delay in the negotiation of a new contract to continue and finish the experimental effort has precluded the acquisition

of any additional data. We can only hope that the experimental program

will have been completed in time for the results to be given when this paper is read in September.

Concurrent with the charged particle studies, eight samples (four

each of aluminum backed and silver backed) of five mil FEP were ir- radiated in the in situ chamber at a rate of 5 ultraviolet suns with an Xenon lamp, to a total dose of 600 ESH. Four samples were irradiated

in vacuum and four in an oxygen atmosphere. We believe a total ultraviolet dose of 600 ESH is sufficient to confirm the previous work which

has indicated that FEP films are very resistant to damage or changes

brought about by ultraviolet irradiation. The irradiation in an oxygen

environment was undertaken to determine whether or not irradiation in the presence of an active gas would introduce any differences.

351
EXPERIMENTAL RESULTS

Even though the beam was nonimiform and total dose uncertain, the preliminary charged particle irradiations on bare FEP Teflon films in- dicated quite clearly that the samples have a tendency to turn brown.

The intended flux rate was 3x 10^ particles/cm^ -sec. and total in- tended dose 1x10^^ particles/cm^ ; the actual dose appears to be

greater than this figure, but its true value is unknown. This browning was distributed over the irradiated portion of the samples , total dose

differences being indicated by a varying degree of brownness. This

browning is generally more apparent on the bare samples than on those

with metal backing. The widespread browning, when it occurs, appears

to be distributed through the material and the material remains clear. Often, however, the surface of charged particle irradiated teflon be- comes cloudy or milky in appearance; i.e. , a fine white powder ap- pears on the surface which can be wiped off. Its presence often helps

obscure the browning. This powder may be depolymerized teflon or a chemical reaction product of the monomer or, even, some residual

catalyst. The transmission of the browned teflon decreases slightly in the visible and near infrared, and (as would be expected) much more

strongly in the ultraviolet. The ultraviolet cutoff moves noticeably to longer wavelengths.

The results of the ultraviolet (Xenon lamp, 600 ESH exposures) ir- radiations are: i) five mil silver backed teflon ultraviolet irradiated in vacuum and left in vacuum is essentially stable, showing improvement

after re-exposure to a normal laboratory atmosphere; i.e. , its reflectance increases. The integrated solar absorptance decreased on

the order of 7. 5 percent. ; ii) five mil aluminum backed teflon ultraviolet irradiated in vacuum and left in vacuum tends to improve slightly; i.e. , the reflectance increases overall. The reflectance actually de- creases in the ultraviolet but increases in the visible and near infrared.

The integrated solar absorptance decreases on the order of six percent.

However, when these samples were reintroduced to a normal laboratory atmosphere the reflectance changed, having decreased from its initial value in the ultraviolet and visible regions while remaining essentially the same in the near infrared. The net effect was a degradation. The integrated solar absorptance increased on the order of 5. 8 percent; iii) five mil aluminum backed teflon ultraviolet irradiated in oxygen tends to improve; i.e. , the reflectance increases. The de- crease in integrated solar absorptance was on the order of 8. 5 percent;

and iv) five mil silver backed teflon ultraviolet irradiated in oxygen

For clarification see preceding section.

352
undergoes a large degradation; i.e. , the reflectance decreases. The

increase in integrated solar absorptance was on the order of 76 percent.

The spectral reflectance of the two electron irradiated samples,

on which the localized central brown spots first appeared, was measured from 0.25 to 2.4 microns. The initial value of the integrated

solar absorptance of the lOKeV irradiated sample, measured in vac- uum, was 0. 071. After an exposure of 1x10^^ e/cm the integrated

solar absorptance had become 0. 076; and finally, after a total dose of 1x10 e/cm the integrated solar absorptance had increased to 0. 108—

an increase of 52 percent. However, this in situ measurement was made prior to the discovery of the central brown spot and, therefore,

included it. The samples were remeasured in air (as is customary)

off center; i.e. , the beam did not strike the central brown spot. The

initial and final values of the integrated solar absorptance were 0. 070

and 0. 089 respectively; representing an increase of 27 percent. The

general appearance of the sample off the spot was typical for charged

particle irradiated samples and, hence, the latter measurement is probably reasonable. The initial and final values of the integrated solar

absorptance for the 25KeV irradiated sample, measured in air and off spot, were 0. 071 and 0. 104 respectively; representing an increase of 46 percent. However, this sample received approximately twice the dose of the lOKeV irradiated sample or about 2x10 e/cm . The

electron irradiation studies are not yet complete, but our experiments

have indicated that avoidance of the brown spot formation is probably

not possible. We shall simply make all measurements off spot and

note them as such.

Microscopic examination of the first sample which developed the

browning of the central spot showed it to be clear and uniformly brown

while the overall surface of the sample was somewhat cloudy. Several

additional samples were electron irradiated at varying dose rates in an

attempt to determine the nature of this browning and what might be done

to prevent it. All of these samples were five mil silver coated teflon. Sample number one (1) was irradiated with SOKeV electrons at a

rate of 1x10^^ e/cm^ -sec. to a total dose of 1x10^^ e/cm^ . A central

spot appeared but it was not brown. The entire surface of the sample,

except for the central spot, did appear cloudy; i.e. , covered with a

thin layer of fine white powder. The central spot itself looked, upon

microscopic examination, as if the electron beam had not struck it at

all. Sample number two (2) was irradiated with SOKeV electrons at a

rate of Ix 10^^ e/cm^ -sec. to a total dose of Ix 10^^ e/cm^ . A centrally located, extremely bright, white opaque spot was formed. It had no visible structure. No attempt was made to determine its thick- ness or if it were composed of a white powder which could be wiped off.

353
Sample number three (3) was covered with a coarse mesh screen which was groimded. This screen was in contact with the metal retaining ring clamping the sample, and thereby, to the sample holder itself. This sample was irradiated with 30KeV electrons at a dose rate of 1x10 e/cm -sec. to a total dose of 1x10^^ e/cm . The grid pattern

of the screen was quite visible on the sample surface. In the center of a majority of the grid squares a brown spot had been formed. Under

microscopic examination the central brown spots do not appear clear

and dendritic patterns are quite visible around each of these spots.

These dendritic patterns are similar to the well known Lichtenberg

figures created by electrical discharge in insulators. In fact, this is precisely what they are. Bridges or conducting paths exist across the

gaps left by the grid lines. It is indisputable that these paths have been

created by electrical discharge.

Sample number four (4) was covered with a coarse mesh screen and it and the sample holder were biased at +300 volts during the ir- radiation. The sample was irradiated with 30KeV electrons at a rate

of 1x10^^ e/cm^ -sec. to a total dose of 1x10^^ e/cm . No centrally

located brown spots were formed this time within each grid pattern; in

fact, it appeared that little or no damage exists near the center of each

grid pattern. Damage does appear to be most severe near the grid

lines themselves. Dendritic pattern or electrical discharge paths exist

aroimd most grid markings: These are very large in some cases and

bridges exist across the grid lines. Sample number five (5) was irradiated with 25 KeV electrons at a

rate of 1x10^^ e/cm^ -sec. to a total dose of 1x10^^ e/cm . A small

barely visible brown spot exists in the center of the sample and the sample appears cloudy overall, but clear. Sample number six (6) was covered by a coarse mesh screen placed

2-3 mm in front of but not touching it. The sample holder and the screen were grounded. The sample was irradiated with 25 KeV electrons at a

rate of 3xl0^^e/cm^ -sec. to a total dose of 1x10^^ e/cm . The entire sample now showed an overall browning. The grid outlines of the screen were visible but were a darker shade of brown than were the central portions of each grid square. Surface scratch marks developed

dendritic growth patterns extending outward and normal to the scratch

marks; these dendritic patterns were not present on the same scratch

marks wherever they extended into a portion of the sample which was

not irradiated. The structure stopped abruptly at the boundary between

irradiated and unirradiated parts of the sample . This structure looks as if electrical discharge occurred from the teflon itself to the surface

scratches which acted as a conducting path to ground. We know that there is visible arcking of the sample during irradiation because a view

port was installed during a portion of these side experiments so that the samples could be observed during irradiation. This arcking coincided

354
with a jump in the instantaneous current being collected by the Faraday

cup and measured. At intervals of from five to thirty seconds (depending

upon the irradiation rate) the current measuring indicator would jump

suddenly to a higher value, perhaps fluctuate a bit for a second or two

and then settle down to its steady state value. We have no alternative

but to conclude that electrons are being stored in the teflon. Some

continually leak off, to be sure, to be captured by the Faraday cup, and

much of this is probably via surface conduction. However, enough

charge is apparently stored to create local electric fields of enormous

values; values so large that discharge is enevitable. Sufficient repetition of this gives rise to the dendritic patterns or Lichtenberg figures

in the teflon, and may be causing the browning of the sample center. Presumably, at low enough flux rates, the leak current would be sufficiently high to prevent significant charge accumulation. Visual observation of a decreased amount of arcking at lower flux rates confirms

this. If, however, this leak current is predominately due to surface

conduction caused by vacuum system contaminates it may not occur in space unless the contaminates found there due to the outgasing products

of the spacecraft itself are sufficient. On the other hand, space bombardment is predominately neutral beam outside the Van Allen belts and no localized fields may ever materialize.

The results of the proton and simultaneous proton-ultraviolet ir- radiations are presented in Table 2. These results tend to confirm the presumption that damage is en- ergy dependent, certainly for low proton energies. The 25KeV data seem to indicate an anti-synergistic effect; i.e. , that simultaneous

ultraviolet and protons tends to retard the damage that occurs with

protons alone. Certainly, no positive synergistic effect is evident.

However, the 5 KeV data do not support such a conclusion. With the limited statistical sample herein represented it is impossible to say. We hope to have more to say upon completion of the work.

355
TABLE 1 WAVELENGTH RANGE (nm) DETECTOR LAMP

250 - 360 Lead Sulfide Xenon

350 - 750 Photomultiplier Tungsten

750 - 2500 Photomultiplier Tungsten

TABLE 2 Integrated

Solar Absorption

Measured

in Vacuum

Sample #24

25 KeV

protons

Sample #25

25 KeV

protons

Sample #26

5 KeV

protons

Sample #27

5 KeV

protons

Type 5 mil Silver 5 mil Silver 5 mil Silver 5 mil Silver

Initial 0.074 0.079 0.071 0.072

IxlO^^p/cm^ 0.079 X 0.072 X

IxlO^^p/cm^

plus uv

(13.9ESH) x 0.079 X 0.071

5xl0^^p/cm^ 0.083 X 0.075 X

5xl0^^p/cm^

plus uv

(19.4ESH)

X 0.081 X 0.075

IxlO^^p/cm^ 0.092 X 0.079 X

IxlO^^p/cm^

plus uv X 0.095 X 0.082

(26.4ESH)

356
QUARTZ LENSES VACUUM PUMP

PROTON & ELECTRON

SOURCE

^^GATE VALVE

SCANNER AND ELECTRONICS

SPECTROPHOTO- METER

SAMPLE

(Lowered for Measurement) INTEGRATING

SPHERE

Fig, 1—Test facility for in situ measurement

bl°lll°lb

1

Fig. 2—Beam scanner installed in facility 357
\

358
Paper No. 25

THE INTERACTION OF LOW ENERGY ELECTRONS WITH

POLYMERIC PERFLUORINATED ETHYLENEPROPYLENE (FEP)'

John B. Schutt ^

REFERENCE: Schutt, John B., "The Interaction of Low Energy Electrons With Polymeric Perfluorinated Ethylenepropylene (FEP),"

ASTM/IES/AIAA Space Simulation Conference, 14-16 September 1970. ABSTRACT: Photomicrographs are presented showing the effects of 30 Kev electrons on FEP. Depicted are Lichtenberg patterns, bubbles

and a buckling effect. The attempt to explain these observations is derived from a theoretical analysis based on a diffusion mechanism

for charge flow using beam current as a source term. The solution is used to demonstrate that charge is stored centresymmetrically with

respect to ground. These effects are discussed in terms of an

"electrolysis" mechanism resulting from charge storage and beam

energy thermalization. Explicit expressions are given for surface

charge storage and voltage across the sample resulting from such

charge accumulation. Buckling is treated from a thermodynamic

basis using the Helmholtz free energy and assuming the distortion develops isothermally. "Electrolysis" is postulated via a mechanism

selecting the tertiary carbon as the active site for mobile moiety

generation.

The usefullness of optical data for purposes of thermal design

taken from samples irradiated with low energy electrons in the laboratory is questioned. Because FEP is an electrical insulator, these

experiments must, therefore, be carried out in a neutral laboratory

environment, or be characterized in the space environment from

carefully designed electrical experiments.

KEY WORDS: degradation mechanising, electron bombardment,

electron degredation, polymeric perfluorinated ethylenepropylene. proton bombardment, thermal control coatings

^Goddard Space Flight Center, Greenbelt, Maryland 20771

The numbers in parentheses refer to the list of references ap- pended to this paper.

359
Introduction

Few spacecraft thermal control materials have received as much

undocumented attention as perfluorinated polyethylenepropylene (FEP).

This unusual ascendency of FEP into prominence occurred because en- couraging thermal information obtained from flight data*^^-^) were

available prior to laboratory tests qualified from optical measurements.

That the performance of metallized FEP as a second surface

mirror appreciably surpasses white paints cannot be doubted-/ yet,

its qualification in the laboratory cannot be judged as altogether routine. For example, irradiating with ultraviolet photons or 2 - 30 Kev

protons gives reproducible data under carefully controlled conditions,

whereas irradiations carried out with electrons characterized similarly to their proton coimterpart will not give a degradation sxiffi- ciently uniform for optical measurement » This treatment provides an

elucidation of this "unconventional" behavior of FEP toward electrons.

In the following section, experimental observations of electron

irradiated FEP are presented and discussed with the detailed objective of providing a basis for a theoretical development of the charge

storage function c(r, z, t). Subsequently, the stability problem is discussed in a heuristic manner using some notions of control theory and

simple electric circuits. A brief thermodynamic description is presented and a solution to the complete problem is outlined accounting

for chemical reactions or "electrolysis" of the film. Finally, recommendations are put forth regarding electron stability testing of electric insulators.

Experimental Observations

Perfluorinated polyehtylenepropylene copolymer units comprise

the FEP<^^) network in a manner yielding a material susceptible to sublimation. That FEP is indeed subject to vaporization, resulting

from thermalization of a 30 Kev proton beam, is shown in Figure 1. The area surrounding the puncture contains proton beam generated

charge rearrangements detectable by optical means as additional ultraviolet absorption sites having the usual long wavelength cutoff extending deep enough into the visible portion of the spectrimi to create

a brown color. The puncture occurred near the center of unmetallized

material or at the point furthest from ground. The experiment was

(2,3,4,5)r]njjg numbers in parentheses refer to the list of references ap- pended to this paper.

360
repeated using a grid placed about 30 mils above silver coated 5 mil

FEP. Experiments were performed with both grid and sample holder

at groimd potential and with the grid at +300 volts above ground. Results were similar and reproducible for both cases. Figure 2 shows

the configuration. Figure 3 is a photograph of the area comprising

one section of the grid.

Although not clearly shown in Figure 3, the mesh centered portions of the FEP generate ultraviolet centered optical changes possessing absorption tails extending far enough into the visible portion

of the spectrum to show the familiar brown color. Again, heating ap- peared to take place in portions furthest below ground electrically.

Clearly visible are dentrite-like Lichtenberg pattems*^^^ emanating

from the center of the sample and extending toward portions shadowed

from the electron beam by the grid. Figure 4 reveals the existence of these dentritic patterns amid clusters of bubbles. It appears likely

therefore that charge is stored at least in the neighborhood of the surface of the material, preferably at its geometric center relative to the mesh. Referring again to Figure 3, it is apparent that the surface of the sample is not in relative focus owing to the buckled central region

where charge storage and thermalization occur. Figure 5 shows this effect schematically. Notice that the metallic film maintained its adhesion to the FEP.

In summary then, it appears reasonable to assume that 5 mil FEP

whether metallized or not can at least store charge in the neighborhood of its surface and in a region furthest from ground. The conversion of beam kinetic energy into thermal energy in the presence of stored charge gives rise to an "electrolysis" or bubbling of the material and, therefore, the formation of dentrites capable of carrying

charge to ground; ground itself being made more accessible via the grid by means of a buckling mechanism. Figure 6 depicts the veined

network common to virtually all beam shadowed areas where charge

also arcs to ground. It is the charge storage problem and its relationship to the grounding of the sample that forms the basis for the fol- lowing section.

The Charge Storage Problem

The phenomenology of electron storage in electrical insulators

and dielectrics can be treated in a manner formally identical to a

^^^The numbers in parentheses refer to the list of references ap- pended to this paper.

361
corresponding thermal problem. Our starting point is, therefore, the

diffusion equation:

3 cCr, z, t")

^ = DV2c(r,z,t) + VF(z) (1)

where:

, c(r, z, t) = nimiber of charges at the point (r, z, t)

t = time, D = a diffusion constant and,

(r, z) = an arbitrary point in cylindrical space. Because we take for simplicity that charge is incident normal to the surface at z = 0, the angular dependency is automatically eliminated

from the equation. The function F(z) or source term is the solution of dF(z)

- "dT" " kNoF(z), (2)

the usual penetration equation governing spatial dependencies norma

to the incident particles. The parameter k represents a number

characterizing the capture cross-section of the storage site, while Nq

is the site density. Solving (2) and substituting into (1) gives

= DV^c - NgkEoe'^^o' = DV2c-f(z) (3)

where is proportional incident particle momentum. We obtain the

solution of (3) as the simi of two functions,^''

c(r,z,t) = u(r, z) + w(r, z, t) , (4)

with w = 0 when z = 0. It follows that

B2 1 3 32u f(z)

B r (5)

(^^The nxmibers in parentheses refer to the list of references ap- pended to this paper.

362
and

iBw l3w

+ —^ +

A solution to (5) is sought in the form

^ " I] <^^(r)V^.(z) (7)

with 0- (r) forming an orthonormal set. Substitution of (7) into (5) shows that the 0. 's are Bessel functions of zero order with i//. being

a solution of

^'^i (^.\' 1 Ji(^i) 2 f(z)

where a. is a zero of ]q{/j.R), since the sample is assumed at ground

potential for r = R. The parameter /x is a separation constant, Y^ijsr) was eliminated as a solution, since c(0,z, t) must be finite. The

Greens function solution of (8) was obtained using the usual symmetric

boundary conditions,

G<(0) = 0 and G>(d) = 0, (9)

where d is the thickness of the sample. The result is

G = 0i^-z) .^l y sinh^tz + d(z-^) , sinh/x(d-z)(10)

= G<(^^z) + G>(z,0.

with 6 representing the Heaviside step function. The particular solution to (8) is then given by

V^p(z) = f 'g<(^,z) g(a d^ +1 ^G>(z,^) g(^) d^'

363
M sin h/xz

2/x sin h/xd (M+kNj (kN,-^)

1-e (kNo-M)z M sin h/x (d-z) f 1 2/xsinh/xd L(^Nq-/x)_

^-(kNo-M)z _ ^-(kNo-yx)d

1

(/x + kN^)

^-(M+ kNo)z _

^-(/x + kNo)d

(11)

where

M =

1 J,(a.) 2FoN,k

The homogeneous solution is of the form

<//„ = A e^^ + B e'

with the total solution written as

(12)

or

Ji(^i)

DkN,

1 +

g-/^d sin hfxd

sin h/x z

(13)

where is given by equation (11) . The complete solution is now re- duced to the form

c(r,z,t) = E ( + V^pJ^i + (14)

3t4
where the solution for w remains to be determined.

If we choose

w(r,z,t) = Wj(r) w^Cz) W3(t) (15)

then

^1 = AJo(/^0 + B,Yo(/^r)

and = 0 since must be finite for r = 0, and as before (^R) =

0 to satisfy our grounding condition. Therefore,

with 0 < X < 1 in place of 0 < r < R. Continuing the partitioning to obtain equations for and w^, we obtain

32 w_

2 3 z2

+ X^w. = 0 (16)

and

1 ^^3

If we write

W2 - A2 sinX.z + B2 cos \z

and

W3 = A3 e-(^'^^')°^

equation (15) becomes

365
A' sin ^ z + B' cos k z (18)

The constant is determined from the groimding requirement that w = 0 when z = d for 0 < x < 1 (0 < r < R) giving = 0. The

parameter X. is determined from the zeros of sin \ d or

k = 27Tn/d

aad

(r,z,t) = Z L ^

2 , a. . 2 fj„(a.k) sin (2"" 2n77-j- (19) 7)'

X n

Thus far we have demonstrated that our charge storage solution can be put into the form

c(r, z, t)

^ ^A. e^^ + B.^ e"^^^ e-^^l + Jo(a,k)

(20)

2n77 z\

1 n

To determine A. , B., and A 2 we apply our grounding condition

c(r,d, t) = 0, 0<r<R, 0<t<oo (21)

followed by the surface requirement

c(r,0,t) = Fq/UNoD, 0<r<K, 0<t< (22)

and finally the initial condition that the sample carries no initial charge, or

366
c(r,z,0) = 0. 0<r<R, 0<z<d (23) Applying the boundary conditions (21), (22), and (23) in the order

listed gives

A. =

B. = - A. e^^^^

1 1

and

2FoJi(-i) 1 kN

(2nn) R

(24)

with A. being determined from a Fourier-Bessel's expansion and A^. from both Fourier and Fourier-Bessel's expansions.

Discussion of the Relationship Between Charge Storage and Stability

There exist two conduction paths for the escape of absorbed

charge to ground: along the surface and through the bulk. Brief

consideration determines that these routes can be represented most

simply as two resistances connected in parallel. Let us write expressions for local resistances in the form

p(d - z) _ _ p(R - r) ^buik ~ K 2 andR^^^f^^^ - R^ rz

where the locality is given by (r, z) for a sample of size (R, d) and p

is the resistivity. Combining these expressions in parallel, the local equivalent resistance, R(r, z) becomes

Rrrz^ - P (d-z)(R-r)

^^"'^^

77r 2z(d - z) + r(R - r)

^^^^

367
Taking the limit for z we obtain

R(r,0) - R = limz->0

while if in addition the limit for r approaching zero is taken,

R(0,0) = R(r,0) = 00, lim r - 0

a divergent expression results/^ ^ In both cases the local resistance

about the origin gives a divergent result, suggesting that the closer an

electron is to the origin the more difficult its migration to ground be- comes. Hence the conjecture that charge is stored in the center of the sample or farthest from ground appears to be a reasonable one. Using this approach, however, the divergence of R(0,0) cannot be re- moved; but, having anticipated this result the diffusion equation was

solved in the previous section to remove this divergence. From equations (14) and (15), the storage function is most compactly written in the form

(26) c(r,z,t) = 2] 0(^x){0h(z) +0p(z) . + ^

2Z ^^^i'')^2(2)^3(t).

where use has been made of the identity 4>{p--^ x) = Wj(a. x). Our

boundary condition on the radial solution to equation (15), demanding

the convergence of c(r,z,t) at c(r,0,t) and c(0,0,t), eliminated the divergent Neumann function. Inspection of equation (11) for 0p , equation (13) for s^H, and the result Jo{0) = 1» gives an expression for the number of charges stored in the neighborhood of the origin:

c(0,0,t) = ^h(^) +

(^^The numbers in parentheses refer to the list of references ap- pended to this paper.

368
2 Ji(^i) ^_o_

(27)

. ^-(kNo-/x)dj ^ M 1 _ ^-(kNo+M)dj 1 2a^'

with M being given in equation (11), By inspecting the above result the generalized surface charge

storage expression is obtained:

c(r,0,t) = ]2 c(0,0,t) J^Ca^x). (28)

i Equation (26) shows that the charge stored at (0,0) is indeed finite and

that

F R(r,0) = e c(r,0) C (29)

where

e = electronic charge,

C = surface capacitance at (r , 0) , and

F = electron current. Equations (27) and (28) show that FR(0,0) = V(0,0) is finite. However, R ^ p = 1/a

where cr is the conductivity. Writing

a = c(0,0, t) e fx^

where e and /^^ represent electronic charge and mobility, respectively, we can conclude that R(0,0) is finite. Next we wish to calculate

the number charges stored on the surface of the sample. The result is obtained by integrating c(r, 0, t) over the surface of the sample:

369
c

s c(*,0,t) Jo(a,. X) r d r de

J,(ai)

a. (30)

i

The number of charges per unit area is then given by

Ji(«i)

a. (31)

Since the voltage supported by a capacitor is given by definition as

where q and C are charge and capacitance, respectively, we auto- matically have an expression for the voltage across any undamaged

sample.

a quantity subject to calculation and measurement.

At the on set of this section, we discussed the conduction paths as

a parallel arrangement of resistances, purposefully neglecting their

respective path capacitances. It is clear from the photographs discussed in the first section, that material instabilities can be generated

under quite moderate irradiation conditions. Further, it has been

suggested that thermalization and charge storage cooperatively induce

material changes. In a heuristic manner, then, we suggest that the

initial charge storage sites become inductively altered once saturation

has been reached, and propose a black box scheme to model the course

of events. Shown in Figure 7 is a simple feedback loop represented by

the transfer function (p) or a time delayed degradation mechanism

serving to modify the voltage generated across the circuit with transfer function (p), the initial charge storage mechanism. The circuit

transfer function (p) is given by

V = qC, (32)

V = ec^ C, (33)

370
Yo(P) = Yi(P)

1 - Y^(p) Y^Cp) ' (34)

where the stability conditions are related to the voltage discharge fre- quency (p) to ground via the grid. It follows that stability will be determined by the zeros and poles of the function

Y^(p) Y^Cp) - 1

The transfer functions are accordingly generated from the circuit

analogy selected to represent the physical situation. This aspect of the problem will not be carried further here, but rather reserved for a future publication.

Before concluding this section, by reviewing the problem from a mechanisms point of view, the buckling phenomenon will be briefly

discussed using a thermodynamic approach. As a starting point we

assume that the occurrence of buckling could take place isothermally,

that is.

B c(r, z,oo) 1 /V

(35)

where

F = Helmholtz free energy,

py = vapor pressure of the FEP,

/X = chemical potential,

c (r,z,oo) = terminal nimaber of stored charges (t = co), V = volume, and

e = dielectric constant. Next we make the substitution for the dielectric constant

1 + 47rP (36)

and take

371
to obtain

where P is the polarization. Equation (37) expresses the charge buckling effect in terms of the vapor pressure p of the expanded surface.

To rewrite the equation in terms of the vapor pressure Pa, of the un- deformed material and the surface energy, use is made of the relationship between surface pressure and curvature:

, Pv = (38)

where y is the surface energy and and R2 the principal radii of curvature. Our final result is

1 11 / V\2 ^ ^ V

T/^riTj (39)

If FEP is "electrolyzed" under the influence of an electron beam

and an expansion or buckling results, the surface curvature effect enters in a way such that the vapor pressure of the material will tend

to be maintained at its original level. To simplify the argument we

take = 0 and Rj = R2 or

3 c

when

27 PV /VV

R7>2.-d-nTj

Pv

B c(r, z,oo)

3 c(r, z,co) ^

_27

R.

27tP- (40)

372
Be

Then ^— < 0, and the tendency toward volume charge storage is diminished and y becomes a factor in the behavior. On expansion d

will increase allowing for some increase in P and V if field strength

is maintained. The relationship of the surface effect to bulk storage

suggests buckling and charge escape to the grid is manifested by an

"electrolysis" effect. We are now prepared to relate the stability

problem to the chemical nature of FEP. Quite generally we can write

the copolymer in the form

_^CF-CF— C CF,^ (41)

assuming the ethylene and propylene units are combined in equimolar

amounts. If FEP stores appreciable charge prior to damage, then a small degree of imsaturation may be present. Next the tertiary carbon

becomes suspect as a source of charge imbalance by virtue of the known stability of the carbanion. We are led to assume

CF,

+ e + cf: (42)

CF.

+ 2( + cf: (43)

and similarly for extraction of the fluoride moiety. A "crystallization" or cross linking reaction can be written for tertiary radicals,

2—

(•-)

(2e)

(44)

373
with the products CF3^~) and F^") becoming suitable candidates for transporting charge to the grid^^\ where discharge or the "electrolysis" effect is formally completed:

2F" F2 + 2e (45)

(CF3), + 2e. (46)

Summary and Conclusions

Throughout this paper the assumption of beam uniformity and

stability was maintained. That electron beam stability can" be a problem is well known by experimenters in this field. However, it is felt that the paths to ground are the most critical aspect in the work,

otherwise the observation of buckling might not have been made.

Clearly, the grounding paths and physical effects in the material are in turn related to the charge storage and beam energy thermalization

effects. Vaporization is probably related to the proclivity of hydrogen

deficient organic materials containing side groups with small steric

factors toward sublimation, (this behavior is related to the well known

flammability problem). FEP is a highly charge compensated material

by virtue of the orientation of its bond dipoles. That these orientations are easily disturbed is evidenced by the well known T~ ^ dependency .(^o) Combined with the charge storage property a temperature effect be- comes reasonable.

Examining the buckling effect once more, the dominance of surface

storage over bulk storage is suggested, since this distortion moved the surface of the sample closer to the grid or suspended ground as op- posed to the bulk ground through the silver to the substrate. Heating

combined with charge repulsion can then be used to justify the expansion. Once distorted, charged particle diffusion channels may be

formed and the "electrolysis" effect may proceed. The observed

browning effect follows as the result of FEP moiety reorientation and

"electrolysis".

The electron irradiation of electrical insulators cannot be treated

from an experimental approach in a routine manner. A spacecraft

passes through space carrying a neutral charge cloud defined with re- spect to some finite volume, and departure from electrical neutrality

(9.10) rpjjg numbers in parentheses refer to the list of references ap- pended to this paper.

374
over the surface of this cloud likely exists only as minor electrical

fluctuations. Consequently, it appears reasonable to conjecture that for laboratory test data on electron irradiation to be meaningful, data

should most probably be obtained using a neutral beam environment.

If laboratory simulation of low energy electron and neutral charge

distribution fluctuation effects develop into a formidable task, then

data from carefully designed spacecraft experiments provide an at- tractive alternative for t^vo reasons: first, meaningful thermal design

data are obtained and second, realistic inputs are obtained for materials development.

Acknowledgment

The author wishes to express his appreciation to Messers. Rebne

Thompson, Segun Park and Bruce Bean of Electro -Mechanical Research Incorporated of College Park, Maryland, for helpful discussions. Their elaboration of experimental observations not readily

recorded photographically were instrumental in establishing the scope

of this paper.

References

1. The experimental portion of this work was supported by funds

supplied by the OAO Project Office, Goddard Space Flight Center,

Greenbelt, Maryland. Data were obtained under Contract No,

NAS 5-9469 with the Electro-Mechanical Research Corporation,

College Park, Maryland,

2. Carrol, W. F., "Mariner V Temperature Control Reference Design, Test and Performance", AIAA paper 68-791 presented

AIAA 3rd Thermophysics Conference, Los Angeles, June 1968.

3. Private communication from Miss Archie Fitzkee, Thermal Control Section, Thermophysics Branch, of OGO-6 Flight Data on

Silvered 5-MIL FEP.

4. Private commimication from i\Iiss Archie Fitzkee, Thermal Control Section, Thermophysics Branch, of OGO-6 Flight Data on

Silicated Treated Zinc Oxide in IMethyl Silicone.

5. Usually referred to as FEP Teflon Type A, Teflon is a Registered trademark of E. I. duPont de Nemours Company.

6. Gross, B. and Murphy, P. V., "Electrical Radiation Effects in Solid Dielectrics," "Nukleonik", 2. Band 7. Heft, 1961, S, 279-285,

375
7. Carslaw, H. S. and Jaeger, J. C, "Conduction of Heat in Solids",

Second Edition, Clarendon Press (Oxford), London, 1959, pp. 29-

30.

1 3 / dp\

8. This difficulty may be overcome if — ^ k lexists, a result suggesting solution via a diffusion mechanism.

9. Lilly, A. C. and McDowell, J. R., "High- Field Conduction in Films

of Mylar and Teflon", Journal of Applied Physics, Vol. 39, 1968,

pp. 141-147.

10. Debye, P. D., "Polar Molecules", Rheinholt Press, New York,

1929, pp. 27-30.

376
377
BEAM

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380
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Figure 7. Simple Feed Back Loop Representing Effect of Material Changes on Intrinsic Charge Storage Capability

382
Paper No. 26

THE DEGRADATION OF ALZAK BY SHORT WAVELENGTH

ULTRAVIOLET RADIATION

M. J. Donohoe, Roy Mcintosh, Jr. and John H, Henniger, Thermophysics Branch, NASA Goddard Space Flight Center, Greenbelt,

Maryland.

REFERENCE: Donohoe, M. J., Mcintosh, Roy and Henniger, John H., "The Degradation of Alzak by Short Wavelength Ultraviolet Radiation",

ASTM/IES/AIAA Space Simulation Conference, 14-16 September 1970

ABSTRACT: Experimental evidence is presented to further support

earlier findings on the UV induced degradation of Alzak, which was

found to be inversely proportional to wavelength between 220nm and

SOOnm/^^ UV and vacuum UV exposures at selected wavelengths between 123. 6nm and 300nm indicate degradation to be inversely proportional to wavelength down to 185nm. Equivalent exposure to 123, 6nm

radiation however, produced less degradation indicating a change in the damage dependency below 185nm. Degradation of up to 12% change

in reflectance at 295nm is reported for Alzak exposed to less than 50

equivalent sim hours of vacuum UV radiation.

KEY WORDS: ultraviolet degradation, vacuum UV, thermal control

coatings, Alzak,

There has recently been increasing interest in the damage of thermal control coatings produced by short wavelength ultraviolet

radiation. This has occurred, in part, as a result of an apparent discrepancy in data returned from a coatings experiment aboard ATS-I.^

In addition, certain other findings indicate that, at least for some

Numbers in parentheses refer to the list of references appended to this paper.

383
coatings, degradation appears to be inversely proportional to the wavelength or irradiance/^^ Since the solar spectnmi is rich in the vacuum and extreme ultraviolet regions, some experimenters have

postulated that this could be a matter of great importance/"^ ^ If such

a hypothesis could be demonstrated, it would mean that one more fac- tor should be added to the growing list of requirements for sjniergistic

ground based testing of spacecraft coatings. At the Gk)ddard Space Flight Center work has been underway for some time in an attempt to fully characterize the ultraviolet degradation of one of the important thermal coatings, namely, Alzak. Alzak is the designation given to a particular type of anodized and chemically

polished aluminum sheet which is manufactured by the Aluminum Company of America. This material was chosen for the outer skin of the OAO spacecraft because it can be easily worked by ordinary sheet

metal techniques and because it has an attractively low solar absorptance (a^ )/hemispherical emittance (e^) ratio of approximately 0.22. Work has been previously reported which showed that the degradation

of alzak was strongly dependent upon the wavelength or irradiance at least down to = 220nm.^2^ The experiments which are described here were designed to provide data down to the solar Lyman a region

(121. 6nm).

EXPERIMENTAL .

. The experimental apparatus used in this study is shown in Figure 1.

It consists of an irradiation cell and three UV sources with wavelengths at 123. 6nm, 185nm, 206. 2nm and 240nm. Not shown is a DK-2

Spectroreflectometer with an Edwards type integrating sphere used for

in-air reflectance measurements on the samples. A 1-1/2 liter stain- less steel vacuum chamber with a side port forms the irradiation cell. The cell is ion pumped with a trapped mechanical roughing pump. All vacuimi seals are the copper gasket type and the valves have viton

elastomer seats. Irradiations take place in the 10"^ torr region as measured by an ionization gauge. The irradiation port is a 2" diameter sapphire window or a 1" diameter lithium fluoride window, depending upon the wavelength or irradiation.

Tlie sample was screw mounted on a temperature controlled copper substrate opposite the irradiation port. Sample reflectance

measurements were made before, during and after the irradiations by

^Thermoelectric module — Cambion Corp., Cambridge, Mass.

Numbers in parentheses refer to the list of references appended to this paper.

384
removing the sample from the chamber. The sample could be re- moved, replaced and pumped down within 30 minutes. No detectable

recovery in alzak took place during this time interval. During the earlier studies, in-situ reflectance measurements of Alzak damaged

by irradiations from a 123.6nm source and a xenon arc indicated that recovery of damage was negligible when measurements were made

within 8 hours after exposure to air. Reflectance measurements were

m_ade with an estimated error of ±l/2^c due to variations across the sample face. The UV irradiation sources were two low pressure microwave

discharge lamps (krypton, iodine) and a high pressure xenon arc lamp^

with a band pass filter^ centered at 240nm. The relative spectral out- put of the three sources is shown in Figure 2, together with a solar UV

irradiance curve for reference.

The krypton lamp emits three lines, llTnm, 123. 6nm, and 247nm.

The 117nm line is weak, being approximately 10% of the 123. 6nm line intensity. The 123. 6nm line is the primarj^ irradiating wavelength and

is used to simulate the solar Lyman a line at 121. 6nm. The 2nm dif- ference between the t^vo lines is assumed to have negligible effect on

the results. The 247nm line contribution to sample degradation is evaluated by masking half the sample with a sapphire filter to eliminate the 123. 6nm line, but pass the 247nm line. A reflectance measurement on both sample halves then isolates the 247nm line contribution.

The krypton lamp output spectrum is blank from 247nm to approximately 330nm. The lamp is normally run gettered with a cold finger

immersed in liquid nitrogen to eliminate contaminant lines. Krypton

lamp intensities could be varied by the microwave generator from a

fraction of an equivalent Lyman a sun (i.e., one equivalent Lyman a sun 5 X 10"'^ watts/ cm 2) to approximately 20 equivalent Lyman a suns. The LiF wiadow life was greatly reduced at the higher intensities due

to the window sealing technique used.

The iodiae lamp has a sapphire window and emits a group of lines centered at 185nm and a strong line at 206. 2nm. The lamp is normally

run with a cold finger immersed in water ice to maintain the proper

pressure in the iodine reservoir shown in Figure 1. The lamp spectrum above 206.2nm is blank up to 400nm. Differentiation between the degrading contribution of the 185nm lines and the 206. 2nm line is achieved using a gas filter of butene-l.*^"^^ The butene-1 cell absorbs

the 185nm lines while reducing the 206. 2nm line intensity by only 10%.

Hlodel XR-150, Eimac Corp., Palo Alto, California

3 Optics Technology, Palo Alto, California

Numbers in parentheses refer to the list of references appended to this paper.

385
The iodine lamp output was variable up to 1/2 an equivalent sxm at 185nm and 206.2nni (i.e., 40 Lyman a suns). The 185nm lines contributed approximately one half of the total, the balance was the 206.2nm line intensity.

The xenon arc lamp used, had a continuum at 240nm and the filter centered at 240nm had a band pass of 30nm. Maximum output was

approximately equal to 1/2 an equivalent sun at 240nm (i.e., 75 Lyman

a suns).

The intensity monitor for all the lamps was a sodium salicylate

coated photomultiplier (1P28) referenced against a calibrated nitric oxide ionization chamber using the 123. 6nm krypton line. The photomultiplier could be rotated to a position between the sample and the irradiating source. The calibration measurements were performed

with the ionization chamber in the sample position as shown in Figure 1. The sodium salicylate quantum efficiency was assumed constant over

the wavelength region of interest. This assumption is supported by the

literature.(^) The photomultiplier calibration was checked periodically during the experiments and found to remain within 5% of the

original calibration.

Calculation of the equivalent sun irradiance for each source was

based on Johnson's data using a lOnm band width of each line. The

185nm lines were treated as one line between 180nm and 190nm.

Actually the band width of each line was known from the vacuvim

monochrometer measurements. Use of the actual band width, i.e., 2.5nm just increases the equivalent sun hour exposures. Since line sources were used to simulate the sun's continuum spectrum at select wavelengths, the lOrnn interval was used. Another convenient way of looking at the data would be in terms of Lyman a suns since part of the experiment is aimed at studying the effect of Lyman a.

RESULTS

Plotted in Figure 3 are the changes in reflectance at 295nm of Alzak samples exposed to different irradiating wavelengths. Exposure

to 123. 6nm light used to simulate solar Lyman a produces the slowly

developing damage shown in the lowest curve. Exposure to combined

185nm - 246. 2nm light produces the steep curve at the left. This curve was constructed from 3 samples as indicated by the legend.

Shown in Figure 4 is a plot of % AR at 295nm versus wavelength

for constant incident energy.

Numbers in parentheses refer to the list of references appended to this paper.

386
Before discussing the data in Figure 3 and Figure

4, two additional experiments should be mentioned.

The first is a single Alzak exposure to 123.6nra

irradiation at an intensity equivalent to the 185nm206. 2nm irradiations shown in Figure 3. For a 36-1/2

exposure at 1.4 x 10" watts/cm^ (i.e., = 30 times

Lyman-Q^) . Alzak showed a 4% AR at 295nm compared to

the 11% shown in Figure 3. The second result is a UV screening test on Alzak using a filtered 2.5kw xenon arc lamp (Spectrolab Solar Simulator model X-25)

The UV irradiation tests performed through quartz vacuum chamber ports produced a 12% AR in Alzak after

57 equivalent sunhours at a one sun rate. Therefore,

this sample saw the complete solar spectrum starting

at the quartz cut off of the simulator and vacuum

system optics. The UV content (i.e., below 220nm)

for this test was not known, but the point is presented to emphasize the damaging effects of short wavelength UV. The energy output ratio in watts/cm^ of

the simulator versus the microwave discharge lamp is

approximately 10,000 to 1. DISCUSSION

The conclusions which can be drawn from the data

presented are two: first, on a solar irradiance basis

Alzak is damaged faster and further by wavelengths

in the 180 - 210nm region than by Lyman-o' (see Figure 3), and second, on an equivalent incident energy

basis, Lyman does less damage than the 180 - 210nm

wavelengths (see Figure 4). These two points, coupled with the previous studies on Alzak above 220nm,

provide a general behavior picture for Alzak degradation; i.e.. above approximately 300nm no degradation

is observed for long exposures (100 hours) , and below

300nm increasing degradation with decreasing wavelength is observed which peaks somewhere between 150-

210nm (on a solar irradiance basis) and reaches a lower value at Lyman-o^ wavelengths. In addition,

most of the damage caused by the 180-210nm wavelengths

occurs within the first 50 hours of exposure and may

approximate the damage observed under standard solar

irradiation tests.

387
From the standpoint of solar environmental testing of Alzak, it is evident that Lyman-cv wavelengths

need not be included in laboratory testingo Every

effort should be made, however, to include the wavelengths between 150 and 200nm. Other materials may behave differently and should at least be checked

for their behavior under Lyman-cy wavelengths

At this point in the studies it is only of

academic interest to determine the precise shape of

the Alzak damage curve because of the large reflectance changes observed using the iodine lamp„ However, efforts are being made to develop a bromine resonance lamp to provide a line source at 163nm

which will bridge the wavelength interval between the

iodine and krypton irradiation sources.,

REFERENCES

(1) Mcintosh, R., "The Degradation of Alzak as a Function of the Wavelength of Irradiance," Symposium on Coatings in Space, December 1969, Cincinnati.

(2) Reichard, P. J., Triolo, J., Progress in Astronautics and Aeronautics, Vol. 20, Academic Press, New York, 1967, 491.

(3) Swofford, D. C, Johnson, S. W., Mangold, V. L., "The Effects of Extreme Ultraviolet on the Optical Properties of Thermal Control

Coatings", AIAA Thermophysics Conference, June 1968, Los

Angeles, Paper No. 68-783.

(4) Liuti, G. and Mentall, J. E., "Monochromatic Iodine Lamp", Rev.

Sci. Instr., 39, #11, 1767, Nov. 1968.

(5) Samson, J. A. R., Techniques of Vacuum Ultraviolet Spectroscopy,

John Wiley and Sons, New York, 1967, p. 212.

388
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Paper No. 27

THE PHOTO-RESPONSE AT AN ALKALINE PAINT INTERFACE

WITH ALUMINUM

Walter A. Wappaus

REFERENCE: Wappaus, Walter A., "The Photo-Response at an

Alkaline Paint Interface with Aluminum," ASTM/IES/AIAA Space

Simulation Conference, 14-16 September 1970,

ABSTRACT: When a spacecraft coating comprised of aluminum oxide

and potassiimi silicate is applied to an aluminum alloy (6061) substrate

and subsequently partially coated with evaporated gold electrodes, a measurable potential is developed between the substrate and the gold

electrodes. This effect is observed when the coating is in darkness

and in vacuum. A dark current is also observed. Prior to observation, the specimen is heated in vacuum to 150°C for 24 hours.

When the coating is irradiated in situ with UV shorter in wavelength than 3600 Angstroms, the current and voltage is found to in- crease. Upon exposure to prolonged irradiation this current level decays exponentially with time. The voltage similarly decays with

time in an exponential manner indicating that photo- degradation of the electronic transport properties is occurring.

The transport mechanism proposed to account for this behavior is a migration of protons toward the aluminum substrate and negative

hydroxy1 ions toward the gold electrodes resulting in the electrolysis

of water in the coating. In the presence of UV this process is accelerated resulting in an increased current flow of three orders of magnitude. Since the current flow during UV radiation decays exponentially, this suggests a diffusion limited process.

j KEY WORDS: alkaline paint, conduction of paints during ultra-

{ violet exposure, electrode potential of paints , photo degradation, thermal control coatings, ultraviolet degradation

! ^Goddard Space Flight Center, Greenbelt, Maryland 20771

393
INTRODUCTION

The stability of spacecraft coatings has been the object of much

investigation. For successful satellite application a coating must exhibit stable properties for long periods of time when subjected to a space environment. One parameter, important to thermal control of a

satellite is the absorptivity, or a, of the spacecraft coating, a is an

optical property and usually changes when the coating is subjected to a space radiation environment. The object in coating formulation is to prepare a compound such that its time rate of change of a is very

small when subjected to space radiation. The coating studied in this investigation, a mixture of aluminum oxide and potassium silicate, is stable with respect to a in a proton and electron radiation environ- ment. However, it is unstable when subjected to ultraviolet irradiation. An attempt is made here to study the conduction characteristics

of this coating by monitoring some electrical parameters of this coating while being subjected to ultraviolet radiation. This paper dis- cusses an in- situ method for recording the behavior of a voltage across and a current through the thickness of a spacecraft coating

applied to an aluminum alloy (6061) substrate. This coating when

sprayed on an aluminum alloy substrate is representative of a typical

spacecraft skin whose surface is painted for thermal control purposes.

In this paper conduction schemes are suggested and a model for

ultraviolet degradation is proposed. As far as is known this particular

approach in studying the in-situ degradation of spacecraft coatings has

not been previously employed.

EXPERIMENTAL PROCEDURE

The coating which is discussed here is formulated from 1.0 micron alpha phase aluminum oxide ^ and potassium silicate^ mixed in a

four to one solid ratio; that is. Weight AI2O3

= 4 Weight K2O + Weight Si02

The mole ratio of Si02 to K2O is 3.3; i.e.,

K2O / mole

= 3.3

^Manufactured by the Linde Division, Union Carbide Corp.

^Sylvania Corporation's electronic grade PS-7.

394
The pH at this ratio is 12.8 using water as the solvent. The

preparation is sprayed on an aluminum alloy disc to a layer thickness

of about 0,1mm. WTien air dried this coating has good hiding properties as well as low absorptivity values.

xA. set of gold electrodes is evaporated on the surface of a coated

disc. The electrode configuration is shown in Figure 1. The disc is then moimted in a specimen holder which is designed to position a heater in contact with the imcoated side of the specimen. The heater

is electrically insulated from the aluminum disc. The specimen

holder firmly positions electrodes contacting the evaporated gold on

the coated surface. A small amount of silver paste serves to insure

electrical continuity between these contacts. The holder is mounted

in a vacuum chamber such that it is electrically insulated from the chamber. An ultraviolet transmitting quartz port is provided through

which radiation from a 500 watt medium pressure mercury lamp"^ is directed normally onto the coated surface. The two independent gold surface electrodes are electrically

joined and connections are led outside the vacuum chamber via an

electrical feed through. Another lead is attached to the aluminum

substrate disc of the coated specimen and led externally in a similar manner. These two leads permit electrical measurements to be made

across the coating layer.

The vacuimi is ion pumped and capable of reaching the 10"^ torr region when empty, A sorption pump is used for roughing. All metal

components are stainless steel. A light proof enclosure fitted with a

shutter is placed over the quartz port and optically aligned with the

ultraviolet beam. The specimen is in total darkness when not being

irradiated with ultraviolet light. The pumping system and the light source are shown in Figure 2. The radiation tests are designed to show the change of electrical

parameters of the coating specimen as a function of total irradiation

exposure time. At the start of the experiment the chamber containing

the specimen appropriately mounted in the holder is roughed out and

subsequently pumped down to the 10"^ torr region. The specimen is then heated to 150°C for a 24 hour interval and then allowed to cool. The pressure usually lowers to 10"' torr after this procedure. The

mercury ultraviolet source is positioned so that the specimen receives a one sim irradiance. The shutter is opened for a given time

interval. The current and voltage are measured prior to and at the very beginning of the radiation time interval. Both light and dark current and voltage measurements are recorded with an electrometer.^

^Hanovia model 6730A10

^Keithley Instrimients, Inc., ^lodel 610B

395
At the conclusion of the irradiation time interval the shutter is then

closed and a period of time is allowed to elapse before the next radiation interval in order to allow the specimen to return to initial temperature conditions. This cooling interval varies according to the length of time of radiation exposure, as the radiation interval is in- creased the cooling interval is also increased,

RESULTS

It is observed that, with the coating in darkness, a current flows

through the paint compound. The sense of this dark cxirrent in the external measuring circuit is from the aluminimi substrate electrode to the evaporated gold electrode. The dark current, when measured just

prior to each radiation interval, was in the 10" ampere region for the duration of the experiment. A dark voltage is also observed and is developed across the two different electrodes. The gold evaporated

electrode is of positive polarity. The dark voltage when measured

just prior to the start of the experiment was found to be 0.35 volts. This dark voltage decreased linearly with time to a final value of 0.15 volts at the conclusion of the testing.

The results of the photo -response are presented in graphic form

in Figures 3 and 4. Figure 3 shows the relationship between the current flowing through the coating, while being irradiated with ultraviolet light, versus the total accumulated time of this radiation. The

time as plotted is the cumulative sum of the radiation intervals. Similarly, Figure 4 shows the relationship between the voltage developed across the coating electrodes, while being irradiated with ultraviolet

light, versus the total accumulated time of radiation.

When a blocking filter was put into the light path, filtering out radiation of wavelength less than 3600 A, no photo -response was detected. The original color of the coating was a chalk white, but at the conclusion of the irradiation the color had changed to a light brown.

Upon re-exposure to light and a normal laboratory atmosphere for

several months, the color of the coating did not bleach but remained

brown.

DISCUSSION - The coating material in contact with the aluminum substrate exhibits properties similar to a dry cell battery with a power output less than 10"^^ watts. The voltage which is built up and maintained across

the electrodes seems to occur spontaneously without the need for the application of an external voltage to initially polarize the system.

396
Contact potentials are certainly contributing to the voltage build up but

their contributions are small compared to the total voltage in the dark

condition.

The mechanism for current production in the coating is different

during the dark and illuminated conditions. La order to discuss these

conduction schemes it is necessary to describe the various anodic and

cathodic reactions occurring under these different conditions. The

following reactions are suggested to account for the conduction mechanism.

Dark Conduction

In considering a reaction at the cathode, one that is immediately

suggested is, H^O

Al + X KOH > Al (OK)^ + -2 H2 T , 0 < X < 3 (1)

due to the fact that the coating is strongly alkaline. Guided by (1), with

the release of hydrogen at the cathode, the following competing cathodic reaction is also suggested,

2 H+ + 2 e - H2 T (2)

The protons, which are furnished by dissociated H2O, migrate toward

the aluminum. At the gold anode the following reaction is suggested:

40H- - 2H2O + 4 e + O2 T (1) (3)

complementing (2). Here the hydroxyl ions migrate toward the gold

electrodes releasing electrons for current flow in the external measuring circuit, accompanied by the release of oxygen as a gas. The

hydroxyl ions are supplied by dissociated KOH or H2O. In consideration of the K"*" ion and its migration to the cathode an equation similar

to (2) may be written, resulting in a greater capacity for hydrogen for- mation. Reaction (1) is diminished as the water content of the coating

is depleted. The conduction is ionic, resulting in the increased ionization of water at the cathode. The rate of these reactions determines

^^^The nimibers in parenthesis refer to the list of references appended

to this paper.

397
how many electrons are furnished for current flow in the external

measuring circuit but the mobility of the ions in the coating limits the conduction process.

Light Conduction

The above processes in the coating occur also during the illuminated condition but these are dominated by an additional reaction which

supplies a much greater source of current. Since the photoresponse was detected at wavelengths less than

3600 A an energy threshold greater than 3.4ev is required for photoconduction. A suggested reaction for this photo-conduction at the anodic surface is 20H- + hv - 2(0H) + 2 e + 2e (4)

2H,0, + hv - 2H2O + 0. T (5)

Here additional electrons are provided for external conduction and

additional gas is evolved. The cathodic reaction is the same as in (1) and (2). A contributing mechanism suggested for increased conduction in this condition is that the OH forms an amorphous -like valence band

permitting hole conduction.^^^

It is further suggested that reactions may still occur to provide

conduction even if all the water is depleted by electrolysis. The fol- lowing reactions are speculated:

KOH + hv - KOH"" + e (6)

2K0H+ - (K0)2 + 2H^ (7)

2H^ + 2 e - H2 T (8)

During the photo -response the data suggests that the resistivity de- creases since an increase in voltage accompanies the increased output

of current. (^^The numbers in parenthesis refer to the list of references appended

to this paper.

398
The curve in Figure 4 shows that there are two distinct rates of voltage decay indicating that the ability of the ions to polarize in the coating is changed after a certain length of radiation exposure time.

The curve in Figure 3 deviates initially from the exponential slope.

This suggests that the ions are not initially diffusion limited and that surface conduction might take place due to contamination on the surface of the coating. The surface of aluminum oxide is strongly hydroxylized which might contribute to this effect.^ "^^ The dark current reactions seem unaffected, however, after a similar radiation exposure. Water is present in the highly absorbent coating even after prolonged bakeout. As the photo -response progresses the coating dehydrates and this seems to be the dominating factor in attenuating the current and voltage output. This is supported by the fact that, at the conclusion of testing, the vacuum pimip was shut off and the pressure

in the chamber was allowed to rise slowly due to small leaks in the system. At 10" torr the voltage and current values were higher than

the final values observed during the testing schedule performed at 10"^ torr. At the cathode, equation (1) suggests that along with the creation of aluminum ions, impurities in the aluminum alloy electrode

would be mobilized via a related mechanism, and made available to in- fluence optical properties. Preceding toward the anode, dehydroxylation of the aluminum oxide according to the scheme

H

presents sites for charge transfer absorption. The doubtful oxidation

states of the impurities are conducive to color center formation. The manner in which the impurities function is important.

GENERAL CONCLUSIONS

The decay of the conductivity processes of this aluminum oxidepotassium silicate coating is related to the dehydration of water from

this coating. The conduction of the coating is normally ionic but when

the coating is stimulated by UV, hole conduction is the predominant

contributing factor to the conduction process. The coating is self

<^^Xrhe numbers in parenthesis refer to the list of references appended

to this paper.

399
polarizing and is enhanced due to UV radiation. Under UV exposure,

dehydroxylation of the surface of the coating coupled with the release

of impurities at the cathode and their subsequent diffusion may combine these effects to form color centers in the coating, with a net resultant increase in absorptivity.

ACKNOWLEDGEMENTS

The author is indebted to Dr. John B. Schutt for his illuminating

discussions and helpful suggestions. The author is also grateful to Charles M. Shai for preparing and applying the coating and to Robert N. Sheehy and his group for applying the gold electrodes to the coating.

REFERENCES

1. Latimer, W. M., "Oxidation Potentials", Prentice- Hall, New York

1952, p. 348

2. Ruff, I., J. Phys. Chem, 69, 3183 (1965)

3. Carter, J. L., Lucchesi, P. J., Cornell, P., Yates, D. J. C, and

Sinfelt, J. H., J. Phys. Chem. 69, 3070 (1965)

400
Fig. 2— Photograph of experimental apparatus; sample chamber and

light source.

401
EQUIVALENT SUN HOURS

0.25

" 0.50 1.0 2 5

10 20 50 100

102 10^ 10" 105 10' ;^ADIATION TIME — (SECONDS) Fig. 3—Relationship between UV photo-response of current vs. radiation time.

EQUIVALENT SUN HOURS

0.25 0,50 1 0 2 5 10 20 50 100 "

\ \

\ \ \ \ \ \ \

1.2 — ,

10 10^ 10^ 10" 10^ 10' RADIATION TIME - (SECONDS) Fig. 4—Relationship between UV photo-response of voltage vs. radiation time.
Paper No. 28

PLAINS AND STATUS OF XIIAB ad hoc CQ-CIITTEE 0}i TESTING FOR

PIIEDICTION OF :iATERIAL PERF0R:L\NCE IN STRUCTURES AND

CO.'Ii-ONENTS

Robert S. Shane

National Materials Advisory lioard, National Research Council - National Academy of Sciences, V.'ashin^ton, D, C.

ABSTRACT: National Materials Advisory Board - National

Research Council - National Academy of Sciences - a_d hoc

Committee on "Testing for Prediction of Material Performance

in Structures and Components" - Interim Renort

At tiie request of the Office of the Director of Defense

Research and Engineering, the National Materials Advisory

Board of the National Academy of Sciences has undertaken to review "Testing for Prediction of Material Perform.ance in

Structures and Components".

The basic purposes of the study have been identified as:

(a) make a selective survey of the needs for new and /or

improved predictive (including accelerated) testing techniques;

(b) consider predictive methods for forecasting the performance

of materials in structures and components under combined

stresses in service; (c) identify the factors of predictive

testing and explore approaches for development of methods and

techniques of predictive testing; (d) appraise the benefits of

predictive testing as v;ell as the risks of predictive testing;

(e) provide guidance in establishing a policy for Government

and/or Industry interaction; (f) recommend mechanisms for

implementing this policy.

x"m ad hoc Committee has been appointed and convened. The

principal problem areas appear to be (a) interaction phenomena,

(b) design of multi-factor tests wiiich forecast definitively

the behavior of structures and components in the presence of

knov7n or foreseeable failure modes, involving materials

(c) lack of systematic accumulation of existing information so that gaps in needed knov;ledge become apparent and remedial

action can be taken before design choices are irrevocably made.

403
(d) peripheral failure modes are frequently overlooked.

i^^L\B plans to co-sponsor the A3T:I National Symposiun on Predictive Testing, Anaheim, California, April 21-23, 1971.

KEY WORDS: combined environments, components, material performance, predictive testing, structures

(Complete paper not available)

404
Paper No. 29

MONITORING SERVICE TESTING BY NONDESTRUCTIVE TESTING

Robert W. McClung

Oak Ridge National Laboratory, Oak Ridge, Tennessee

ABSTRACT: Applications of nondestructive testing techniques

to raonitor tests include the use of x-rays v/ith both film and

television detection to monitor thermal cycling of fuel rods,

neutron radiography on irradiation experiments, eddy-current

measurement of space interval between components, and interim

monitoring by ultrasonics of the thickness of a vessel

in a reactor experiment. Studies during fabrication development include multiple interim measurements of compaction density of fuel rods and impregnation density of graphite using

quantitative gamma and x-ray attenuation as well as use of eddy-currents to monitor changes in quality of sodium bonding

in a fuel rod. Recommended approaches include the use of eddy-current methods to monitor changes in electrical properties, dimensions or flaw character, ultrasonics to measure

changes in elastic properties, dimensions or flaxj character,

penetrating radiation to observe or measure changes in density,

or dimensions and other forms of energy for similar applications.

KEY WORDS: com-Daction density, eddy current measurements,

fuel rods, imisregnation density, monitor tests, neutron

radiogr?»phy , ultrasonics, x-rays

(Complete paper not available)

405

Paper No, 30

SIGNATURE ANALYSIS - NON- INTRUSIVE TECHNIQUES FOR INCIPIENT

FAILURE IDENTIFICATION APPLICATION TO BEARINGS AND GEARS

Bo Weichbrodt and Kempton A. Smith'

REFERENCE: Weichbrodt, B. and Smith, Kempton A., "Signature

Analysis - Non-Intrusive Techniques For Incipient Failure

Identification Application to Bearings and Gears", ASTM/IES/

AIAA Space Simulation Conference, 14-16 September 1970. ABSTRACT: This paper describes how acoustic and vibration

signals generated by operating machinery can be used to diagnose its internal condition without using internal sensors which might perturb the system of interest. It is shown how characteristic "signatures" can be

electronically extracted from the overall noise and vibration

signals. By interpreting these signatures in the light of

engineering knowledge of the machinery under study, it is often possible to identify incipient failure modes long before

final failure. This technology opens up many new possibilities

to avoid unscheduled maintenance, improve product quality and

reduce testing time.

The paper discusses in detail specific applications to bearing and gear diagnostics and shows how the diagnostic

process can be electronically implemented and automated.

Several other application area are briefly discussed.

A major part of the work which is reported in this paper

was sponsored by the Condition Monitoring Engineering Operation

of General Electric 's Aircraft Engine Group, for development of

jet engine diagnostic systems.

Manager, Signature Analysis Unit, General Electric Research

and Development Center, Schenectady, New York.

2 Signature Analysis Unit, General Electric Research and

Development Center, Schenectady, New York.

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KEY WORDS: diagnostics, reliability, signature analysis,

vibration, bearings, gears.

INTRODUCTION ^ Both mechanical and electrical systems generate secondary

effects during their operation. These secondary effects, in contrast to primary effects, such as pressure and flow in a pump, are not directly utilized in the operation of the system.

Examples of typical secondary effects are vibration, acoustic

noise, heat generation, electrical and magnetic radiation.

(Obviously, in some cases these are also primary effects.) In fact, not only are the secondary effects not used, but

in many cases these effects also represent major problems, and

one would rather see that a system could be operated without

them.

However, luckily there is also a positive side to these

secondary effects, as this paper will attempt to illustrate.

Generated inside a machine they are usually transmitted to the

outside, carrying with them information about the internal

action. Provided that the external observer can properly

measure and interpret these effects, he has a new tool to gain

an insight into the internal condition without actually having

internal sensors. During the last several years much work has been directed

towards the development of machinery diagnostic techniques,

using these secondary effects. This has generated both

successes and frustrations. In the successful cases, it has

been possible to establish effective communications with the

internal action in a machine. This has led to several results.

One has been the capability to predict failure of machinery

much earlier than was previously possible by monitoring only

the direct operating parameters. A second result has been the

capability to predict life performance by testing only on the new product, i.e., better Quality Control and checkout techniques. This is of particular importance for space systems.

A third result, and maybe the most important, has been the

capability to actually identify specific failure modes without

disassembly, by only studying the external secondary effects.

This paper will attempt to illustrate some of the more important results which have come out of machinery diagnostics

work done at the Research and Development Center of the General

Electric in Schenectady, New York. The diagnostics program, which initially was concentrated entirely to acoustic and

408
vibration techniques, "Mechanical Signature Analysis", now is broadened into mechanical, thermal, electrical and magnetic

effects used for diagnostic purposes. It has been possible to develop technology and hardware, some of which is already being

used in aerospace and industrial applications.

However, in this paper we will concentrate on mechanical

effects, and how they can be used to diagnose machinery and

structures. Since this is the area where work was first

started, it has currently generated the largest number of

successful applications. Subsequent papers will discuss non- acoustic diagnostic techniques as they are developed.

After a brief discussion of signature extraction techniques, we will discuss in detail the application to rolling

element bearings and gears and in particular illustrate this

discussion with results that have been obtained in diagnosing

long life spacecraft gyros. After these detailed discussions

we will briefly show how the diagnostic technology, developed

at the General Electric Research and Development Center can

also be applied to a much wider variety of systems, including

nuclear reactors and artificial hearts.

SOUND AND VIBRATION AS DIAGNOSTIC TOOLS As most readers will easily recall, sound and vibration

have been used for diagnostic purposes for a long time. A

classical example is the skilled automobile mechanic who with

the help of a screwdriver and his "naked ear" can listen to an automobile engine and detect valve leaks and other defects long

before they show up as performance changes.

Without elaborating on this example, it illustrates the

proven fact that in order to successfully diagnose machinery by

sound and vibration, it is necessary to recognize quite subtle

differences between its normal and abnormal "signature". The

automobile mechanic does not listen to the engine's overall

loudness, but rather to the individual impacts from valves

hitting valve seats and other similar effects. The same is true for most machinery diagnostics applications.

Thus, one of the primary requisites for successful diagnostics is to know what subtle sound differences to listen for. This, in fact, is the most important requisite of all, without

which very few results have been obtained. It is necessary to understand the kinematics and kinetics of the mechanical system

under study, and also to know something about the signal trans- mission through the machinery structure. In many cases it has

been necessary to actually derive a dynamic mathematical model

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of the system, and to mathematically calculate the differences

caused be certain malfunctions.

Once the sound generating mechanisms and expected signa- tures are known, the next step is to devise means for extracting and analyzing these signatures from a complex total noise

signal. This sometimes requires very special techniques, as will become evident later in this paper. Finally, the techniques which have been developed must be

electronically implemented and in many cases automated.

When applied with ingenuity and understanding of machinery

dynamics, diagnostic techniques based on sound and vibration

can detect machinery failure in its very early stages. At

that time there is usually still time to take preventive action

to avoid catastrophic failure or unscheduled maintenance.

These non- intrusive techniques also have demonstrated their

capability to actually identify specific failure modes, which

is of prime importance since it strongly influences the preven- . tive action to be taken.

SIGNATURE EXTRACTION VS SIGNATURE ANALYSIS

The task of analyzing subtle sound or vibration patterns

in complex noise signals really consists of two problem.s. The

first problem is to separate from the remaining noise the

specific components in the signal, the "signature", which is generated by the part under study (i.e., one gear in a complex

transmission) „ This is the signature extraction task.

The second problem is to analyze the features of the signature, once it has been separated.

The second problem is obviously closely related to the

specific system under study, and the solutions tend to vary

from one case to another. However, the first problem, that of

signature extraction, is more general, and some specific ex- traction techniques have been developed which are used in most

applications. In order to clarify the following discussions,

a brief description will first be given of the most frequently

used signature extraction techniques.

SIGNAL PROCESSING FOR SIGNATURE EXTRACTION

Because signature analysis as a practical tool utilizes

secondary effects as a means of communicating with an operating

component, it is frequently necessary to detect and extract very low level signals from high noise level backgrounds. As

a result of this need, several specialized techniques of signal

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processing have been developed which allow particular features

of a signal to be enhanced.

Differing applications may of course call for different

processing techniques, but the majority of these techniques

fall into two categories; waveform analysis and spectral

analysis. It is important to understand these techniques so that an appreciation can be gained for the wide variety of

possible applications, previously not feasible without these

techniques

One of the most powerful tools available for extracting a periodic component waveform from a complex signal is the technique of signal averaging or summation analysis. A signal with

known periodicity T is buried deep in noise. To extract that

signal by averaging, the total signal plus noise is divided

into n epochs of duration T. Each epoch is sampled at 1000

points between times (n-l)T and nT. The samples are stored in

a digital memory of 1000 words which algebracically adds

together all n samples.

An event which always occurs with a period T will always

be added in the same sample location; i.e., the resulting sum

will be n times the average amplitude of that event. Events which do not repeat with a period T will result in a sum pro- portional to'Vn. Thus the desired signal has been enhanced

over the noise by a factor of ^ n. This process is illustrated

in Figure 1. In practice it may be desirable to incorporate a weighting

factor of one type or another or to select the period T as a multiple of a repetition period. A typical weighted average

de-eraphasizes old samples with respect to the samples just

entered, allowing the operator to observe changing waveforms in

real time. In many cases waveform analysis is able to provide only a part of the total characterization of a signature. Much additional information may be obtained by real time spectral analysis. Frequency filtering provides optimum signal-to-noise

relations and permits strong and weak components of different

frequencies to be detected simultaneously. The Research and

Development Center has developed a unique facility for the

thorough exploration of signal characteristics by the use of

frequency pattern analysis. The analyzer itself combines high

quality crystal filtering with a high scanning rate to preserve

both transient and steady state features. The analysis para- meters are extremely versatile. Signal frequencies from 1/2

to 125,000 hertz have been analyzed with effective filter

bandwidths ranging from 1/4 to 3,600 hertz. Many display for- mats are available, including cross section plots of amplitude

411
versus frequency, several forms presenting amplitude by intensity modulation, amplitude level and contour plots spectrum

peak displays, and other combinations. In addition, computer

processing of the analyzer output is available for the automatic extraction of information in signal spectra. An example of

this type of analysis is included in the discussion of bearings

below.

DIAGNOSTICS OF ROLLING ELEMENT BEARINGS

Ball and Roller bearings usually follow a characteristic

wear/failure pattern. Since the sliding between metal surfaces

is minimized, the wear rate is usually very low. But the rolling contact load between balls/rollers and the race is instead

very high, and leads to sub-surface fatigue. Therefore, rolling element bearings have a tendency to fail by fatigue rather

than wear- out. Obviously, the statements above are very general, and can

be challenged in special cases. But they describe the typical

situation for bearings which are properly designed and maintained. All bearings can fail through many other mechanisms if

they are not properly used, for example through corrosion or lack of lubrication.

Figure 2 shows a simplified but typical wear rate curve

for a rolling element bearing.

Early in life the wear rate is often at a maximum. This

is the running- in period. The maximum may be absent for very

high precision bearings. After this initial period comes a long period with very low wear rate. This represents the major

portion of. the bearing life. During this period the wear rate

is low, but both races and balls accumulate fatigue cycles.

The end of the low-wear period is marked by the appearance

of some surface defects, usually on one of the races. This

may be fatigue spalling but also corrosion spots or other

effects. The surface defects are small at the beginning, but

disturb the smooth rolling process, increase the friction, and

therefore grow and increase the wear rate. When a sufficient amount of surface defects have accumulated, the bearing fails

to perform its function. It may jam or fail to hold the

position of the shaft, so that secondary damage occurs. It may

overheat and burn the lubricant. This is the end of the bearing

life. As discussed above, the bearing goes through a number of

stages on its way to final failure. Obviously, it is desirable

to detect the problem some time before the final failure occurs.

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But when? The timing is critical, because too late may be catas trophical but too early causes premature removal and is therefore uneconomical.

It can easily be shown that there is more than one answer

to this question. In fact, different applications may result

in completely different detection philosophy and completely

different detection techniques.

Consider the situation where a radar antenna rotates resting on one roller bearing. Or a jet engine in an aircraft,

where the main rotor is axially held by only one ball bearing. In both these cases final failure of the bearing would cause complete performance failure and heavy secondary damage. It is desirable in this case to detect the appearance of the first

fatigue spalls, which would correspond to point A on the wearrate curve. At that point most of the bearing life is expended, and the risk for rapid failure high. This is the time for

requesting a replacement bearing for the antenna, or to schedule maintenance of the jet engine.

Consider now another situation - a spacecraft gyro undergoing final checkout before acceptance for a mission that

requires long life. The components which usually fail first

are the rotor bearings. In this case there is obviously no reason to look for fatigue spalls in the bearings, which are

new. Instead one will want to look for the existence of

initial imperfections such as misalignment, lack of roundness

of bearing surfaces, etc., which may increase the load at some locations on bearing races and thereby cause early failure.

These are two entirely different definitions of a bearing

malfunction, suited for different applications. Others may fit

better in other cases. For the purpose of cathegorizing detection methods it has been found practical to distinguish between

local defects (spalls, corroded spots, Brinelling) and

distributed defects (misalignment, lack of roundness, unequal

ball diameters) as Figure 3 illustrates.

Many different techniques are used to detect bearing malfunctions in operating machinery:

Oil contamination: Spalling of bearing surfaces generates

metal chips which follow the lubrication oil and can be

collected in filters or with magnetic plugs. Smaller wear

particles may be detected with spectroscopic or radiographic oil analysis methods.

Temperature : The increased friction which precedes bearing failure generates heat that increases the temperature

413
of bearing and lubricant. Oil temperature is often

measured.

Overall Noise Level : When the spalling has proceeded

sufficiently the bearing gets noisy. The problem is that

bearings are usually buried deep down in machinery, and

there may be many other noise sources. All these techniques and others, have been tried with

varying degree of success in many cases. The problem is gen- erally that the malfunction indication comes too late, so that

there is no time to take preventive action before total fail- ure has occurred. Another major problem is that these techniques do not identify the failure mode. This is unsatisfactory

since failure mode identification is often necessary to decide

what corrective action should be taken.

With this background, the work at the Research and Development Center has been directed toward developing methods for

earlier detection of bearing malfunctions than what has pre- viously been possible. This has involved diagnostic techniques

based on vibration, or structure- borne sound, generated by the

bearing and picked up with an external sensor. Analytical

predictions have been made of specific vibration patterns gen- erated by good bearings under various load conditions, and

bearings with defects as discussed above. A number of projects,

carried out at the R&D Center with the cooperation of several

product departments, have given opportunities to compare analytical predictions with actual test results. It has been generally found that it is possible to make analytical predictions of

vibration patterns from bearings with a number of defects. By

employing selective data reduction techniques to extract the

calculated patterns from the overall vibration signal, it has

been possible to detect defects earlier than with other known

methods

In the following a few examples will be given to demon- strate how bearing malfunction signatures can be predicted and

extracted from complex vibration signals. The examples will

cover both local defects, very briefly since these have been

discussed in earlier papers, (Reference 1) and distributed

defects where the work has been done more recently.

DETECTION OF FATIGUE SPALLS AND OTHER LOCAL DEFECTS

A local defect in the inner or outer race of a bearing

generates an impact every time a ball (or roller) rolls over it. Likewise, if the defect is in a ball surface, an impact is generated every time it hits the inner or outer races. If the

bearing rotates with constant speed each defect generates a regular set of impacts with an impact repetition rate that

414
depends upon bearing speed and location of the defect. Assuming

that no sliding occurs in the bearing, those fundamental repetition rates can be calculated. The result is: Outer race f =^f (1--^ cos p) (cps)

. e2r PD malfunction

Inner race f^ = ^ f^ ( 1 + |^ cos p) (cps)

malfunction

Ball S = Id ^r " ^Id^^ "^^^ ^"^P^^ malfunction

where is the relative speed between inner and outer race (revolutions per second), and BD, PD, p, and n as defined in

Figure 4. The vibration pattern, or signature, which can be expected

from a single local defect in a bearing then consists of a series of transients with a repetition rate as calculated above.

Figure 5 shows such an ideal pattern.

The problems in practical cases are that the signature

from the bearing defect is often hidden behind background noise,

and also that the repetition rate can only be calculated approximately-o It has been discussed in an earlier paper (Reference

1) how a diagnostic system may be designed to extract this

signature from a local defect, and thereby make the diagnostic

system more selective. In addition to the mere transient re- petition rate, there are many characteristic features associated with each individual transient. These details of the

transient waveform depend on the load pattern for the bearings,

and on the structural response to the impacts between rolling

element and defect. As an example, consider the effect of an unbalanced rotor,

supported by the bearing under study. In addition to the

gravity load, which always points in one direction, there will

now be a rotating unbalance load acting on the bearing. This

combination of loads will cause a modulation of the ball-race

contact forces. A local defect on the outer race, for example,

will see a 1/rev. modulation of the loads on the balls rolling

over the defect. This effect will cause a modulation of the

amplitudes of the impact transients. The calculated signature

will look as shown in Figure 6a. The modulation of the envelope is confirmed by Figure 6b which shows an analysis of vibration data measured on a jet

engine with a large spall in one of the main ball bearings.

The signal was analyzed through signal averaging as described

previously in this paper, in this case programmed to enhance

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events repeating with the envelope repetition rate rather than

the individual transients.

The previous discussion has involved prediction of repetition rates and transient amplitudes, but not any details of the

individual transient waveform. If the mechanical structure supporting the bearing is known, it may be possible to calculate its response to an impact at the bearing race. This is going to result in a number of "ringing frequencies" with which

the structure will ring after the impact. It has been found in one case, for example, that the structural response to the

impact from a local defect in a jet engine bearing can be de- scribed with a very low number of ringing frequencies.

Figure 7 illustrates this case with analysis of a vibration

signal measured inside a jet engine under idle conditions. The

engine had a spall in one main bearing.

The signature shown in Figure 7 was extracted from the same raw vibration data as shown previously in Figure 6 through signal averaging as described in Reference 1, but

without rectification before the summation. The analysis was here programmed to display the individual impact transients.

These are only a few examples, serving to illustrate the

point that vibration signatures can be analytically predicted,

and extracted from vibration data measured on machinery with

local bearing defects. By knowing as many characteristics as possible in the predicted signature, the sensitivity of the

diagnostic system can be maximized, and its false alarm rate minimized.

The success in detecting and identifying localized bearing defects has led to the development of an electronic system,

which automatically performs the signature extraction and

interpretation. This was developed in a joint program between

General Electric 's Research and Development Center and Aircraft

Engine Group, Condition Monitoring Engineering. A laboratory

model of this bearing condition monitor is shown in Figure 8. Miniaturized and further automated systems are currently being

developed by the General Electric Company both for aircraft and

industrial uses.

DETECTION OF DISTRIBUTED BEARING DEFECTS

In certain cases it is not sufficient to detect the first

local surface defects in bearings. This was discussed above

in connection with final checkout of gyros. The requirement is instead to detect such initial conditions which might cause high local forces in a bearing and therefore early failure.

These are conditions such as misalignment, lack of roundness of

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bearing races, unequal ball/roller diameters, etc., which are always built into the rotor system to some degree during

assembly.

The capability which has been developed to detect such

internal conditions through external measurements will be

discussed in connection with a specific example.

The study was made to determine the feasibility of using

vibration signals to predict the life of spacecraft components. As a first suitable test vehicle one type of rate gyro was

selected. Eight similar gyros of this type were avaialble for

the study. Rate gyros are critical components in spacecraft navigation systems. Therefore, much work has been devoted to generating rate gyros with long life expectancy. This work has

resulted in gyros with long average life. However, the increase

in average life of gyros achieved through higher quality and

precision is not matched by a similar increase in the uniformity of individual gyro lives. In other words, even if average

life of a certain t5rpe of gyro is long, there is still con- siderable variation around this average for an individual gyro. This large variation becomes of critical importance when the

need for extremely high reliability arises. In this situation, a screening tool with the capability

to select the most long lived individuals within a group of gyros would offer one solution. However, a manufacturer of

high quality gyros has already before delivery employed a number of sophisticated techniques to assure uniform performance. A useful screening tool must therefore be more sensitive than conventional checkout precedures.

The primary purpose of this study was to determine whether

mechanical signature analysis might provide such a sensitive

tool for gyro checkout.

The major failure reason in rate gyros is failure of the

ball bearings which support the rotor. The work was therefore

concentrated to such ball bearing defects which might be present in a new gyro motor and later cause early failure.

Initial vibration analyses of eight rate gyros, all

performing within specifications, showed large variations

between different gyros. This was interpreted as evidence that

vibration analyses provided an extremely sensitive tool.

The initial test also showed that gyros generate a very

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complex vibration spectrum, containing many hundred tones distributed over a very wide frequency band. Most of the effort

in this study was therefore directed towards the interpretation

of the complex gyro signature in terms of such parameters which

might influence gyro life. This work included both mathematical

modelling of the generation of bearing vibration signatures,

and testing of gyros with inserted malfunctions. The analytical

work was necessary in order to establish acceptable confidence

in the results, since the eight gyros which were available did not by any means provide a statistically satisfactory sample.

MATHEMATICAL MODELLING - DISTRIBUTED DEFECTS In order to predict the vibration signature from a bearing,

it is necessary to first consider the source system, i.e., the

bearing and its immediate support structure. This is where the

signature is generated. In many cases it is not possible to locate the vibration sensor close to the bearing, and in such

cases the transmission of the vibrations from bearing to sensor

must also be considered.

Consider the source system in Figure 9. This is a ball

bearing with six balls. The primary forces, which are the source of all vibration, are the six contact forces between

balls and races. In the figure it is assumed that the inner race is stationary and connected to the gyro frame. If instead the outer race

is stationary, like in most electrical motors, the forces

between balls and the outer race are considered.

The ideal situation, with no defects present in the bearing, is that all ball-race forces are equal and constant. (Gravity forces are negligible compared to pre-load forces in

the gyro) . The primary vibration source is then the six constant ball forces, rotating with the same constant speed as the ball cage, and acting on the stationary bearing ring.

When defects are present, the normal ball-race force

pattern will be disturbed. Figure 10 shows the type of pattern

which would be expected for a misaligned and an elliptic bearing race. These disturbances of the normal force pattern are entirely

different from the impacts generated in a bearing with spalls

or other local defects. Therefore another approach to the

analytical modelling must also be taken.

Without going into great detail here, it can be shown that most of the "initial" bearing defects can be represented by

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disturbances of the force pattern similar to the examples shown

in Figure 10. In a general approach to the problem, it is assumed that the ball-race force pattern follows a sinusoidal

profile as shown in Figure 11. More complex solutions can

then be constructed through Fourier analysis. The sinusoidal

pattern may have an arbitrary angular velocity in order to simulate defects located on inner race, outer race, or balls/

rollers

The magnitude of each ball-race contact force is the

following function of Cf

P = P_ ' 1 + € cos m (0 - pt) , ) o

P = Force pattern angular velocity (rad/sec)

0 = Angular coordinate (radians) m = Number of nodes of force pattern around race circumference

e = Degree of force modulation

The dynamic radial deflection § of the stationary bearing

ring may be calculated by repeated solving of Lagrange's

equations for the system.

/ P '

j ,2 o sin ncu(t-0) 1 -f- £Cos m(cot-0-6t) /

i n,sin n,sin M o) v

! n

^ 2

' q + o) q = — cos na)(t-9) 1 + ecos m (a3t-0-6t)/ ^n,cos n^n,cos M — \ k // n 03 ^ ) where q , and q are generalized coordinates, and M

n,sin n,cos n are effective mode-masses. After a number of manipulations an expression for the dynamic deflection § caused by all the ball

forces is constructed:

°° P 8 = Z Z _o

9 n=l M

cos n(cot - Q - 4^ ) 2 2

o) n + £ A cos m ^ (a>p) t-Q } cos n ( cut Q - 4^)

419
where

A

2 2 2, o^2 2

B 2mn (a>|

CO It can directly be seen that the solution for 5 consists

of several series of modulated tones, with frequencies and

modulation patterns depending on the type of defect. By

setting e = 0, the solution for a "perfect" bearing is obtained.

Further study of the effects of signal transmission from

the bearing to the outside of the gyro shows that certain tone patterns propagate easily whereas other patterns are heavily

attenuated through force cancellation.

The conclusions of these calculations is that characteristic vibration signatures, in the form of tone patterns, can be

analytically predicted for most of the defects under study.

This provides a tool with which the very complex vibration

patterns generated by the gyros can be understood.

As an example of the highly complex tone pattern generated

by the two rotor ball bearings in a gyro, Figure 12 shows a high speed spectral analysis of the vibrations measured on the

outside of a gyro. Horizontal scale is time, vertical scale

is frequency. Dark lines indicate tones in the signal. The

gyro was first run at constant speed for twenty seconds, and

then coasted down to zero speed. In order to practically test the calculated bearing malfunction signatures, the eight rate gyros under study were

sent back to the manufacturer for insertion of minor defects

in six of the eight gyros. Three types of defects were selected as important for gyro life. They were:

o Rotor Bearing Misalignment

o Incorrect Pre-Load of Rotor Bearings

o Rotor Unbalance

This selection was made in agreement with the gyro manu- facturer. The three defects were inserted in two gyros each,

thus making six defective gyros. The two remaining gyros

out of the total group of eight were assembled without malfunctions . 420
The severity of the defects was generally such that all gyros were within or almost within performance specifications.

The eight gyros were run at the Research and Development

Center, vibration signals were recorded and analyzed. The

analysis results were compared to previously generated analytical models of predicted signatures for the inserted types

of malfunctions. At the time when this was done it was known

which malfunction types and how many of each were inserted in

the gyros, but not which malfunction was inserted in each

particular gyro. The result of this independent signature

analysis is summarized in Table 1. Thus of the eight inserted malfunctions seven were detected and identified correctly.' As far as the gyro No. 6 with

3 pounds preload instead of 2 pounds, it is conceivable that

the preload actually was lower than 3 pounds. It is generally

very difficult to control bearing preload during gyro assembly,

this is one of the major reasons why this type of defect was selected for the study.

The seven correctly detected defects were detected by

ranking after study of spectral analysis plots and interpretation of the data based on previous mathematical modelling. It

was at that time not attempted to assign numerical values to the defects. As shown in Table 1, the evaluation resulted in prediction

of several more bearing defects than those which were intentionally inserted. At the time when these predictions were

made it was not possible to get them verified since all gyros

had normal performance. However, all eight gyros were subjected to a life test and then torn down and inspected. The life

test duration was approximately six months, during which none

of the gyros failed to run. Shortly before the end of the

life test, gyro No. 1 was shut down because of markedly

increased overall noise level.

The inspection of the gyro bearings after teardown was

done in considerable detail. As far as the two types of

additional defects predicted by MSA before the life test, the

result was

Lack of roundness, outer bearing race , predicted for gyros

Noo 5 and 6. Completely confirmed by measurements.' The

two tested gyros have bearing housings with 150-200 microinches lack of roundness compared to normally 20-50 microinches for the other gyros. Unequal ball forces , predicted for gyros No. 1, 2, and 8. Some indication also for gyro No. 7. The most likely

421
reasons for unequal ball forces would be unequal ball dia- meters within one bearing. The balls from gyros No. 2, 7, and 8 were measured and found to be well within specifications.' The other gyros were not measured. This may

initially look like a negative conclusion. However, one observation of great interest can be made. Of the eight

tested gyros, only three showed heavy bearing wear after

teardown. The rest were sometimes pitted and brinelled

but did not show significant wear. The three gyros with worn bearings are the same three which showed the characteristic "unequal ball force" pattern when new. Thus even

if the reason for the pattern is not unequal ball diameters, it may still be significant. If it can be further

confirmed that this frequency pattern is reliably

associated with early wear, it might well be the most im- portant result of the study.'

The techniques for bearing condition monitoring which were described above have later been further developed and refined,

among other things to include the effects of multiple defects

occuring simultaneously. These diagnostic techniques are currently being used in a program with NASA to evaluate gyros

for the Apollo project (Reference 2).

APPLICATION TO GEAR DIAGNOSTICS

A pair of properly operating gears, running at constant

speed and load can be expected to generate a signal which is periodic with the meshing frequency. The signal can be sound,

vibration, stress, strain or something else, and all of the

meshing events are similar when there are not mechanical

imperfections in the gears. A typical waveform generated by a gear with ten teeth would look like the example of Figure 13. Given the gear geometry and speeds, the repetition rate

1/T is very simple to calculate. The characteristic waveform

F(T) depends on gear profile, load, speed, coupling between

gears and sensor and many other parameters, and is very difficult to calculate. Some predictions can be made, however,

about the perturbations resulting from certain defects.

Generally gear defects can be separated into two categories

1. Local Defects . These are defects which appear on one

or more teeth but are not uniformly distributed

around all teeth. Gears with these defects, cracked

or deformed teeth, or scoring for example, will show

the effects of those defects only when the defective

tooth is in mesh. This makes interpretation a rela422
tively simple matter; when the waveform for one revolution of the gear is extracted, each tooth mesh can be

compared with the others. If the waveforms are not all the same, within some limit, a defect is present.

2. Distributed Defects . These are defects which are distributed all around the gear and thus the effects

will be seen everywhere on F(T) . A typical example is a gear which has a non- uniform load distribution on

its teeth as a result of eccentricity of its shaft.

Because of the changing load during each gear revolution the amplitude of the mesh waveforms is altered

in a fashion proportional to the load. The resulting

waveform appears as a carrier frequency at the mesh

rate with amplitude modulation at the rotational

frequency. Distributed defects may be either symmetrical or non- symmetrical. In the latter case, interpretation is facilitated by comparing individual teeth. In the former case, interpretation of the signature

may be much more difficult.

A Gear Condition Monitor, shown in Figure 14 as part of a Jet Engine Diagnostic System, has been built at the Research

and Development Center in a joint program with the Condition

Monitoring Engineering Operation of the Company's Aircraft

Engine Group to extract and automatically analyze the signatures

of individual gears in a jet engine gearbox. The monitor pre- processes the signal by bandpass filtering and waveform aver- aging to isolate one gear signature from the total signal

sensed externally on the gearbox housing. The extracted wave- form is then analyzed automatically for the two defect categories mentioned above. One part of the automatic analyzer, which identifies

defects in the first category above, performs a crest factor

analysis. A normal gear tone, with all tooth meshes equal in

shape and amplitude will have a certain ratio of peak to root mean square amplitude. If the shape of a tooth mesh changes,

(e.g., one tooth mesh signal is greater than the next) this

change will be reflected in a different (in general increased)

peak to RMS ratio. This ratio is displayed by the analyzer for

every data analysis performed.

The second portion of the automatic analysis system is designed to detect and display the percent modulation of a gear tone, as an indication of the presence of distributed

defects of the non- symmetrical type. The meter on the front

panel is calibrated to read directly the percent modulation of

the averaged signal stored in the memory of the digital averager. Permanent records of the waveform may be made.

423
Illustrations below are an example of how the signatures

of these defects can be extracted and displayed. The data was collected from a large transmission operated on a test stand.

The housing of the transmission was instrumented with accelero- meters sensing the structure-borne sound generated internally

by the many operating components.

The first figure is the averaged waveform of the planetary

gear train, composed of a 76-tooth sun gear 15.2 inches in

diameter, ten planets 5.4 inches in diameter with 27 teeth

and a 26- inch diameter 134-ring gear fastened to the housing.

The meshing frequency, at normal operating speed is 550 Hz. To extract this waveform, the sensor mounted over the ring gear

support on the housing was interrogated. The signal was filtered and averaged in the waveform analyzer with the basic period

T selected at one revolution of the carrier.

There are two basic features to the resulting averaged

waveform. First, the higher frequency signal is the sound

generated by the tooth meshing, in this case 550 hertz.

Modulating this meshing frequency is a lower frequency,

occuring at ten times per carrier revolution. Each modulation

cycle is generated by one planet in mesh with the ring gear

passing by the accelerometer . Since the acceleration measured

at a point on the ring gear is dependent on the forces between

the meshing teeth, the extracted signal indicates relative

load sharing between the planets. In this case, one planet is obviously carrying less load than the rest. (This display

shows parts of nine out of the ten planets).

The parameters of analysis for this gear were chosen to display the relative signals generated by each planet; hence,

the selection of T at the carrier pass period. Other para- meters are possible, for example, if T were selected at the

tooth mesh period times the total number of teeth on the outer

ring gear, the average waveform would be the average signal

from each tooth. Thus, this technique permits isolation of

defects that could not otherwise be studied.

Figures 16 and 17 are signatures of a bevel pinion gear. This 23- tooth gear is rotating at 13,800 RPM and is in mesh

with a 92- tooth bevel gear. In Figure 16 a repeating pattern

is visible every revolution of the pinion gear. The most obvious part of this repeating pattern are the three spikes

across about one millisecond of the signal. When the trans- mission was inspected after several hours of running, the gear

in question was found to have hard wear patterns on five teeth.

The gear was replaced and the signature in Figure 17 resulted;

some modulations present but generally a much more normal

signature.

424
In this example, the driving gear and the driven gear have

an integral ratio of teeth, making it possible to display the

effects associated with both gears simultaneously. The more common occurence is that the ratio of teeth is non- integral

requiring that the summation process be performed twice, with

the period T selected for each gear. Further discussion of gear diagnostics can be found in a previously published paper (Reference 3).

OTHER APPLICATIONS OF SIGNATURE ANALYSIS TO MACHINERY

DIAGNOSTICS

Similar techniques and logic as described above for gears

and bearings can also be applied to many other machinery systems

The following paragraphs will give a brief summary of some of

these applications with reference to other papers where more detailed discussions can be found.

DIAGNOSIS OF PUMP SYSTEMS

Both reciprocating and rotating pumps generate repetitive

signatures which can be extracted from background noise with

signal averaging, performed so that speed variations are tracked by the processing system. This has been successfully done

for both piston and vane pumps and has led to the development

of an automated diagnostic system for hydraulic drives of

satellite tracking antennas. (References 4 and 5) An example

of signatures from a vane pump in normal and defective condition is shown in Figures 18 and 19.

DIAGNOSIS OF MACHINING PROCESSES

The metal cutting action during machining generated

characteristic mechanical signatures which can be used to automatically detect low level tool chatter or cutting edge wear without stopping the machine for visual inspection. This

capability is of great importance in developing adaptive machine tool controls (References 6, 7, and 8).

ANALYSIS OF HEART SOUNDS

Significant gains have been made in the field of Phonocardiology by the Research and Development Center by applying

the techniques of Signature Analysis to the sounds of normal

and abnormal hearts. It is now possible to specify characteristic waveforms associated with certain types of heart ailments

425
and to provide the medical profession with an advanced tool to aid them in diagnosis.

An outgrowth of the Phonocardiology Program has been the

analysis of the sounds generated by a cardiac assist device.

This program is providing the design engineers with information

regarding the performance of the pumps - speeding the day when

assist devices will be available for implantation.

Then these techniques will be used to monitor the conditixi

of a patient with an assist device and keep the physician in- formed of its operation. Figure 20 is a typical example of

the vibration signature of a left ventricular assist device.

Points 1 and 3 are disc valve closures. At point 2 a valve

is bouncing against the support structure. At point 4 the

pumping bag fills and uncol lapses. Deviation from this normal

signature would indicate such defects as occluded valve seats

or ruptured bags.

APPLICATION OF SIGNATURE ANALYSIS TO NON- REPETITIVE PHENOMENA

Although the applications of signature analysis discussed

above have focused on repetitive signals, there are many times

when a much simpler type of analysis may be possible. In

particular, if the signal to noise ratio is high, no specialized extraction technique may be necessary and the characteristic signature of a non-repetitive event may be recognized.

Two examples of this may be taken from work performed to develop diagnostic techniques for nuclear reactors. Hydraulically operated control rod drives contain several

mechanical systems which are critical to operation of the

reactor, notably the latching mechanism for retaining the rods

and the hydraulic control valves. Figure 21 illustrates the events which can be seen in the direct signal obtained from

accelerometers mounted on the drive housing and the accumulator. At 1 the drive valves are opened, generating an impact transient followed by a high level of hydraulic noise. At 2 the

retaining fingers latch. At 3 the drive valve closes, again

creating an impact transient following by a high level of

hydraulic noise. At 4 the settling valve closes and the cycle

is complete. Here is a case where the relevant acoustic signa- ture can be directly observed with little or no processing.

Often the only necessary processing may be wide band

filtering. Another application to diagnosing nuclear reactors

is in the location of steam leaks in heat exchangers. Several

sensors around the exterior of the exchanger are tuned to "listen" for noise in a particular frequency band which is gen- erated by the leak. Correlation and triangulation between

426
sensors then locates the source of the leak within the heat

exchanger, A similar approach has found application in studies in- vestigating cavitation phenomena. A cavitating fluid transmits

transient bursts of energy to the structure of the container. A sensor on the container walls is sensitive to these transients

and evaluates them by the crest factor analysis technique al- ready discussed.

CONCLUSIONS

The work which has been discussed in this and previously

published papers has proven that Mechanical Signature Analysis

techniques can successfully be applied to almost any t3^e of

machinery. Particularly when machines generate repetitive

signatures, advanced signal processing techniques are available

to extract signatures from background noise, and provide auto- mated diagnostics. Of particular value is the fact that

specific incipient failure modes can be identified, usually

long before changes in operating parameters show up. Today's diagnostic technology can often directly be applied

to many types of bearings, pumps, gears, and reciprocating

mechanical systems. In some cases, such as for bearings and

gears, diagnostic hardware has already been developed and is presently in use for aerospace and industrial applications.

In other cases, research and development efforts are currently

underway.

With diagnostic systems, such as discussed here, considerable savings can be made in terms of reduced unscheduled maintenance, improved product quality, and reduced testing time.

427
Table 1 Results From Mechanical Signature Analysis

Gyro

Number

MSA Independent Results

Prediction of Additional

Inserted Inserted Mal- Conditions

Malfunction function Types Predicted

Rotor dynamic Rotor unbalance Unequal ball

unbalance Second highest in forces

the group

CORRECT

High bearing Preload highest Unequal ball

preload in the group - forces

(4 Ibo instead significantly

of 2 lb.) separated from

all other tested

gyros

CORRECT ; Misalignment Outer bearing Bearing preloac

of outer bear- race misaligned. almost totally

race CORRECT : absent

Misalignment Outer bearing None

of outer race misaligned

bearing race CORRECT : None- gyro Gyro normal as Lack of

normal far as inserted roundness

malfunction types outer bearing

are concerned. race. CORRECT .' High bearing Gyro normal.' Lack of roundpreload(3 lb.. incorrect: ness, outer instead of bearing race

2 lb.)

Rotor dynamic Rotor unbalance None

unbalance highest in group. CORRECT : Normal Gyro normal Unequal ball

CORRECT .' forces

428
Weichbrodt, B. , "Mechanical Signature Analysis, a New

Tool For Product Assurance and Early Fault Detection",

presented at the Fifth Reliability and Maintainability

Conference, New York, New York, July 18-20, 1966,

sponsored by AIAA, SAE, and ASME. It was published in

the "Annals of Reliability and Maintainability", vol 5 "Achieving System Effectiveness", American Insititute

of Aeronautics and Astronautics, 1966. Sturgeon, Lloyd, "Final Report on an Acoustic

Signature Analysis of 32 Gyro Wheel Assemblies",

prepared for the Massachusetts Institute of Technology

Ref. P. 0. No. IL-B-(M)-300207, Document No. 70SD4230,

April 16, 1970. Thompson, R. A., and Weichbrodt, B. , "Gear Diagnostics

and Wear Detection", published by the American

Society of Mechanical Engineers, 69-VIBR-lO, March 30- April 2, 1969. Tatge, R. B., "Acoustic Techniques For Machinery

Diagnostics", presented at the Seventy- fifth Meeting

of the Acoustical Society of America, May 21-24, 1968,

Ottawa, Ontario, Canada.

Tatge, R. B., "Final Report Study for Maintenance of

Hydraulic Motors", prepared for National Aeronautics

and Space Administration, Goddard Space Flight Center,

under contract No. NAS5- 10559, October 1968. Thompson, R. A., "The Modulation of Chatter Vibration^,

published in the Journal of Engineering for Industry , Paper No. 68-WA/Prod- 21, also presented at the Winter

Annual Meeting of the ASME, December 1-5, 1968. Weichbrodt, B., and Weller, E. Jo, "New Development in

Cutting Tools: 'Listen to Your Tools - They're

Talking to You'", ASTME Paper No. MR67-144. Paper

was also presented to the ASTME 1967 Engineering

Conference and Tool Exposition, Chicago, Illinois,

April 24-28, 1967. Paper was published in "The Tool

and Manufacturing Engineer", July 1967. Weichbrodt, B. , Weller, E. J., and Schrier, H. M.

"What Sound Can Be Expected From a Worn Tool?", ASME

Paper No. 68-WA/Prod- 4, presented at the Winter Annual

Meeting, December 1-5, 1968, New York, New York of the

American Scoeity of Mechanical Engineers.

429
f(t)

i^^T ill ill Rectification ^ N-1

1)1=1 f[r+ (n-l)Tj|:

LA

LA

Figure 1 Signal Smranation Analysis

430
431
432
433
\ a;

434
Figure 6a

Predicted Signal

Figure 6b

Measured Envelope

Predicted and Measured Modulation of Impact Sequence From

Local Defect

435
00 . W O

0) ctf ^ 1

•H CO

(U C

0) -H Q W

03 bO O 3 O O '^^ oC (U

(U fl 3 0) o -P

cfl CO e 00

CO

CO bO

<-< C

•H -H

Cfl M

+j cd Q pq

0)bO

fx,

436
437
Figure 9 Ball Bearing with Primary Excitation Forces
439
6000

N7

5000

Nfi

o

Xio

c

cro

u

4000

N4

3000

N3

2000

fN2

1000

fNl

lo I50 llOO Time, sec

Il50 I2OO

Figure 12 Intensity Modulated Spectrogram of Vibration Signal

Steady State and Coast Down of Gyro.

440
f(t)

Figure 13 Gear Mesh Waveform

Figure 14 Jet Engine Diagnostic System

441
442
443
>

u

444
• - •

Figure 18 Part of Vane Pump With Surface Defects

445
VANE PUMP RESPONSES

I* 1/REV. TRIGGER (APPROX, 50 M SEC.) H

NORMAL PUMP

DEFECT: GALLED RING

Figure 19 Mechanical Signatures From Normal

And Defective Vane Pump

446
447
448
Paper No. 31

MEASUREMENT METHODS FOR MICROCIRCUITS

W. Murray Bullis

REFERENCE: Bullis, W. Murray, "Measurement Methods for Micro- - circuits," ASTM/IES/AIAA Space Simulation Conference,

14-16 September 1970.

ABSTRACT: One of the problems encountered with testing in

either simulated or real environments is associated with the

necessity of demonstrating that the failures observed are, in

fact, due to the environment rather than random effects which

would occur even in the absence of the environmental stimulus. • This problem is particularly acute in the case of high-reliability electronic components such as silicon monolithic integrated circuits. For these parts, desired levels of reliability

can be achieved only if all process steps in their manufacture

can be carried out in a controlled and reproducible fashion

This paper surveys the status and availability of standard test methods suitable for making the measurements which are necessary

to achieve such control. Consideration is given to standard measurement methods appropriate for thin film integrated circuits and hybrid microcircuits as well as those for silicon

monolithic integrated circuits fabricated by the epitaxial,

planar process

KEY WORDS: device fabrication, environmental testing, integrated circuits, measurement methods, microcircuits, silicon

Testing in a simulated environment is usually undertaken in order to determine the effect of an environmental stimulus on

the characteristics of the object being tested. Even if the

simulation is assumed to be accurate, to obtain a valid test one

1 National Bureau of Standards, Washington, D.C. 202 34

449
must be able to demonstrate that the changes in characteristics

are , in fact , due to the environment rather than to random effects which occur even in the absence of the environmental stimulus. In particular, it is frequently desired to establish

wheth.er the values of a certain set of parameters change to such an extent that the object will no longer perform its desired

function. Objects in which one or more of the characteristics

change to this degree are classed as failures. In addition an object may fail catastrophically , as by cracking or burnout, and

thus may no longer be able to perform its function.

High-reliability electronic components such as silicon

monolithic integrated circuits are designed to have extremely

low failure rates even under the influence of various adverse

environments. To show that the failure rate induced by environmental stress significantly exceeds the random failure rate re- quires large samples and carefully controlled experiments particularly when the two failure rates do not differ by very much.

As a result, tests to establish performance characteristics

or failure rates under adverse environmental conditions are often unsuccessful. This is a particularly serious problem in . connection with tests in nuclear radiation environments. When

inconclusive results are obtained in underground tests with nu- clear explosives the direct cost in dollars and resources which

is wasted is self evident. Not so obvious is the costly but

fruitless effort expended in design and analysis of the experiment. If the experiment should lead to incorrect conclusions

the cost is even higher. These last costs are also encountered

in tests conducted in simulated environments . Even though the

environment can be carefully controlled in this case, experiments may fail because of parameter spread in the devices being

tested

There are two approaches to reducing the variability of devices and failures due to random causes. The traditional way is

through the application of a variety of test procedures to com- pleted devices in order to screen out defective units. Although

this procedure has worked quite well for relatively simple de-- vices , it becomes less and less effective as the complexity of

the device and the variety of operating conditions to which it

may be exposed increase. As a result, this "testing-in" of re- liability is gradually being augmented or replaced by procedures

which permit reliability to be built into the device. To do

this requires that all process steps in the fabrication of the

device be carried out in a controlled and reproducible fashion.

Such control depends on the ability to determine and measure a large number of characteristics at the various process steps

This involves a large number of tests on a variety of ma- terials and structures . All the incoming materials including

the substrate wafer, process chemicals, and miscellaneous parts must be inspected and accepted. Fabrication steps, such as epitaxial deposition, photomasking , diffusion, oxidation, and

metallization, and assembly and packaging operations, such as

450
wire bonding, die bonding, and sealing, must be controlled.

Finally the finished circuits must be tested.

Development of adequate measurement methods suitable for

determining the necessary parameters during processing has also

a secondary advantage; with the use of such measurements, increased understanding of the device operation can be achieved

and improved values of various parameters can be obtained for

use with models intended to enable device response to be predicted. This is particularly important in relating the results obtained in simulated environments to performance in actual en- vironments. In addition, as such models become more refined,

more efficient screening tests can be devised.

Much testing is presently carried out at the various

producer-consumer interfaces and during the processing of semiconductor devices. Nevertheless, in a keynote speech at the

Electrochemical Society's Symposium on Silicon last year,

Willis Adcock of Texas Instruments emphasized that device production is "limited by broader than desired device parameter

distributions due to the variability of the process-caused

changes in material and device characteristics." (1)'^ In addition he noted that the area of silicon materials science which

is concerned with the answer to the question "'What effect do

the silicon processing operations .. .have on device performance?'

is in an early stage from the manufacturing viewpoint." As a solution he suggests that "quantitative correlations between the

properties of the prepared slice and the subsequent device

characteristics obtained through diffusion, interconnection, and

packaging processes must be found and exploited." He also forecasts that "the greatest impact will likely come in improvement

of the device property distribution through more effective ma- terial processing rather than in the form of the device structure." This effectively closes the circle because it is just

this narrowing of parameter distributions which is essential in

conducting effective tests in real or simulated environments

The need, then, for adequate measurement methods is apparent. Where do we stand? What are we doing to improve the situation?

First, where do we stand? Many test methods are in daily- use and a large fraction of these are written down on paper,

either as in-house procedures, as military standards, or as standards issued by organizations such as ASTM, EIA, IEEE, and

so forth. Unfortunately, users of these methods are all too

often unaware of the limitations of the procedures they are

using. Unless standard forms of the test method have been developed and unless the validity of the procedure has been demonstrated by a successful round robin experiment, these limitations are seldom documented. It is therefore of interest to

2 The numbers in parentheses refer to the list of references

appended to this paper.

451
examine the kinds of methods which are available for use in

various processing steps and the extent to which standards for

them have been developed.

The prospect of increased standardization of measurement

procedures for high-reliability of devices has improved over

the past few years because of the almost universal use of silicon as the basic material for such devices and the wide adoption

of the epitaxial, planar process for fabrication. The standard

methods which are available have been developed principally because of their need in exchanges of materials or devices across

a producer-consumer interface . As a result they cover the various requirements of microcircuit fabrication only incompletely.

The situation is best in regard to measurement methods for

characterizing the silicon substrate wafer which is the basic

starting material for monolithic circuits. Standard ASTM procedures (2) are available for most measurements, such as resistivity, conductivity type, orientation, and crystal perfection

of the wafer. With suitable care and calibration, reliable re- sults can be obtained in most cases. Nevertheless problems do

exist in this area. In his review paper which opened a session

on electrical characterization of silicon at last year's Symposium on Silicon, Paul Schumann of IBM Components Division observed that "the purchase of semiconductor materials is hampered

by improperly specified and inadequately measured parameters,

costing both vendor and customer time and money." (3) For other

important characteristics such as gross geometry, flatness, and

surface condition of the wafer, no standard procedures that have

been validated by round-robin experiments exist. Some other

parameters, such as oxygen content and carrier lifetime, are sometimes specified. Tlie standard procedures for both these

parameters are now undergoing investigation to determine if significant improvements in precision or sensitivity can be made.

Other materials such as water, acids and solvents, carrier

and doping gases, silane or silicon halides, and photoresists

are also necessary in the processing of semiconductor wafers.

The methods for determining the properties of these materials

are in various stages of development. Establishment of suitable

standard procedures is hampered because of the difficulty of detecting and identifying the very small concentrations of impurities which are significant. Of these materials, water has been

most intensively investigated. Although many of the standard

methods for characterizing water were developed for a less pure

water than is used in the electronics industry, tests appropriate to high-purity water have been developed in a few cases. In addition, standard methods are available for determining the re- sistivity of electronic-grade solvents, the concentration of

certain impurities in hydrogen peroxide, and selected properties

of photoresists. These represent only a very small beginning of

an extremely large task. Once the materials for processing are selected it is

necessary to test each step of the process to verify that it is

452
being carried out reproducibly . This involves the determination

of the characteristics of epitaxial, oxide, diffused, and metallic layers of various sizes and shapes. Because details of the

procedures employed in fabrication are frequently highly proprietary, very little standardization has occurred in the measurement of these properties. Nevertheless measurement techniques

with adequate and well-defined precision are essential whether

to permit control and improvement in the processing for proprietary reasons or to permit the manufacturer to demonstrate that

a process had been carried out reproducibly without revealing

any of its proprietary details. This is particularly important

in monolithic microcircuits because spurious interactions be- tween circuit elements can be affected by the details of the

fabrication process. Thus even devices which have the same electrical characteristics when measured under a given set of test conditions might perform differently in circuits whei-e the

conditions differ from test conditions

Because minute particles can cause flav7S (such as pinholes

in thin films of oxides and metals) in the finished circuit,

control of cleanness and particulate contamination during wafer

processing is particularly important. Standard methods and

specifications for clean room areas are available and are widely

used and accepted. It is frequently necessary to control such

other environmental factors as ambient temperature and moisture

content at a specific temperature. No standard test methods

appropriate to the control of these factors in microelectronics

processing are now available.

The principal causes of failures which occur in high-reliability devices are related to problems encountered during

assembly and packaging. As with fabrication, properties of incoming parts and materials must be determined, particulate con- tamination must be avoided, and the processes themselves must be

controlled. Standard procedures for testing incoming parts are virtually nonexistent. Extension of general testing procedures

is made difficult because of the small size and cleanness re- quired of the parts . Particulate contamination can be monitored

by the tests mentioned above. Control of assembly and packaging

procedures is accomplished principally through visual inspections and environmental, mechanical, and electrical tests such

as those found in MIL-STD-750 for discrete devices and

MIL-STD-883 for microelectronics. Though these and similar

tests are widely used, the procedures in most cases are quite

general and considerable variations in test conditions can be

expected to occur as a result of uncontrolled factors. Many of

the electrical tests in these documents are applicable to testing of finished devices. Other standard methods for these

electrical tests have been prepared by IEEE and EIA Committees.

One can conclude that although many standard test methods

are available , much work remains to be done in nearly every area

of semiconductor device processing. Now let us retijrn to the

question what is being done to attack this problem.

453
The ASTM is probably the leader in the development of stan- dard testing methods which have been validated by round-robin

experiment. The primary responsibility for developing ASTM

standards in the semiconductor device field rests with Committee

F-1 on Materials for Electron Devices and Microelectronics.

Most of the ASTM documents mentioned previously originated with

this committee. At the present time Committee F-1 has more than

70 additional documents in various stages of development as standards. These include methods for determining the characteristics of substrate wafers, the purity of various processing

materials , the characteristics of photoresists and photomasks

and the electrical and crystalline properties of epitaxial

layers. In addition test methods for evaluating both thin me- tallic films and the thick films used in hybrid microcircuits

are under development. As with most other ASTM standards, the test methods developed by this Committee benefit from, both user and producer participation and thorough evaluation before adoption . The IEEE has recently reactivated its Electron Devices

Standards Committee and is in the process of establishing task

groups to work in a large number of areas of semiconductor device characterization. The Society of Automotive Engineers has

a group which is working on specifications for certain electronic materials and processes. The International Society f6r

Hybrid Microelectronics is also concerned with standardization

activities but is relying on ASTM Committee F-1 for most test method development appropriate to this area.

The Electronic Industries Association is extremely active

in the standardization of test methods for semiconductor devices . The various committees of this organization work closely

with the military in the development and review of test methods

to be incorporated in military standards. Unlike the committees

concerned with discrete devices which are restricted to producer members, the committees concerned with microelectronics now

have both user and producer representation. A number of these

committees have undertaken round-robin testing of methods under

development

The Electronic Technology Division of the National Bureau

of Standards is also assisting many of these standardizing

committees as well as other government agencies in the develop- ment of the technology necessary for the establishment of new

or improved standard measurement methods. This work is being

carried out under the joint sponsorship of the National Bureau

of Standards and several other government agencies (U).

From these various activities can be expected to come a set of measurement methods which will assist in the production of

better characterized semiconductor devices with narrower parameter spread. The availability of such devices will assist in

relieving one of the problems associated with testing high-reliability electronic components in real or simulated environments . - V

.

454
(1) Adcock, W. A. in Semiconductor Silicon^ R. R. Haberecht

and E. L. Kern, Eds., The Electrochemical Society,

New York, 1969 , pp. 36-5^+. (2) A listing of ASTM standards pertinent to silicon and silicon processing is given in Bullis , W. M. , and Scace , R. I., PvoG. lEEE^ Vol. 57, No. 9, September 1969, pp. 1639-1546.

Most of these standards are published in Part 8 of the

Book of ASTM Standards which appears anually in November.

(3) Schumann, P. A., in Semiconductor Silicon, R. R. Haberecht

and E. L. Kern, Eds., The Electrochemical Society,

New York, 1959, pp. 662-592.

(4) Joint Program on Methods of Measurement for Semiconductor

Materials, Process Control, and Devices, supported by the

National Bureau of Standards , the Defense Atomic Support

Agency, the U. S. Navy Strategic Systems Project Office,

the U. S. Navy Electronic Systems Command, and the National

Aeronautics and Space Administration. For the most recent

report on this work see Nat. Bur. Stand. (U.S.), Tech. Note 527, Issued May 1970.

455

Paper No. 32

LIFE TESTING USING CONTINUOUS ACCELERATION

1 o Rabinowicz and B. Shiralkar'^

REFERENCE: Rabinowicz, E. and Shiralkar, B. "Life Testing

Using Continuous Acceleration" ASTM/IES/AIAA, Space

Simulation Conference, 14-16 September, 19 70.

ABSTRACT: One method of carrying out accelerated tests consists of increasing the stress at a uniform rate with time

until failure. It is sho\im that from such tests the life

under normal stress conditions can be predicted by plotting

the results on a stress- time curve on which the stress axis

has been modified so that area represents amount of damage.

Tests with electric light bulbs, xvhose failure obeys a cum.ulative damage criterion, give good agreement bet^aeen predicted and measured life. The danger of the continuous

acceleration procedure is that the stress may become so high

that new failure modes are introduced.

KEY WORDS: life testing, accelerated testing, mechanical

reliability, cumulative damage, failure.

Department of Mechanical Engineering, Massachusetts

Institute of Technology, Cam.bridge, Mass.

2 Nuclear Energy Division, General Electric Co., San Jose California

457
Introduction

The concept of continuous acceleration was introduced

into life testing by Prot (1) in 1948, v;lio detemined the

fatigue limit of mechanical specimens, not in the classic manner by running tests at constant stress levels, but

rather by a procedure v/hich involved starting at a low stress

level and then increasing the tress continuously with time,

until failure occurred. Tests at a number of different

acceleration rates v/ere carried out (Fig. 1), and then a mathematical procedure made it possible to deduce the fatigue

limit from the measured failure points. The Prot procedure

has been tested many times (2) - (8) , and its theory has

also been frequently evaluated (9) - (14) . Since that time two additiona] methods of running

continuous accelerated tests have been developed. Fig. 2 shows a situation in which a specimen is first tested under

a normal stress for a certain length of time which corresponds

to its required life. After that time, the stress is

accelerated at various rates until failure occurs, and from

the location of the failure points the total life under

normal stress conditions is estimated. This method is

frequently used in situations in which government or other

specifications require that a mechanism must run for a certain length of time on a test stand under constant operating conditions (the operational rectangle) . Once the

specimen has met the test, merely keeping the specimen run- ning at normal stress is v/asteful of valuable test facilities,

while merely removing it from the test stand gives no information about how close to failure the specimen was at the

end of the operational rectangle. Clearly, a change in

operation characteristics to a high stress level is indicated, and a simple procedure for doing this is to raise

the stress on the mechanism, continuously until failure is experienced. An alternative procedure uses specimens which are operated for part of the operational rectangle under nornial stress conditions and then acceleration is applied at a con- stant rate until failure occurs (Fig. 3). This method

is especially useful v;hen a device is operated under field

conditions for a certain length of time and is then brought

back to the laboratory for testing to failure. To save

time on the test stand, it is helpful to apply acceleration

The numbers in parentheses refer to the list of references

appended to this paper.

458
procedures right from the start. There is some discussion in the literature on whether

acceleration must be carried out in a continuous fashion or whether it may be brought about by small but finite steps.

However, no difference between these two procedures seems

ever to have been demonstrated.

In the testing carried out in our program, the third

type of continuous acceleration, that based on Fig. 3, was

used, but in fact a limited amount of experimentation was carried out using the acceleration method of Fig. 2, and it seems that our procedures and modes of analysis are equally applicable to these methods of acceleration. Our

study was part of a larger effort, part of which has been

published (15) , in which various methods of carrying out accelerated life tests were investigated, all the methods

being based on the assumption that failure occurs as a result of cumulative damage (16) . The cumulative damage

criterion assumes that if a specimen is tested for a time

under a stress at which its life would have been L]^, and

then for a time t2 under a stress at which its life would

have been L2, and so on, then if failure occurs at the n^^

stress, we have

Theoretical Analysis

Accelerated testing procedures described in the literature generally start with the assumption that the effect

on life of increasing the stress is known, i.e., it is assumed that the acceleration function is known. However,

it was the aim of our study to develop accelerated test procedures which did not involve such assiomption but rather,

to develop empirical information on the influence of ac- celeration on life, and then to apply this information in as restricted way as possible to the analysis of the continuous

acceleration procedure.

An important aid in comprehending accelerated life

tests is the constant damage diagram. This is a diagram,

rather analogous to the pressure-velocity plot used in

thermodynamics, in which area under the curve represents the amount of damage. The ordinate of such a plot is a suitable

function of the stress, while the abscissa is the time of

operation of the mechanism.

The procedure for obtaining such a diagram is shown in

Figs. 4 and 5. In Fig. 4 are shown the results of running

life tests at various constant stress levels oq to 03. The

times to failure are Lq to L^. To lay out a vertical scale

which obeys the constant damage relationship, any hyperbola

459
obeying the relationship

xy= constant

is drawn, as in Tip,. 5, and then points are found on the

vertical axis of Fig. 5 so that the horizontal distance from

the axis to the hyperbola matches the parameters Lq to L^,

and these points correspond to the transformed values of'oQ

to 03. Once the positions for the stresses Oq, etc. are found, other values can be inserted by interpolation. It remains to shov; that the diagram, we have obtained is a constant damage diagram.

It is clear from Fig. 5, that the operational rectangle

for specimens tested at constant stress oq and are equal

in area, since this is a property of the hyperbola, and thus

if constant stress testing is undertaken, equal am.ount of

damage yields equal area on the diagram. Nov/ assume that

testing is carried out under a more complex relationship

until failure occurs (Fig. 6). The continuous curve can be

approximated by a discontinuous series of steps, of length

to to t3. The total area A,, under the curve is then given

by

At = Et.f (o) ~ ^2) But it is Imown from the construction of the constant damage diagram that constant

f(a) =

L .

' (3)

Hence, equation 2 becom.es merely

A^ = constant X and since the summied term is constant according to equation

1, so is A|-

The main advantage of constructing a constant damage

diagram rather than performing an equivalent mathematical

manipulation for computing the amount of damage are tv^ofold. First, a graphical visualization of damage is obtained and

in some circumstances this can be very helpful. Second,

it becomes possible to cstim.ate the. amount of damage v;hen the theoretical function of damage against tim.e is not known

or is too complex, but failure times for various stress

levels are knovm.

Experimicntal Program

The experimental testing was carried out with light

bulbs. These were chosen from aoiong several devices that were initially evaluated because of the. following properties

A. It is easy to accelerate the damage rate in a continuous fashion merely be increasing the voltage.

B. As v/ill be seen below, it is found that the damage

mechanism remains the same over a v/ide range of

460
voltages

C. Tlie life of commercial light bulbs is larj',ely uniform

witli the standard deviation of life values less than one half of the mean life value.

D. The test specimens are very cheap and detection of

the failure point is an extremely simple matter.

The eventual method arrived at for testing the light

bulbs consisted of connecting them in parallel to the output

of a Variac voltage transformer. The total current passing

through all the bulbs was recorded continuously as a function

of time, and the reduction in current which occurred when

any bulb failed was directly detected.

Results

The first property studied was the statistical distribution of failure times for the light bulbs. Tlie laean life

of bulbs tested at a voltage of 110 volts was found to be

1190 hours, and the various lives v;ere distributed about

the mean in a normal distribution with a standard deviation

of about 500 hours. These param.eters agree v;ell with

published data (17) . A second batch of rather different bulbs

gave a mean life of 2050 hours. These m.ean lives were too long for our program and hence it was decided to operate at

a standard voltage of 130 volts, i.e., it was hypotliesized

that these bulbs had been manufactured for use at 130 volts.

The mean life under these conditions was 245 liours (Fig. 7).

hife tests on samples of 20 l^ulbs were then carried out

at voltages of 145, 165, 190, 210, and 290 volts. At the

higher voltages the life is drastically shortened and also

the distribution of lives changes towards a bimodel shape

suggesting that two different failure modes are operative

(Fig. 8). These two failure mechanisms may be thinning and

breaking of the tungsten wire (as at lo\;er voltages) and

detachment of the wire from its binding post (a nevj mechanism of failure)

The life results at various elevated voltages were used

to lay off the vertical axis of the constant damage diagram

shown in Fig. 9, and this diagram, was used to plot the

subsequent experimental data. To carry out the accelerated tests, the procedure shovm

in Fig. 10 was employed. It v;as assumed that the operational rectangle consisted of service at 130 volts for 120 hours,

and the acceleration procedures consisted of increasing the

voltage at a rate of 4.3 volts per 4 hours. In fact, two different acceleration procedures were employed- in one case

the voltage was increased every A hours while in the other

case the voltage i-jas increased three times as much every

12 hours. However, these produced no noticeable differences

and the results have been combined. The failure points

obtained are indicated in Fig. 10 and the voltage-time

461
curves are replotted on the constant damage diagram of Fig. 9. It is now required to estimate the life of the bulbs

under nonaccelerated conditions, or more specifically, to see

if a method can be found for estimating the true life under

nonaccelerated conditions. The procedure is a relatively

simple one, and it involves the use of at least two accelerated testing modes (say B and C) . It is assumed that the

position of the curves at high voltages is well established,

but that the low voltage values are quite uncertain, since

life tests at voltage of 130 volts have not been carried out. This state of affairs is represented on Fig. 11. On comparing curves B and C on Fig. 11, it will be seen

that they are identical in area, except that one has a region

P and the other a region Q. If these areas are equated, then

an estimate of the true position of the 130 volt line becomes

available. Now compare curves C and D. These are identical

except that one has region R and the other has region S. On equating the areas of these curves, an estimate of the position of point D becomes available. This procedure may be

repeated by comparing the other pairs of curves, namely, A

and B and then A and C. Table 1 lists the value of the estimated life at a constant voltage of 130 volts obtained by comparing the curves A to C, also the directly measured life at 130 volts

(curve D) . It will be seen that the various estimates

agree well.

It is of interest to note that before the use of the constant damage diagram was developed an attempt was made to use the results shown in Fig. 10 to deduce the life at the

normal stress of 130 volts by various extrapolation procedures. For instance, Fig. 12 shows a plot of the log of the

maximum voltage reached as a function of the log of the time

of failure. It was hoped that a plot such as this might

yield a straight line relationship thus allowing empirical

extrapolation procedures to be used, but in fact no linear

relationship was obtained. Many other plotting procedures were similarly unsuccessful.

Discussion

It has been the aim of this paper to show that a useful analysis of continuous accelerated life tests can be

produced by first establishing a constant damage diagram and

then analyzing accelerated life tests by means of such a diagram. It is not necessary that the acceleration function

be known because during the initial stages of the testing

program, a few points of the acceleration diagram can be

experimentally determined, and can then be used.

As a matter of interest, we have found that the failure

of light bulbs occurs by a process which follows the

Arrhenius equation

462
rate = Ae'^/^^^ (5)

v/here 0 is the activation energy of the process

R is the gas constant

T is the absolute temperature

A is a constant

Tliis equation raust be modified to allow for the fact that for

a wire heated electrically and cooled by radiation

Heat in a Heat out a and substituting this into Arrhenius equation gives rate = Ae -'''^ W"

life = Se ^ (6)

Thus, the ordinate in Fig. 9 is in fact proportional to e300/v V and from the constant of 300, we may deduce an energy of activation of 120 kcal/mole. Tliis figure is not

out of line with activation energies for creep or evaporation

of a strong metal like tungsten.

It may be of interest to compare the results of this

investigation vrith earlier work. Prot assumed that a plot

of alternating stress against testing time would be a hyperbola. He then established that such a diagram has constant damage characteristics and used this fact to devise his plotting procedure. Hoxvrever, since the fatigue

life of materials at stress levels betxsreen the yield strength

and the fatigue limit follows a relationship of the type:

life a (7) rather than the hyperbolic relationship

life a (8) It is clear that 'T'rot's assumption is not a satisfactory' one. One of the advantages of the constant dam.age diagram is that it gives a realistic appreciation of how drastically

accelerated testing raises the level of damage above the

norm.al one. Anyone seeing the high spikes of Fig. 9 must

wonder whether continuous acceleration has not taken the

process into a region in v/hich different failure laws apply. One of the interesting findings of our study is that

during accelerating testing, the distribution of failure

times is not normal (Fig. 13a) but rather the long life part

of the histogram is abruptly trupkated . The data of Fig. 13 may be replotted in terms of the area under the constant

damage curve rather than the time to failure, and the

resultant relationship is much closer to being a normal one

(Fig. 13b). This is perhaps an additional benefit of

working with the constant damage diagram.

463
It should perhaps be pointed out that in all our tests,

the average time to failure, which is the quantities used in

plotting figures (9) -(12) is the mean value where the

distribution is highly skev7ed. The constant damage diagram of Fig. 9 as well as the

simple stress-time plot of Fig. 10 from which it was

derived, differ from some of the plots used in thermodynamics

in that the position of a failure point depends on the path

by which that point v/as reached. To illustrate this point, we show in Fig. 14 the results of tests on small electric

motors, which were operated either unloaded or under an ap- plied torque. Twelve motors or more were run at each loading condition. Two curves are shoxm, the higher one indicating the stalling load for units operated in the unloaded

condition until the time indicated, the lower curve indicating the failure curve for units operated continuously under

a uniform torque. In diagonal testing, as shown, the failure

locus falls between the tv;o previous curves, as might have

been expected. A number of previous authors had not emphasized this influence of the mode of testing on the location of the failure point (6) , (7) , (11)

Perhaps an evaluation of the v/hole procedure of

continuous acceleration is in order. The concept by which

we keep raising the stress to 'punish' a specimen which

has been a long time on the test stand, is psychologically a very attractive procedure, and perhaps the wide attention

which Prot's procedure has received is partly attributable

to this factor. However, there is a certain inefficiency

in having a modest initial accelerated stress, and then

compensating for this by having a very high final stress

level. The alternative, two-step approach (15), is less

likely to get the test into a realm v/here different modes of

failure are brought into play, as a comparison of the two modes of testing shows (Fig. 15).

Acknowledgements

The authors x>7ish to thank the Pratt and Whitney Division

of the United Aircraft Company for sponsorship of this v/ork,

and Px. Shevchenko, N. Gordon, P.F. Brovm, K.A. Domeisen, and

E.P. Saling, Jr., of that company for their advice and in- terest. Drs . P. Could and C.S. Reichenbach have made many

helpful suggestions.

464
References

1. E.M. Prot, Fatigue Testing Under Progressive Loading, A New Technique for Testing Materials, Revue de Metallurgie,

45, 481-489, 1948. (Translated by E.J. Ward, WADC-TR-52-

148, AD-2176, 1952).

2. F.H. Vitovec and G.J. Lazar, Strength, Damping and Elasticity of Materials under Increasing Reversed Stress with

Reference to Accelerated Fatigue Testing, Proc. ASTM,55

844-865, 1955.

3. N. Enomoto, On Fatigue Tests Under Progressive Stress,

Proc. ASTM, 55, 903-917, 1955.

4. L.S. Lazar, Accelerated Fatigue of Plastics, ASTM Bulletin

220, 67-72, 1957.

5. N. Enomoto, A Method for Determining the Fatigue Limit of

Metals by Means of Sepwise Load Increase Test, Proc. ASMM,

59, 1959.

6. J.J. Bussolini and M.J. Ciarlariello , "Step Stress Testing

to Failure for Reliability Analysis of Electronic Equipment," pp. 432-441, Aerospace Reliability and Maintainability Conference, SAE, New York, 1964.

7. G.E. Best, G.R.Bretts, H.T. McLean and M.M. Lampert,

Determination and Application of Agii^ Mechanisms Data in

Accelerated Testing of Selected Semiconductors, Capacitors,

and Resistors, pp. 293-302, Proc. Eleventh Symp . Reliability and Quality Control, 1965.

8. M.N. Stepnov and Y.A. Shukhmin, Use of the Prot Method

ill the Fatigue Testing of Light Allcyn , Industrial

Laboratory, 33, 1018-1020, 1967.

9. W.A. Hijab, A Statistical Appraisal of the Prot Method for

Determination of Fatigue Enducrance Limit, J. Applied

Mechanics, 24, 214-218, 1957.

10. H.S. Endicott and J. A. Zoellner, A Preliminary Investigation of the Steady and Progressive Stress Testing of

i4ica Capacitors, pp. 229-240, Proc. Seventh Symp. Reliability and Quality Control, 1961.

11. G.A. Dodson and 5.T. Howard, High Stress Aging to Failure

of Semiconductor Devices, pp. 262-272 of Proc. Seventh

Symp. Reliability and Quality Control, 1961.

12. R.R. Gatts, Cumulative Fatigue Damage with Progressive

Loading ASME Paper 62-WA-292, 1962.

13. G.E. Best, G.R. Bretts and H.M. Lampert, Physics of

Failure and Accelerated Testing, Electro-Technology,

81-92, Oct. 1965.

465
14. G.A. Deryagin, U.K. Shtovba, E.I. Shneerova, Use of Rapid

ilethods in Determining the Fatigue Strength of D 16 T Aluminum Alloy, Industrial Laboratory, 3^, 1021-1024,

1967.

15. E. Rabinowicz, R.Il. McEntire and B. Shiralkar, A Technique

for Accelerated Life Testing, ASIIE Paper 70-Prod--10,

19 70. 16. M.A. Miner, Cumulative Damage in Fatigue, J. Applied

Mechanics 12, A159-A164, 1945.

TABLE 1. Predicted Life of Bulbs at 130 Volts

Source of estimate Life

Figure 9,

Figure 9, Figure 9

Figure 7,

Curves B and C Curves A and C Curves A and B

210 hours

Direct measurement

220

310

245

466
Time

Opera! lona

rect ang le Time

Fig. 1- -Schematic illustration of the Prot method

of accelerated testing.

Fig. 2 --An acceleration method

especially applicable to devices

which must first meet a specified duty cycle.

Fig, 3 --An acceleration

method to be used on devi- ces which have used up part

of their lives in field tests.

Fig, 4- -Typical plot of time

to failure as a function of constant test stress.

467
Time to failure Fig. 5- -Method of obtaining a constant damage stress scale

from the data of fig. 4.

Failure

poi nt

Time Fig, 6- -A complex curve on

a constant damage diagram

may be approximated by a number of rectangular steps.

hyperbola generating constant

area diogrom

Life 3CC hours

o

0 30

Life

Fig. 7- -Histogram for 60

light bulbs tested to failure

at 130 volts- The distribution is nearly normal.

Fig. 8 --Histogram for 20 light

bulbs tested to failure at 290

volts . The shape of the diagram resembles that of fig. 7, but there are some very early

failures.

468
200

150

^ 100

Hypothetical

/

operational /J

path D

100 150 200 250

Life in hours

10 --Plot of the stress -time functions of light bulbs tested under continuous acceleration. Thirty specimens were

failed along each path, and each median failure point is shown.

Fig.

469
50 IOC 150 200

Life in hours

Fig, 11- -Method of computing the life at 130 volts . P and Q, and then areas R and S, are equated.

250

Areas

1 727

130V

100

Time to foilure.

OD

300

log scale Fig. 12- -Failure points from

fig. 10 replotted on a log -log

scale . Points A, B, and C do

not readily extrapolate to D.

10

(a) 80 82

Life 84 86

in hours 10

- 5

(b) I 8

Area

88 90

9 10 I I under curve Fig. 13 --Histogram of lives of 30 bulbs tested along path B of

fig. 10 a) Life given in hours

b)Life given as area under

curve B of fig. 9.

470
0 100 200 300 400 500 600

LIFE IN HOURS

Fig. 14- -Stalling curve, life curve at constant torque, and

life data under continuous acceleration for electric motors.

The stress -life function depends on the testing history.

o

uCOo

<vO

co

V)coo

t/f to<v in

Time to failure

Fig. 15- -Comparison of two-step and continuous acceleration testing . The full curve and the broken curve have the

same area, but the broken curve (two-step testing) does not

reach as high stress values.

471

Paper No. 33

ESTABLISHMENT OF M OPTIMUM DURATION FOR SPACECRAFT COMPONENT

THERMAL VACUUM TESTS

W. L» Harvey'

REFERENCE: Harvey, W. L., "Establishment of an Optimum

Duration for Spacecraft Component Thermal Vacuum Tests,"

ASTM/IES/AIAA Space Simulation Conference, 14-16 September

1970.

ABSTRACT: The duration of planetary space missions precludes

performing component thermal vacuum tests which simulate the

mission time period. Many arbitrary approaches have been

taken to establish adequate durations for these tests. This

paper outlines an approach which considers the probable effects

of the thermal vacuum environment on the components and their

constituent materials. In general these effects are found to be related to outgassing rate which is time and temperature

dependent. A test cycle is proposed which considers this

situation and utilizes any prior "bake out" heat exposure to reduce the test duration to a reasonable period.

KEYWORDS: Thermal vacuum tests, components, simulation,

environment, volatile constituents, outgassing, degassing,

"bake out", thermal conduction, evaporation, sublimation,

volatilization, decomposition, absorption, adsorption.

INTRODUCTION

Long duration space missions propose a unique problem to the engineer charged with establishing an appropriate test program to qualify and flight accept the spacecraft

components. Since it is obviously not feasible to extend

thermal vacuum test durations to coincide with a multi-month

Staff Engineer, Martin Marietta Corporation, Denver, Colorado.

473
Interplanetary cruise or orbital period, some compromise must

be achieved

A review of government and industry efforts at achieving

this (as reflected in OGO, Mariner, Apollo, and other spacecraft test requirements) has indicated the absence of any

consistent rationale in this area. Therefore, the purpose of

this paper is to propose an approach to resolving this potentially costly and time consuming problem. Basically, the

problem can be approached from two directions.

The first (and perhaps most reassuring) would be to correlate failure rates with time in the thermal vacuum environment

from extensive ground tests. Unfortunately, this data is not readily available and, where available, is not in such a form

that positive conclusions can be drawn from it. Such factors

as component functional state, prior test history, and thermal

cycling conditions have been observed to have a very pertinent

effect on the occurrence of component failures under thermal vacuum conditions. For example, a component may be exposed to several days thermal vacuum conditioning while completely non- operative. It is then turned on and fails shortly thereafter.

Was its failure a function of time in the chamber or initial

operation under the thermal vacuum conditions? Similarly, a component may progress through a series of environments, including launch vibration, which simulates the spacecraft

mission. It fails during simulated cruise thermal vacuum

conditions. Should its failure be attributed to stresses

induced by the thermal vacuum environment or to residual stress from the prior vibration testing?

The second approach involves restating the problem:

"Establish a test (duration) which simulates the effects of

the mission thermal vacuum environment rather than the precise

environment itself." A step has already been taken in this

direction with the realization that the hard vacuum of space

(10-14 torr or less) cannot be readily simulated for normal

component testing. Rather 10"^ to 10"^ torr^ is generally

Several rationales exist for supporting this compromise

rather than the fact that it is exceedingly difficult if not impossible to provide a system for testing relatively dirty

components at pressures much less than this. For example, at 10"^ torr the mean free path of gas molecules becomes several meters. Thus, gas molecules will be returned to the test surface only by condensing on the chamber walls and then re- subliming. This also means that the thermal conduction of the gas is quite small relative to radiant heat transfer. At 10"^ torr the dielectric strength of the chamber atmosphere has increased to the point where operating voltages in the kilovolt

range may be applied without danger of arc or glow discharged > Even at 10"^ torr, however, a monomolecular absorbed gas layer

may be built up on an initially clean surface in 2 seconds. (2) 474
considered adequate to simulate the effects and is nominally

required for this type of testing.

In order to establish a test which adequately simulates

the effects of the environment, it is necessary to determine

what these are. Not surprisingly, these effects are directly

related to the two prime characteristics of the space environment, i.e., temperature extremes and the absence of an atmosphere. The synerogistic effect (absence of a convective

atmosphere) appears to be minimal and definitely biased to- wards altering the effects of the thermal extremes. Consequently, it would appear that the primary effects of the space

environment may be categorized as either thermal or vacuum

related.

THERMAL EFFECTS

The thermal effects are essentially the same as those

produced by any temperature condition. High temperature

effects result in alteration of materials characteristics and

include softening which may be either reversible or irreversible, loss of tensile strength or evaporation of volatile

constituents as well as differential displacement which may

create a physical stress. Low temperatures cause loss of

resiliency or ductility and increased brittleness in addition

to developing physical stress. While each of these effects is significant, two factors should be considered. First, each

effect, with the exception of differential displacement, can

be studied and evaluated at the materials level. The effects

of differential displacement, however, can only be evaluated

effectively at the component assembly level. Second, loss of

resilience including dissolution of plasticizers can be

produced under ambient pressure conditions as well as under

vacuum.

VACUUM RELATED EFFECTS

The remaining thermal effect, evaporation of volatiles,

leads directly to consideration of the vacuum related effects

of the space environment. A review of references on materials

for space usage indicates that exposure to space vacuum may

produce essentially two categories of effects. First, and

most prevalent, is weight loss through evaporation, sublimation, or volatilization of the material or its molecular

fractions following decomposition or dissociation. The second

effect, which is less common and certainly less understood, is

The numbers in parenthesis refer to the list of references

appended to this paper.

475
a direct change in mechanical properties without significant

weight loss

The evaporation process is generally associated with the boiloff of entrapped water vapor or other residual fluids.

Sublimation may occur where low molecular weight fractions

exist in the polymer. Volatilization occurs where volatile

constituents have much higher vapor pressures than the cured

polymer. These volatile constituents may include: (4)

Condensed reaction products from the curing process

Processing additives

Stabilizers

Residual solvents

Thermal or radiation induced decomposition or deploymerization products

Unreacted curing agents ^ Impurities ; In all of the situations the resultant rate of weight los

through outgassing is directly proportional to the vapor press

ure of the outgassing materials as indicated by the KnudsenLangmuir equation.

The total effect of weight loss on the engineering properties of the material presents another area of inconclusiveness. It is generally agreed that weight losses of 1 to 2 percent do not produce property changes of engineering importance unless large volumes of the material are present.

In the latter case, the condensable volatiles produced by even

a small weight loss of the bulk material might affect the

performance of some other element of the system,

Inorganic materials at nominal temperatures generally

appear not to be adversely affected by vacuum conditions. In

fact, electrical properties of inorganic dielectric materials

may be enhanced through the boiloff of water. Changes in mechanical properties of polymers not accom- panied by a weight loss include variations in elongation

characteristics, tensile strength, hardness and elasticity.

These poorly understood effects are generally believed to be a function of the relative rates of molecular cross linking and

cleavage. Most sources, however, seem to feel that some weight loss must occur before vacuum conditions will have a significant detrimental effect on mechanical or electrical

properties. Consequently, most materials studies have involved themselves with measuring weight loss or cataloging

volatile condensable materials . ^"^-^ For the purposes of analyz

ing and solving the problem of optimum component thermal vacuum test durations, weight loss and resultant degradation

of engineering properties appear to be the prime parameters of concern.

476
THERMAL VACUUM EFFECTS ON COMPONENTS

Having considered the potential effects of the thermal vacuum environment on constituent materials it becomes neces- sary to determine the probable way in which these effects

could manifest themselves at the component level. These include the following:

1) Physical failure as the result of mechanical degradation through outgassing and/or the introduction of stress from differential thermal displacement .

2) Degradation of lubricating properties (destruction of

absorbed or chemisorbed oxide films with no healing

possible) leading to cold welding, galling or seizure of bearing surfaces. (3)

3) Corona and arcing resulting from the loss of dielectric strength of an insulating material or through the

development of a conductive atmosphere immediately

adjacent to an outgassing source. (^j^)

4) Contamination of critical surfaces through deposition

of outgassed products . ^"^ '^^ At this point, the problem of establishing an optimum

thermal vacuum test becomes one of establishing a test which

adequately evaluates both the thermal and outgassing characteristics of components within a reasonable period. Temperature

cycling, with the component operated while stabilized at each extreme and monitored during the transition time, appears to provide an excellent means of producing any potential problems

related to thermal stress effects.

Outgassing, however, is time dependent. Therefore, it

would appear that the test duration should be based on the time

necessary to initiate, develop and essentially complete out- gassing. The outgassing situation at the component level

becomes a great deal more complex than it was at the materials

level. Component geometry may provide for repeated condensation and resublimation of outgassed molecules in even a well

vented "black box". Operating characteristics of the component

may cause various parts within the component to operate at different temperatures during various phases of the functional

profile. This could result in outgassing, condensation, re- sublimation, etc., occurring in a never ending cycle. While

this is an extreme example it does point up the problem of

basing a test duration on the time required to complete out- gassing. In this case, the component might never complete

outgassing . Worst case effects, however, could be evaluated

at several points during the component operating cycle . Such

a test might indeed be more severe than one wherein total

477
outgassing had been accomplished, including complete resublimation of miscellaneous autgassed products from critical

internal surfaces. Certainly if arcing and corona discharge

are potential problems, it would be better to conduct a test which evaluates the effect of outgassing while conductive gas molecules are still in the vicinity of high electrical potential sources.

OUTGASSING CHARACTERISTICS

From the foregoing it would appear that the development of

outgassing is one of the prime if not the most important

objective of the vacuum portion of a thermal vacuum test. As

has been indicated previously the actual mechanics of outgassing vary from metals to non-metals and from non-metallic

material to another. Outgassing in metals is related to the

diffusion coefficient and true surface of the metal. ^^-^ The

diffusion coefficient in turn is a function of temperature and,

in some cases, concentration of gas within the solid. In addition to the same factors, outgassing of non-metals may be related to thermal degradation or molecular decomposition.^

The complexity of the situation is such that extensive materials thermal vacuum testing has been accomplished and will

continue to be accomplished in order to select the most stable

(and low outgassing) materials for space usage.

Numerous references reflect the weight loss (outgassing)

characteristics of commonly used spacecraft materials. Metals,

wherein outgassing is primarily a function of absorbed gasses,

outgas at a log linear rate. Dayton^^'^ has indicated that

outgassing of clean metals is complete after 15 to 25 hours.

Polymers, elastomers, etc., wherein outgassing may be related

to chemical decomposition at higher temperatures also outgas at a log linear rate but at a higher initial level. Testing of

these materials, which has been accomplished over the last 10 to 12 years, has established an extensive catalog of space

qualified materials. As indicated in the references, outgassing of the low weight loss materials generally reaches equilibrium in as little as 25 to 50 hours depending on the temperature at which the evaluation is performed . (H > ^2)

Without exception, outgassing may be initially accelerated

and subsequently reduced or eliminated through the application

of heat . (^j 11 > This "bake-out" either at ambient pressure

or under vacuum conditions, is a commonly accepted method of

degassing materials.

The benefits of any such degassing process with polymers

are further enhanced in that the subsequent reductions in out- gassing appear to be permanent. Readsorption and reabsorption

of ambient atmospheric constituents is generally less than one half that originally adsorbed . '

' The weight gained by

re-exposure to the atmosphere after degassing is quickly lost

478
when, the material is again subjected to vacuum or elevated

temperatures. Figure 1 shows this phenomena for a silicone

base polymer . ^ This may be considered typical of most organic poljrmers. One means of accomplishing this beneficial preliminary

degassing occurs naturally where terminal sterilization is re- quired for planetary probes. Current planetary quarantine

constraints require exposures of up to 60 hours at 135°C.

Such exposures may increase by an order of magnitude or more

the initial outgassing rate and resultant weight loss of common aerospace polymers. (5) Thus, where preliminary heat

compatibility tests are required, a major portion of the probable outgassing would be accomplished even before thermal vacuum testing is initiated. This factor would permit an extensive reduction in the thermal vacuum test duration.

TYPICAL THEEMAL VACUUM TEST PROGRAMS

Before proceeding to establish an optimum thermal vacuum

test program a review of several major spacecraft programs

appears to be in order.

The following reflects the test durations employed during

some of these programs.

Component T/V System T/V

Program FAt'^ Qualification FAr'" Qualification

OGO 24 hrs 24 hrs 7 days 12 + days

Pioneer 24 hrs 24 hrs 14 + days 17 + days

Mariner Series 2-3 days 13 days 13 days 13 days

Surveyor 8 hrs 3 days 4 days 4 days

Nimbus, IMP, Etc. 2 days 2 days 13 days 13 days

Apollo (LM) 8 hrs 24 hrs

Lunar Orbiter 16 hrs 18 days 58.5 hrs 5 days

Discussions with the contractors involved with these

programs indicated that even though they might run component

qualification and system FAT or qualification for more than

two days, the majority, if not all, of the failures encountered

during the latter portions of the test could be related to one

or more of the following circumstances.

1) Initial operation of the component during the test (even though it had been exposed to thermal vacuum

for several days).

FAT indicates flight acceptance environment tests applied to all flight components prior to installation in the flight

spacecraft

479
2) Operation during or immediately following a temperature

cycle

3) Accumulated deposition of outgassed materials from

some bulk source (not related to a specific compo- nen t )

In general, most contractors contacted indicated that the

majority of the problems encountered during thermal vacuum

testing were temperature related with only a small portion

attributable to the vacuum environment. As an example, one contractor analyzed 134 FAT discrepancies occurring on 108

components with the following results . ('^)

46.37o - Thermal problems (drift, open circuits, etc.)

5.27o - Vacuum (primarily arcing and corona on RF equipment)

31.3% - Test error (procedure, test equipment, etc.)

17.27o - Not thermal vacuum related - primarily residual

from prior vibration test

100. 07o It miay be noted that all of the unmanned spacecraft

programs required system thermal vacuum tests to both quality

and flight accept the spacecraft. Initially it was hoped that some guidance could be obtained from these system test durations in establishing component test durations . Unfortunately, a system cannot be treated as just a "big component" in

this environment. Test durations at the system level must be

of sufficient duration to permit thermal stabilization (if

appropriate) in various operating modes. The effects of out- gassing and deposition of bulk sources must be considered.

Time must be allowed for degradation of thermal control sur- faces. All of these factors make a system thermal vacuum

simulation vastly different from that required for a component.

Component thermal stabilization may be accomplished in

hours instead of days. The effects of self generated heat

become apparent in minutes instead of hours. Bulk source out- gassing either is not a factor as the bulk source is not present or manifests itself only in relation to its parent

component

System thermal vacuum testing does provide an indirect

benefit to the engineer planning a component thermal vacuum

test. The system test establishes an additional period of vacuum exposure outgassing, and resultant materials degradation

if this is indeed a potential problem. Obviously, a system

test may be a late point in the program at which to find a component design or materials application problem. However, it

may also provide the "frosting on the cake" for those concerned

with the risk of running an optimized component thermal vacuum

test. 480
SUPPORTIVE MATERIALS TESTING

Throughout this entire analysis much reliance has been

placed on materials test data. It should go without saying

that similar reliance must be placed on actual materials testing in planning a valid component thermal vacuum test. Just as

the system test is a late point in time to find a component

design problem, so is a component test the wrong place to un- cover a materials misapplication. Ideally all materials going

into a spacecraft component will have been prequalified individually to the mission environments. This qualification can

involve actual long duration thermal vacuum tests (coupled with

ultraviolet radiation if appropriate) or cataloging of pre- viously qualified materials. The component test may then be

properly concerned with evaluating the interaction of the

materials within the specific design.

This is not to say that all materials must have a low

weight loss, non- thermal decomposition pedigree. This would

be an ideal goal. Rather it should imply that all materials

going into a specific component be identified. The presence of any unknown or suspicious materials or unusual application

should then require consideration of additional testing.

THE OPTIMUM TEST

Certain positive conclusions can be drawn from the pre- ceding discussion which should be considered in establishing

the optimum test duration for spacecraft components. These

include the following:

1) Adequate data concerning constituent materials should

be available so that some prediction of their probable

outgassing rates and time to equilibrium can be made

2) Any prior thermal exposure should be considered for

3) The component should be operated and monitored during

the pump down period as a means of detecting any

potential conditions of arcing or corona. (This is a more effective evaluation than depending on component

outgassing to establish potential conditions for this

phenomena .

4) Chamber pressure should be reduced to 10 ~' torr or less to assure that all heat transfer occurs through

gases

481
radiation and none through convection.

If no specific temperature cycle is required to simulate mission conditions the component should be exposed to the maximum test temperature initially for as long a period as reasonably possible. (A period of

12 to 24 hours would further promote outgassing and if

accomplished following a prior high heat test should

result in the earlier establishment of an equilibrium

outgassing rate.)

The component should be operated near the conclusion

of the high temperature exposure and its operation

continued during reduction to the minimum test temperature. This will permit determination of any discontinuities over the total temperature range.

The transfer from high to low temperature should be

accomplished at as high a rate as possible consistent

with mission parameters. This action will promote

any stresses developed through differential thermal

displacement

Time at the low temperature should be minimized to that required to stabilize the component. (Outgassing or other degradation is retarded at low temperatures in constrast with the action at higher temperatures .

Transfer from low to high temperature should also be accomplished as rapidly as possible. (This action

will promote any non-linear stresses not developed

during the high to low temperature transfer.)

The foregoing temperature cycle should be repeated

until the accrued time under high temperature conditions is sufficient to assure that the rate of out- gassing for all constituent materials has reached

equilibrium.

Or, if this is not predictable, the test should be

extended to a minimum of 72 to 144 hours (four to eight cycles with nominal stabilization times)

During this period most common materials will have

achieved equilibrium outgassing 40 to 80 hours

earlier. As a result the test duration will provide

a margin over uncertainty of two or three times.

Operation during or following chamber repressurization

need not be required unless this is necessary to simulate the mission. (Some solid state devices have

482
been known to develop an unexplained failure mode

under these conditions.)

The test described here would provide adequate evaluation of

component characteristics during the design development phase

or could be used to demonstrate design maturity during qualification. Based on the data derived during these test programs

the same or a shorter test could be used for flight acceptance.

If no anomalous conditions attributable to vacuum produced

effects are detected during either of these tests, consideration could be given to either a shorter test (24 to 72 hours - two to four cycles) or to a temperature cycling test under

ambient atmospheric pressure conditions.

CONCLUSIONS

Based on this study, it now appears that the factors to

be considered in establishing an optimum duration for component

thermal vacuum testing should not be based totally on the

actual mission environment itself. Rather, the factors of

component design, materials used, prior environmental conditioning, and potential failure modes all should be considered. The

test should be biased toward as complete evaluation of component

characteristics early in the test profile as possible under the mission constraints. If It is desirable to establish a broad

test to envelop a complete spacecraft family of components, it

would appear that this could be based on consideration of the

criteria used to select constituent parts and materials plus

any prior thermal conditioning encountered as the result of

planetary quarantine constraints.

483
APPENDIX

Definitions

Component An assemblage of parts and materials which

perform a specific function within a spacecraft

system, i.e., a "black box."

System A complete spacecraft element capable of

accomplishing a major phase of the mission

either separately or in conjunction with other

systems

Outgassing Spontaneous evolution of gas from a material in

a vacuum. Degassing Deliberate removal of gas from a material,

usually by heat under vacuum. Adsorption Condensation of gas on the surface of a solid.

Absorption Binding of a gas in the interior of a solid.

0 I « i > . . 1 0 8 16 ''A 32 40 48 Hours

Fig. 1. Readsorption study of Sylgard 182.

484
REFERENCES

(1) Santeler, D. J. et al, Vacuum Technology and Space Simulation , NASA SP-105, 1966, pg . 13, Chapter 9, pp. 223-229.

(2) Jaffe, L. D. and Rittenhouse, J. B. in Materials for

Missiles and Spacecraft , E. R. Parker, Ed., McGraw-Hill,

1963, Chapter 9.

(3) Rittenhouse, J. B. and Singletary, J. B. , Eds., Space

Materials Handbook , Wright-Patterson Air Force Base,

1968.

(4) Boundy, R. A., The Effects of the Space Environment on Materials , SAMPE, Volume 11, April 1967, pg. 194.

(5) Lee, S. and Fewell, R« , The Effects of the Space Environ- ment on Materials , SAMPE, Volume 11, April 1967, pp. 151-

156.

(6) Dodge, H. S., Annals of the Assurance Science, Eighth

Reliability and Maintainability Conference , AIAA, SAE,

ASME, July 1969, pp. 101-104.

(7) Duraetz, R. A. and Tuckerman, A. J., The Effects of the

Space Environment on Materials , SAMPE, Volume 11, April

1967, pp. 228.

(8) Dayton, B. B. , Transactions of the Eighth National Vacuum

Symposium of the American Vacuum Society , 1961, pg . 42.

(9) Dayton, B. B., Transactions of the Ninth National Vacuum

Symposium of the American Vacuum Society, 1962, pg . 293.

(10) Fulk, M. M. and Horr , K. S., Transaction of the Ninth

National Vacuum Symposium of the American Vacuum Society , 1962, pp. 324-338.

(11) Shipple, C. L. and Curtindale, E. G. , Transactions of the

Tenth National Vacuum Symposium of the American Vacuum

Society , 1963, pp. 63-68.

(12) Boebel , C., Mackie , N. and Quaintance , C, Transactions

of the Tenth National Vacuum Symposium of the American

Vacuum Society , 1963, pg. 62.

(13) Griffith, J. S,, "Some Tests for Increase in Friction of

Mechanisms of the Mariner Mars 1969 Spacecraft in the JPL

Molsink Space Simulation Chamber," Technical Report 32- 1446, JPL, Pasadena, California, September 1969.

(14) Thome, J. A. and Whipple, C. L. , The Effects of the Space

Environment on Materials , SAMPE, Volume 11, April 1967,

pp. 244-247.

(15) Vanderhorck, M. , Private Discussion, TRW, January 20,

1970.

(16) Timmims, A. R., "Time Required for an Adequate Thermal

Test of Flight Spacecraft," NASA Report TN D-4908,

National Aeronautics and Space Administration.

485

Faner No. 3U

EQUIPMENT TO STUDY HIGH VELOCITY DISLOCATIONS IN MATERIALS

1 2 V. R. Parameswaran and J. Weertman

REFERENCE: Parameswaran, V. R. and Weertman, J.: "Equipment to Study High Velocity Dislocations in Materials",

ASTM/IES/AIAA Space Simulation Conference, 14-16 September 1970.

ABSTRACT: The micros true tural and physical properties of a material deformed under high strain rates are quite different

from those of the slowly deformed material. This is attributed to the motion of a shock front consisting of an array of

high speed dislocations through the shock loaded material.

An equipment was constructed to shock load single crystals

of metals by shooting small steel pellets using a gas gun. The slip bands generated on the surface of the crystals by

the stress pulses give a measure of the dislocation mobility

in the m.aterial. In this paper the equipment used to produce

the stress pulses and that used to measure and record the stress pulses are described.

KEY WORDS: High speed dislocations. Impact stress pulse.

Piezoelectric stress gauge. Quartz, Slip bands, Dislocation

velocity.

"Research Associate, Department of Materials Science,

Northwestern University, Evanston, Illinois 60201.

'Walter P. Murphy Professor of Materials Science, and

Professor in Geology department, Northwestern University,

Evanston, Illinois 60201.

487
INTRODUCTION

Many high-energy-rate deformation techniques are being

used in aero-space industry to deform hard and refractory

materials such as tungsten, niobium, titanium and their all

oys , nickel and cobalt base alloys and high strength steels

When a material is deformed under high strain rates by one of the high velocity forming techniques such as explosive

forming, electro-hydraulic forming, electromagnetic forming

or pneumatic -mechanical forming, the plastic deformation

of the material is rendered much easier than when deformed

by conventional low-strain-rate techniques. This is due to the fact that under high strain rates, the dislocations in

the material move with very high velocities close to their

limiting velocities in the material. The limiting velocity

V, of an ordinary dislocation pushed by a fixed applied

stress lies in general in the transonic velocity range

(c ^ V ^ c , where c is the shear sound wave velocity and

c is the longitudinal sound wave velocity in the material)

For a screw dislocation, the limiting velocity is equal to V = c. If the applied stress is made large enough, the

limiting velocity of a gliding edge dislocation approaches

V = c . The stress field of a moving screw dislocation

shrinKs in the direction of motion. Because of this shrinkage, the force parallel to the slip plane between fast

moving screw dislocations in the same slip system goes to zero. The plastic flow of material is rendered easier under

high strain rates due to this behaviour of high speed

dislocations

High speed dislocations are also thought to be the

agents that bring about dif fusionless transformations such

as martensitic and bainitic transformations and twin trans- formations. Such transformations are brought about by the

motion of an array of partial dislocations moving with

sonic or supersonic velocities on an interface which can

give up energy. Eshelby (1) has given the structure of such

supersonic dislocations. Smith(2) has given the structure

of the dislocation interface between the shocked and unshocked regions in very high pressure shock wave phenomena.

The dislocations in Smith's interface travel at supersonic

speeds with respect to both the shocked and unshocked regio

in order to stay in the interface. For a review of the

properties and behaviour of high speed dislocations, see Ref.(3-5).

The numbers in parenthesis refer to the list of references

appended to this paper.

488
In order to study the mobility of high speed dislocations

in single crystals of metals, an equipment was built to subject these crystals to impact stress pulses of short

duration of a few microseconds. The length of the resulting

slip bands on the surface of the crystals give a measure of

the dislocation mobility. The results of the velocity measurements at different temperatures in lead and lead-indium

alloys in both the normal and the superconducting states

are described elsewhere (6,7). In this paper, the details

of the equipment used are given.

A. THE GAS GIM: A schematic diagram of the equipment is shown in Fig. 1. A stress pulse is applied by shooting a round nosed steel

pellet A of diameter 3.2 mm onto an assembly consisting of

a piezoelectric stress gauge B, the single crystal C under

study, and a lead attenuator bar of length 12 cm. The assem- bly is kept in position by a perforated steel disc E of

thickness 3 mm and 4 phosphorbronze springs F. Fig. 2 shows

a photograph of the assembly. The pellets are shot out by means of a gas gun consisting of a vertical barrel G, a pressure chamber H, a solenoid valve I, and an electromagnet

J. Pressure in the chamber is indicated by the gauge K. 'A pellet is held by the electromagnet over the barrel and

the chamber is pressurized to the desired value by nitrogen

gas. (Helium gas was used in tests at very low temperatures).

Then, simultaneously, the current to the electromagnet is cut off and the solenoid valve is momentarily opened. Fig. 3 shows the electrical circuit to operate the solenoid valve

and the electromagnet. The pellet, shot out through the

barrel by the gas pressure in the chamber, strikes the top

of the piezoelectric stress gauge. The piezo-voltage developed is fed to a dual beam oscilloscope which has been

kept ready to be triggered by the pulse. X-cut quartz was

used as the piezoelectric transducer. Details of the piezoelectric gauge are given in section B. Figures 4 & 5 show

the photographs of the equipment.

The stress pulse in the form of a piezo-voltage vs time spike was photographed by a polaroid camera attached

to the oscilloscope. The piezo-voltage is directly proportional to the applied stress. The dual beam scope used was Tektronix 555 with vertical plug- in units of the type L and

1A2 respectively, for the upper and the lower beams. The

beams were triggered using the external triggering source

position, since in the internal triggering source position,

the beams were found to be triggered by any external transients in the vicinity of the system such as that caused by

the switching of the electromagnet. A triggering circuit

489
was used to feed the piezo-voltage to the external source positions and the vertical amplifiers of the scope. Fig. 6 shows the triggering circuit used. The two beams were

traced at different speeds to study the main pulse and the

overall nature of the pulse after a longer duration.

The value of the applied shear stress was changed by

altering the propellant pressure in the chamber The length

of the steel pellets was changed to vary the duration of contact of the steel pellet with the piezo-assembly , and

hence the duration of the stress pulse. The stress pulses

applied to the crystals were essentially single pulses, as can be seen from Figs. 7 & 8. The stress pulse can be considered as a half sine wave,

the stress increasing sinusoidally with time, reaching a maximum or peak value and then decreasing to zero. A mean

stress T was calculated by dividing the area under the curve by the total time t of the pulse.

The magnitude of the stress applied by this equipment is of the order of a few kg/mm^ and the duration is of a few

microseconds.

B. THE PIEZOELECTRIC STRESS GAUGE

Many investigators (8-14) have used X-cut quartz discs

as stress measuring gauges. Most of these investigations were carried out to study the voltage and current output

obtained from quartz crystals by application of large stresses of the order of a few kilobars. Quartz has been

extensively studied and it has been found that the piezoelectric strain coefficient dn remains nearly constant

from 4.2°K upto 473^ (15,16). Also, d-^-^ is practically

independent of stress. The dielectric constant of quartz

is large enough so that edge corrections do not have to be

made for bars or plates that are reasonably thin in the

field direction. The crushing strength of quartz is about

300 - 400 kg/mm and so quartz can withstand high stresses

in compression. Due to these advantages of quartz, we decided to use X-cut quartz crystals received from M/S Hunt Corporation, Carlisle, Pennsylvania.

The quartz discs received were of diameter 1.25 cm

and thickness 3.2 mm, with the two opposite faces lapped

and polished to be parallel to each other, within 0.0001 cm. These faces were plated with a vapour deposited layer of

silver of thickness of about 0.0075 cm. Fig. 9 shows the

details of the piezoelectric stress gauge. The electodes

(1)

490
attached to the quartz crystal are made of pure copper. On top of the upper electrode is a lead nose piece of diameter

1.25 cm and length 3.75 cm. On top of this is a hardened

tool steel disc of diameter 1.25 cm and thickness 3.2 mm. These nose pieces make the wave front of the stress pulse

reaching the quartz crystal essentially plane. The lead nose piece also cuts down the secondary pulses due to waves reflected from the sides. Since the double transit time in

the bar supporting the piezoelectric stress gauge is much

longer than the duration of contact of the steel pellet with

the assembly, the difficulties due to wave reflections at

the impact point are eliminated. Further, since the impacting

pellet and the impacted surface were made from hardened tool

steel drill blanks , there was hardly any indentation produced

due to impact, showing the absense of plastic flow of

material.

For bonding the quartz crystal to the top and bottom

electrodes, either a thin layer of Duco cement mixed with

silver paint and thinned down with a solvent, or Emerson

and Cummings Eccobond solder 56 C mixed with catalyst 9 was used. The other joints were made with thin layers of

Duco cement. After mounting the assembly into the outer copper tube, a suitable epoxy potting compound such as Eccobond 45 mixed with catalyst 15 was poured around the

assembly inside the copper tube. The assembly was then

allowed to cure overnight.

The stress gauge was calibrated against a dynamic

stress of known value. This was accomplished with the help

of a dynamic fatigue testing machine, where the assembly

was subjected to a known alternating compressive stress.

Fig. 10 shows the voltage-time curves for the calibration

load cell of the fatigue machine (upper trace) and that of

the piezoelectric stress gauge (lower trace) photographed

simultaneously during stressing, using a dual beam scope. The calibration was done both at room temperature and at

77 with the assembly immersed in liquid nitrogen. Fig. 11 shows the stress calculated from the calibration curve

for the load cell plotted against the peak voltage from the

piezoelectric stress gauge. The piezo-voltage varies

linearly with the applied stress and the slope of the straight line gives the stress required to produce an output of one volt from the piezoelectric stress gauge.

C. EQUIPMENT FOR TESTS AT LOW TEMPERATURES

For the tests at 77°K the assembly A to F fixed by rigid

rods to the top plate L in Fig. 1 was lowered into an open

dewar containing liquid nitrogen. The top part of the

491
piezoelectric stress gauge was above the liquid level. For

the tests at 4.2°K the assembly was immersed into a liquid

heliiam dewar placed inside an outer liquid nitrogen dewar.

Fig. 5 shows the photograph of the assembly with the dewars

in position. For the tests conducted above 4.2°K a heater

was made by winding 100 cm length of fine manganin resistance wire on a grooved copper tube of length 1.9 cm and

diameter 2.2 cm. This heater element was connected to an external variable D.C. power supply capable of supplying

0-25 volts at 0 - 800 milliamps. This heater was placed

around the sample. A thermocouple with one element copper

and the other element gold with 2 at.% cobalt was used to measure the temperature of the sample while heating. The

thermocouple voltage was measured using a Leeds and Northrup K-3 Universal potentiometer and a Leeds and Northrup

D C null detector. Ice with water was used as the reference

junction and the standard calibration chart prepared by

Richard Byrnes of m/S Andonian Associates was used after

suitable zero point correction, to read the temperature

from the millivoltage measured by the potentiometer. Since

the temperature fluctuated, the measured and the actual

temperatures could differ by as much as 1°. CONCLUSIONS

Impact stress pulses of the order of a few kg/mm^

and duration of a few microseconds could be applied to single crystals of metals by means of a simple equipment

built in the laboratory to study dislocation mobility in

metals The stresses were measured by using a piezoelectric

stress gauge using X~cut quartz as the transducer. The gauges were calibrated against a known alternating stress. Absolute

values of the stress measured by the gauges are conservatively

estimated to be within + 10%. It is noteworthy here that

recently, Brunton and Wilson (17) have used X-cut quartz to measure stresses in their study of the kinetics of twinning

in zinc and tin crystals. They calibrated their quartz

crystals against known impulsive loads produced by dropping

hard steel balls. They quote a value of 328 N/V to produce

one volt in the oscilloscope. This corresponds to about

0.27 kg/mm^ for a quartz crystal of diameter 1.25 cm,

as against our value of 1.96 kg/mm (from the slope of the

straight line in Fig. 11) to produce one volt in the

oscilloscope. The difference could be due to the facts that

(1)- we enclosed the quartz crystal with an epoxy potting

compound within an assembly, and (2)- we used a lead nose

piece over the quartz crystal. However, since the assembly

was calibrated against a known stress, this difference does not affect our main conclusions of the dislocation mobility

measurements (ref. 6,7).

492
ACKNOWLEDGEMENT

The authors would like to express their sincere thanks

to M/S Hunt Corporation, Carlisle, Pennsylvania, for

providing the quartz crystals used in the piezoelectric

stress gauges. This research was supported by the Advanced Research

Projects Agency of the Department of Defense through the

Northwestern Materials Research Center.

REFERENCES

1. J. D. Esh^lby : Proc. Roy. Soc. (London), B 69, 1013

(1956)

2. C. S. Smith t Trans. Met. Soc. AIME , 214, 574 (1958)

3. J. Weertman : RES POSE OF METALS TO HIGH VELOCITY

DEFORMATION , Edited by Shewmon and Zackay, Interscience

(1961)

4. J. Weertman : J Appl. Phys . , 37, 4925 (1966)

5. J. Weertman : J. Appl. Phys., 38, 5293 (1967)

6. V. R. Parameswaran and J. Weertman : Scripta Met.

3, 477 (1969)

7. V. R. Parameswaran and J. Weertman : (to be published)

8. R. A. Graham : J. Appl. Phys., 32, 555 (1961)

9. R. A. Graham : Rev. Sci. Instr., 32, 1308 (1961)

10. R. A. Graham : J. Appl. Phys , 33, 1755 (1962)

11. 0. E. Jones, F. W Neilson and W. B. Benedick : J. Appl. Phys., 33, 3224 (1962)

12. 0. E. Jones : Rev. Sci,. Instr., 38, 253 (1967)

13. R. A. Graham, F. W. Neilson and W. B. Benedick : J. Appl. Phys., 36, 1775 (1965)

14. J. F. O'Brien and R. J. Wasley : Rev. Sci. Instr.,

37, 531 (1966)

15. W. G. Cady : PIEZOELECTRICITY , McGraw-Hill, New York,

London, (1946)

16. W. P. Mason : PIEZOELECTRIC CRYSTALS AND THEIR APPLI - CATIONS TO ULTRASONICS , D. Van Nostrand, New York, 1950

17. J. H. Brunton and M. P. W. Wilson : Proc. Roy. Soc. (London), A 309, 345 (1969)

493
FIG. 1-- Schematic diagram of the experimental set up for the application of the stress pulse.

V 494
FIG. 2-- Photograph showing the piezoelectric

stress gauge B, the single crystal C and the lead damper bar D.

495
no V AC

110 V AC DC TO COIL OF TO SOLENOID VALVE ELECTROMAG^ET

FIG. 3-- Electrical circuit used to operate the

electromagnet and the solenoid valve.

496
FIG. 4-- General view of the equipment.

497
FIG, 5-- Photograph of the assembly with the

dewars in position at the bottom.

498
FROM PULSE

Ch A

'OUT to" EXT. TRIG. -VE SLOPE

Ch B

FIG. 6-- Triggering circuit.

499
FIG. 7
EPOXY POTTIN-G

X CUT QUARTZ

CRYSTAL

COAXIAL CABLES

TOOL STEEL DISC LEAD NOSE PIECE COPPER TUBE

CRYSTAL ELECTRODES

PIEZOELECTRIC STRESS GAUGE

FIG. 10-- Voltage-Time pulses from the calibrated

load cell (upper trace) and the piezoelectric stress gauge (lower trace).

501
VOLTS

Calibration curve for the piezoelectric

stress gauge

502
APPENDIX

The equipment descibed in the text was used in the

study of the mobility of high speed dislocations in metal

single crystals. The method was to apply impact stress

pulses of a few microseconds duration to the single crystal

and to measure the lengths of the slip bands produced on

the surface. A slip band is produced by the expansion of dislocation loops on a slip plane. The length measured on

a surface gives the distance through which the screw components of these loops threading the surface have travelled

along that slip plane for the duration of the stress pulse.

This length divided by the total duration of the stress

pulse gives the average velocity of the leading dislocation

in that band. Fig. A-1 shows typical slip bands produced

on the surface of a lead single crystal by the impact

stress pulse.

Dislocation velocities were measured in high purity

lead and lead-indium alloy single crystals at temperatures

varying between 4.2 K and 298°K. Experiments were done in

both the normal and the superconducting states of the metals.

The velocities measured were in the range: 10-^ - 105 cm/ sec. RESULTS

Fig. A-2 shows the dislocation velocity v plotted

against the mean stress t (defined by equation (1) in the

text), for pure lead at 298°, 77° and 4.2°K. The velocity

varies linearly with stress in the lower regions of stress. At higher stresses, the variation is less than linear and

the dislocation velocity approaches a limitting velocity

equal to the shear sound velocity in lead at each temperature.

In the linear region of the velocity-stress curve, the resistance to dislocation motion is proportional to the velocity

itself and the slope of this linear region gives a drag

coefficient or damping constant B defined as: B = Th/v (2) where t is the shear stress applied to maintain the velocity

V. Fig. A-3 shows the damping constant B plotted for different temperatures for pure lead. The circles indicate the values

of B calculated from velocity measurements at temperatures

above 7.2°K in the normal lead. The value of B at 4.2°K is that of the superconducting lead. In the case of the lead-indium alloys (lead with 0.1,

1, and 10 at7o indium), the dislocation velocity was found

to vary linearly with stress. The damping constants

503
calculated are given in Table I. The value of B increases

with increasing alloy content. For a particular stress,

the dislocation velocity in the alloys was always smaller

than the corresponding value in pure lead. (The complete

details of the results are given in a forthcoming paper to be published soon) . . DISCUSSION

In the region where the dislocation velocity varies

linearly with stress, at temperatures above about 60°K,

most of the drag acting on the moving dislocations is caused

by the effects of phonon viscosity and phonon scattering

(Nabarro(l), Mason and Rosenberg(2) , Leibfried (3) )

These drag forces increase linearly with increasing

temperature. At liquid helium temperatures, the phonon drag

reduces to zero. But there is a drag due to conduction

electrons which increases with decreasing temperature

below about 60°K in a normal metal (Mason(4)). In a superconductor, the conduction electrons are paired into loosely

coupled "Cooper pairs" and the electron drag is considerably

reduced. Our experimental results are consistent with these

theories of damping of dislocation motion by phonons and

electrons

REFERENCE

1. F. R. N. Nabarro : THEORY OF CRYSTAL DISLOCATIONS , Clarendon Press, Oxford, 1968.

2. W. P. Mason and A. Rosenberg : J. Appl. Phys.,

38, 1929 (1967)

3. G. Leibfried : Z. Physik, 127, 344 (1950)

4. W. P. Mason : Phys. Rev., 143, 229 (1966)

504
TABLE Values of damping constants in

lead and lead-indium alloys.

Damping constant B (lO"'^ dyne sec/cm^)

298°K 4.2°K

Pure lead

Pb - 0.1 at.7o In

Pb - 1 at.7o In

Pb - 10 at.7o In

3.43 1.52

3.85

4.375

5.43 37.6

505
FIG. A-1-- Typical slip bands observed on the surface of the crystals after impact.

506
FIG. A-2-- Dislocation velocity vs stress

for pure lead.

507
FIG. A-3-- Damping constant B vs temperature

for pure lead.

508
Paper No. 35

SOME EFFECTS OF EQUALIZATION ON ACCELERATION RESPONSE AND

FATIGUE LIFE EXPECTANCY

R. M. MAINS'

REFERENCE: Mains, R. M. , "Some Effects of Equalization on Acceleration Response and Fatigue Life Expectancy," ASTM/IEC/

AIAA Space Simulation Conference, IU-I6 September, 1970. ABSTRACT: A set of calculations is reported for various transmissibilities of a lO-mass system, intended to "be representative

of an exciter, fixture, test article arrangement. The difficulties of achieving equalization at some coordinate "when it has a transmissibility range of 10^ or more are noted. By comparison

of r.m.s. responses to a level power spectral density input, it

is shown that factors of 5 to 10 or more difference in acceleration response result from equalization at different coordinate

locations. These differences in r.m.s. response would produce

corresponding differences in predicted real-time behavior. The

differences in r.m.s. spring stress responses were smaller for

various equalization locations, but still large enough that predictions of fatigue life expectancy (for the system used) would

be much affected by the choice of equalization location.

KEY WORDS: Vibration, test, evaluation, fatigue life, acceleration response, random vibration, transmissivity equalization,

mechanical impedance, transfer impedance, lumped parameter

system.

Introduction

At a symposium on reliability at Corona, California in

195^ (1)*, three papers were presented which discussed random

vibration. Two of them (D.M. Ellett and R.M. Mains) showed the

results of flight measurements and demonstrated the random

nature of the vibration. The third paper (C.T. Morrow) showed

Professor of Civil Engineering, Washington University, St. Louis, Missouri 63130

*Numbers refer to the list of references at the end of the

paper.

509
how the treatment of random noise in communications problems

could be adapted for random vibration in mechanical systems.

Since then, testing and measuring random vibration have become

commonplace, and off-the-shelf equipment is available. Papers

theorizing about random vibration abound in the literature, and

there are several schemes for assessing damage under random

vibration (2,3,^,5). As long as the vibration excitation spectrum is broad-^banded and reasonably level, the responses of an elastic system to that excitation are best measured by the area under the curve of the square of the transmissibility vs. frequency (6). The smaller the area under that curve, the lesser

the r.m.s. response of the phenomenon represented by the curve will be. For some time, the author has been concerned over various

practices in vibration testing, and in particular over the ques- tions of whether, where, and how to "equalize" in a random

vibration test. A study of the transmissibility curves for a system, and in particular the areas under the squared curves, could be expected to shed some light on the equalization problem.

Differences in transmissibilities and in area under squared

curves for various equalization locations would mean differences

in real-time performance and life.

The System and its Analysis A set of systems was devised, as shown in Figure 1, consisting of ten masses and springs in series, and constrained to move

in one direction only. It was intended that these masses and

springs would represent an exciter, fixture, test article com- bination, so nine different combinations of masses were tried

with a given set of springs to see what the frequency patterns

would be. Table 1 shows the various mass combinations and the

natural frequencies for each. It is of interest to note how

little change in the set of frequencies accompanys a considerable change in the arrangement of masses. There were no out- standing reasons for selecting any one of these systems over the

others, so the author chose the first set. In a system such as this, there are two phenomena of interest: the motions of the masses, usually in terms of acceleration;

and the stresses or forces in the springs. The motions are associated with malfunction problems, short circuiting, gyro

precession, relay chatter, and the like. The stresses are associated with fatigue damage accumulation and life prediction,

as mentioned above.

In the random vibration test system, equalizing means the

adjustment of the amplifier gain so that a level spectrum of

input produces a level spectrum of response at the equalization

510
point. In the calculations, mass 1 was taken as the driving

point (exciter armature), and masses 5 and 6 were taken in turn

as the equalization points.

The general scheme of the analysis was to generate the com- plex impedance, invert this to complex mobility, and work with

the mobility to get accelerations and stresses. First the natural frequencies of Table 2 were reduced to get a non-singular

impedance matrix:

03 2 = ^ 2 * (1 _ ^2) (1) r o in which o)^^ = eigenvalue from undam.ped system

C = damping coefficient, taken as 0.03

a)_^2 = reduced frequency, squared

Calculations with only these frequencies would not show enough

shape to the curves, so three more oo^ were calculated to divide

the intervals between each oa^^ uniformly. The frequencies used

are listed in Table 2. The stiffness, K, and mass, M, matrices were constructed in

the usual way, and the damping matrix, C, was calculated as:

I I 2 0=2? [M^rf ] (2)

where the square roots are term-by-term. Next the complex impedance, Z(ia3), was defined as: Z(i.) = Z^^ + iZ^^ (3)

Z^^ = K - oj^M (U)

^IM = ^5) Now Z(ia))*X(iw) = f(ico) (6) in which X(ia)) = X^^ + iX^,, complex response (7)

tihj IM' f(iw) = f^^ + if complex excitation (8) From eq. (6) X(ico) = z"^ ( ioo )*f ( ioj ) (9) If f(iw) is taken as real only, then

-1, E '^RE "RE (10)

511
in "which [(K - A) + ajC(K ~ ^Coj] -1 (12)

(13)

Equations (lO) and (ll) are the responses in inches to one- pound driving forces, and from them the quantities of interest

are calculated.

Results of the Calculations

To examine the feasibility of equalization, Tahle 3 shows

the driving force on mass 1 required to produce 1 g at each mass and each frequency. For any given limit of exciter force, a row

of the matrix shows whether the motion could be equalized within that force limit or not. Mass 5 is the best one for equalization over the full range of frequencies, while mass 6 would

have attenuation above frequency 32, Masses past 6 would have more attenuation in the higher frequencies if equalization were attempted with them. To show the range of driving forces involved. Figure 2 is a semi-log plot of the driving force at mass 1 for 1 g at mass 9- The range from 2.7 lbs at frequency 10, to 656,000 lbs at frequency ho is of particular interest. Figure 3 is a semi-log

plot of the driving force at mass 1 to produce 1 lb of force in

spring 8-9. The range from 0.53 lbs at frequency 10, to 307,000

lbs at frequency hO is of interest again.

Figure H is a semi-log plot of the acceleration of mass 9 for 1 g at masses 1, 5, 6. The differences between these curves

are in the magnitude of the peaks in the lower frequency range. For 1 g at mass 1, mass 9 moved 6.5 at frequency 3 and 6.8 at frequency 10. For 1 g at mass 5, mass 9 moved 86.0 g at frequency 6 and 5.^ g at frequency 10. For 1 g at mass 6, mass 9 moved 15 g at frequency 9 and 12 g at frequency 17. These are significant differences, in that the prediction of real-time

behavior or life from test data could differ by a factor of 5 or more depending upon which equalization point is used.

Figure 5 is a semi-log plot of the force in spring 8-9 for

1 g at masses 1, 5, 6. These curves show little difference in

general level and are similar in many details. The resonance

peaks in the low frequency range shift location with the change

in equalization point.

The driving point impedance, pounds of force at mass 1 to ^ produce 1 g at mass 1, is plotted in Figure 6. This is shown i primarily for the shape and character of the curve. 6 Table k lists the r.m.s. pounds at mass 1 for a level power ^ 512
spectral density of 1 g^/cps at the various masses. These

values are prohibitively large, so that a p.s.d. level of

1 g2/cps could not he physically achieved with this system.

This is because the system between mass 1 and mass 5 acts as a vibration absorber in the frequency range above number 20 or 25,

and it is physically unreasonable to try to force a level spectral density of response. If a level response could be forced

in the test, the results would be grossly misleading for conversion to real time.

Table 5 shows the r.m.s. pounds in the springs for a level

power spectral density of 1 g2/cps at masses 1, 5 and 6. These

values correlate with the similarity of the curves in Figure 5,

in that the peak at frequency 6 for mass 5 produces a larger

r.m.s. response by a factor of h or more. Predictions of realtime behavior or life of the springs would be significantly

different according to which equalization point were used.

Table 6 shows the r.m.s. g's at the various masses for a level power spectral density of 1 g2/cps at masses 1, 5 and 6. Here is where the differences in equalization points show clearly. Equalization at mass 5 is about 6 times as severe as equalization at mass ]

, while equalization at mass 6 is about

li times as severe as at mass 1 and only 1/h as severe as at mas s 5

Conclusions

!tf data from a random vibration test are to be used to predict real-time behavior or life of an equipment, then care must

be taken that the test fixture and the equalization process

used do not lead to erroneous predictions. For the system used

in this study, equalizing for a level response on any other mass

than no. 5 could not be done. Equalizing at masses past 5 would

require that attenuation in the higher frequencies be accepted.

One alternative would be p redesign of the fixture for better

transmissibility in the upper frequency range. Another alternative would be to reconsider the desirability of trying to use equalization at some point as a device for insuring correlation

between test time and real time. For the system calculated, factors of 5 to 10 or more difference in r.m.s. acceleration responses to a level power spectral density input, were noted for equalization at different

locations. These differences in response would produce corresponding differences in prediction of real-time behavior. The

r.m.s. spring stress responses showed factors up to U or 5 difference for the various equalization locations. These

differences would produce appreciable differences in fatigue

life estimates.

513
List of References

"Proceedings of the Symposium on Guided Missile Reliability','

U. S. Naval Ordnance Laboratory, Corona,Calif., October 195^.

Clevenson, Sherman A. and Steiner, Roy, "Fatigue Life Under

Various Random Loading Spectra," Shook and Vibration Bulletin No. 35^ IIJ January I966.

Root, L. W. 5 "Random-Sine Fatigue Data Correlation," Shook

and Vibration Bulletin No. 32^ II, February 196U.

Bieber, R. E. and Fairman, J. H. , "Random Fatigue Data,"

Shook and Vibration Bulletin No. 31, III, April 1963.

Fuller, J. R. , "Cumulative Fatigue Damage Due to VariableCycle Loading," Shook and Vibration Bulletin No. 29, IV, June 1961.

Mains, R. M. , Chapter 12 in Random Vibration, edited by

S. H. Crandall, John Wiley and Sons, 1958, p. 12-32.

514
2: ^

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515
TABLE 2. FREQUENCIES

USED FOR COMPUTATION

FREQ.

NO. CPS

FREQ.

NO. CPS

1 0.4035 21 184.81

2 0.4659 22 198.21

3 22.955 23 200.32

4 32.459 24 202.41

5 39.753 25 204.48

6 45.902 26 206.53

7 55.860 27 217.49

8 64.293 28 227. 92

9 71 .742 29 237.90

10 78.488 30 247. 47

1 1 91 .508 31 262.03

12 102.89 32 275.81

13 113.14 33 288.95

14 122.53 34 301 .51 15 126.30 35 305.68

16 129.95 36 309.80

17 133.51 37 313.86

18 136.97 38 317.88

19 154.57 39 321 .84 20 170.36 40 325.76

516
TABLE 3. POUNDS AT MASS 1 FOR IG AT MASS N

I 2 3 4 5

1 0.27000D 0^ 0.233480 04 0. 182250 04 0. 1 76040 03 0.484950 02

z 0.26998D 04 0.23346D 04 0.456620 04 0.232900 03 0.57756D 02

3 0.26996D 04 0.233440 04 0.116230 04 0.481300 03 0.396720 02 0.26993D 04 0.233410 04 0.572090 03 0.160570 04 0.299650 03

3 0.26991D 04 0.233390 04 0.419540 03 0.387970 03 0.362780 03

b 0.26990D 04 0.233370 04 0.328380 03 0.2128BD 03 0. 105020 03

1 0.269890 04 0.233360 04 0.296930 03 0.171530 03 0.739390 02

a 0.269BBO 04 0.233350 04 0.28490D 03 0.157560 03 0.648650 02

3 0. 269880 04 0.23335D 04 0.279930 03 0.152060 03 0.614590 02

13 0.269B8D 04 0.23335D 04 0.277830 03 0. 149780 03 3.600780 02

1 1 12 13 14 15

1 0. 10258D 03 0.223830 03 0. 161690 03 0. 10453D 02 3.204860 02

2 0.71176D 02 0.780480 02 0.904640 02 0-420330 02 0.426030 02

3 0. 71418D 02 0.567960 02 0.358520 02 0.961940 01 0.293810 02

4 0.102490 03 0.604930 02 0.286940 02 0.521150 01 0. 124950 02

5 0.293730 03 0.11347D 03 0.445570 02 0.693320 01 3.155430 02

6 0.170480 03 0.170470 03 0.105780 03 0.23412D 02 D.61472D 02

7 0.289420 03 0.168080 04 0.377320 03 0.265560 02 0.437710 02

3 0.841600 02 0.110310 03 0.715590 02 0.130550 02 0.2B0240 02

9 0.602390 02 0.703550 02 0.399570 02 0.623130 01 0. 124010 02

10 0.53067D 02 0-597670 02 0.326880 02 0.490200 01 0.959070 01

21 22 23 24 25

1 0.701620 02 0.138260 02 0.137160 02 0. 14891D 02 0.167090 02

2 0.232540 02 0.12850D 02 0.146440 02 0.182440 02 0.235040 02

3 0.218120 02 0.777920 01 0.812910 01 0.922710 01 0.107540 02

4 0.122960 03 0.588180 02 0.525270 02 0.51065D 02 0.51434D 02

3 0.272480 02 0. 107150 02 0.112530 02 0.128000 02 0.14908D 02

6 0.226440 03 0.10581D 03 0.114020 03 0. 132940 03 3.158500 03

7 0.71151U 03 0.49486D 03 0.549680 03 0.618860 03 3.661380 03

8 0.977580 03 0.20228D 03 0.187700 03 0.190760 03 0.206020 03

9 0.332330 03 0.125790 03 0.126290 03 0.138380 03 0.160640 03

13 0.17140D 03 0.55783D 02 0.545110 02 0.580910 02 3.655360 02

31 32 33 34 35

1 0.293320 02 0. 346230 02 0.385850 02 0.37276D 02 3.36194D 02

i 0.385900 02 0.733290 02 0.736500 02 0.450230 02 3.435290 02

3 0.470860 02 0.405270 02 0.312810 02 0.278190 02 0.306090 02

4 0.320060 02 0. 457970 02 0.497250 02 0.566510 02 3.669180 02

5 0.54115D 02 0.952460 02 0.122710 03 0.161720 03 3,199870 03

6 0.99534U 03 0.195540 04 0.278010 04 0.40067D 04 0.509610 04

7 0.66139D 04 0.148400 05 0.236760 05 0.377910 05 0.496500 05

8 0.152680 05 0.423410 05 0.814730 05 0.154550 06 0.212580 06

9 0.11453D 06 0.182480 06 0.149670 06 0.13839D 06 0.152610 06

13 0.73383D 04 0.184870 05 0.29334D 05 0.41171D 05 0.507600 05

517
TABLE 3. CONTINUED

6 7 8 9 10

1 0.34019D 01 0.43420D 02 0.70753D 02 0.11308D 03 0. 182630 02

2 0.38225D 01 0.445040 02 0.653090 02 0.852/30 02 0. 266080 02

3 0.53647D 01 0.55617D 02 0.731610 02 0.788260 02 0. B1758D 02

<t 0. 118200 02 0.9 31770 02 0.102780 03 0.871640 02 0. 209380 02 0.42101D 03 0.251890 03 0. 193000 03 0.121610 03 0. 145560 02

6 0. 10450U 02 0.216170 03 0.713410 03 0.530170 03 0. 147000 02

7 0.625560 01 0.8268'»0 02 0.101950 03 0.714530 02 0. 687450 01

b 0.522670 01 0.623290 02 0.683200 02 0.417370 02 0. 340640 01

9 0.486030 01 0.558370 02 0.588660 02 0«34523D 02 0. 269940 01

10 0.471460 01 0.533590 02 0.554050 02 0.319960 02 0. 246290 01

16 17 18 19 20

1 0. 329670 02 0.438690 02 0.490420 02 0.861110 02 0.224100 03

2 0. 486890 02 0.511010 02 0.517840 02 0.465880 02 0. 363740 02

3 u. 777850 02 0. 166070 03 0. 148660 03 0.111460 03 D. 502730 02

4 0. 234560 02 0.351350 02 0.395910 02 0.730450 02 0.100340 03

5 0.270900 02 0.374890 02 0.424860 02 0.574030 02 0.<>85670 02

6 0. 129950 03 0.206630 03 0. 155740 03 0.290260 03 0.326850 03

7 0. 527910 02 0.465950 02 0.428130 02 0.356630 03 0.696430 03

b 0. 425680 02 0.463720 02 0.520010 02 0. 14811D 04 0.414310 04

9 0. 173400 02 0.172430 02 0.174710 02 0. 19637D 03 0.399510 03

10 0. 131800 02 0.128770 02 0.128150 02 0.12977D 03 0.23500D 03

26 27 28 29 30

1 0. 187780 02 0.301660 02 0.412180 02 0.472930 02 3. 30411D 02

2 0. 303830 02 0.920560 02 0.601330 02 0. 310660 02 0.213490 02

3 0. 124790 02 0.219080 02 0.288370 02 0.329860 02 0.391020 02

<t 0.521800 02 0.498090 02 0.389700 02 0.261390 02 0. 195040 02

5 0. 172400 02 0.282490 02 0.316010 02 0.275970 02 3.253680 02

6 0. 187330 03 0.343870 03 0.427730 03 0.411020 03 0.412120 03

7 0.687210 03 0. 123890 04 0. 185920 04 0.206710 04 0.233650 04

8 0.236830 03 0.868700 03 0.200250 04 0.296080 04 3. 1^14460 04

9 0. 197980 03 0.103330 04 0.331090 04 0.68846D 04 0.142910 05

10 0.784320 02 0.342170 03 0.882650 03 0. 139720 04 0.202030 04

36 37 38 39 40

1 0.357570 02 0.358570 02 0.362420 02 0.367410 02 3. 372690 02

2 0. 4517^0 02 0.4b6820 02 0.531690 02 0.581460 02 0.633550 32

3 0. 357200 02 0.427860 02 0.514180 02 0.613410 02 0.723750 02

4 0.833960 02 0. 106190 03 0. 135130 03 0. 170130 03 0.211200 03

5 0.260090 03 0.345190 03 0.457080 03 0. 597890 03 0. 770080 03

6 0.681910 04 0.929900 04 0.126420 05 0. 169660 05 3.224070 05

7 0.685380 05 0.962810 05 0.134640 06 0. 185680 06 0.251810 06

8 0.300350 06 0.417320 06 0.563490 06 0.755710 06 0^102190 07

9 0. 180090 06 0.225730 06 0.302880 06 0.435620 06 0.655930 06

13 0.662510 05 0.910180 05 0.132850 06 0.20653D 06 3.33430D 06

518
TABLE 4. RMS POUNDS

AT MASS 1 for 1g2/cps

SPECTRAL DENSITY

AT MASS

1 RMS 1 BS. 1 0842+5

5 .87397+4

6 .67102+5

78 .71624f6

.29377+7

9 .20229+7

1 0 .84270+6

TABLE 5. RMS SPRING FORCES

RMS LBS . IN SPRING N FOR RAT 1 OS

SPRI NG

2

1 a /cDS@

2

In1 y // CD^ L-/ ^(3 o ^5 2

1g /cps@6

2

Iq /cps@5

2 In /p n c;(3fS 1 / [J O O w NO. @ 1 @ 1 5-6 6753 37300 7869 5.525 1 .165

6-7 3012 14541 3440 4.825 1 .442

7-8 1 102 4820 1416 4.375 1 .285

8-9 347 1297 469 3.785 1 .352

9-10 67 264 99 3.957 1 .492

TABLE 6 RMS G's AT MASSES

RMS G's AT MASS N FOR RMS G s / .18077+2

MASS

2

Ig /cps@l

2

g /cps@5

2

g /cps@6 lg^/cps@l

2

Ig /cps@5

2

Ig /cps

1 .18077+2 . 1 1440+3 45804+2 1 .000 6.330 2.533

5 .33482+2 .18077+2 39869+2 1 .853 1 .000 2.205

6 .24246+2 . 1 1527+3 18077+2 1 .342 6.380 1 .000

7 .27356+2 .19218+3 .35147+2 1 .513 10.63 1 .945

8 .341 57+2 .23064+3 .52449+2 1 .890 12.77 2.903

9 .41427+2 .24870+3 .73529+2 2.293 13.75 4.070

10 .45800+2 .25680+3 .86931+2 2.534 14.20 4.810

519
0. 100x10 LB/IN

0. 100x10

q:

UJO

X

tjj 0,700x10

0.600x10

0.800x10

0.690x10

0.400x10

-IO

0.268x10

0.155x10

0.720x10

FIG.I. SCHEMATIC DIAGRAM OF SYSTEM

520
521
522
523
524
Paper No. 36

CHARGED PARTICLE TRACK IMPLICATIONS FOR ACCELERATED TESTING

1 2

T. M. Flanagan and V. A. J. van Lint

REFERENCE: Flanagan^, T. M., and van Lint, V. A. J., "Charged

Particle Track Implications for Accelerated Testing/' ASTM/

IES/aIAA Space Simulation Conference,, lk-l6 September 1970.

ABSTRACT: This work considers the ionization effects produced

hy the passage of energetic particles through material, emphasizing the fact that the ionization is formed along particle

tracks rather than being uniformly distributed throughout the

material. Thus, the particle track through a material can be

visualized as a cylinder centered on the particle path disturbed by ionization and carrier trapping. The dimension of

the disturbed region is related to the range of the electron

ejected by the ionization event, while the density of carriers

within the disturbance depends on the ejected electron range

and on the ionization efficiency of the energetic particle.

The disturbed regions then relax with characteristic times

ranging from the very short to the very long. The dose -rate dependence arises from the relationship of the characteristic

relaxation times and the average time for two particles to disturb the same volume. At low doses, each particle track traverses primarily unirradiated material, and the effects are additive. At higher doses, nonlinearities are expected, since

particles penetrating a previously disturbed region may find

the carrier traps in a different charge state than in virgin

material. Another source of nonlinearities is the density of ionization which will vary with particle type and energy. The

size of the disturbance around a track varies from 100 to about

2000 depending on the material. This quantity can be used

to calculate the average overlap time for a given dose rate.

The dose which must accumulate before nonlinearities are apparent can be estimated from the trap density of the material.

Each of these factors must be taken into account when planning

on assessing simulation of radiation effects. KEY WORDS: radiation testing, radiation damage

Ionization Effects Group Leader, Defense Sciences Division,

Gulf Energy and Environmental Systems, Inc., San Diego, Calif.

2Vice President, Gulf Energy and Environmental Systems, Inc., San Diego, Calif.

525
In radiation effects testing,, it is usually impossible and

often not even desirable to reproduce exactly the environment

to which a component or system will be subjected. Especially

for space applications^ where real dose rates are very low and

exposures to the space environment are long^ some form of ac- celerated testing for radiation effects is dictated by economic

necessity. The problem of assessing and perhaps designing a simulator for testing the system or component must then be

faced. This paper develops some simulation requirements by

considering the interaction of energetic radiation with materials from a microscopic point of view.

We consider the effects resulting from a single event

the passage of a single charged particle through a material. (l)

The primary event produces a disturbance (primarily electronhole pair production) at particular sites in the material; a line connecting these sites is the particle track. The density

of disturbances along a track depends on the particle type and

its energy; generally^ increasing density results as the energy

is decreased or the particles become more massive. A transport

of the disturbance outward from the center of the track follows

the passage of the particle and leaves behind the effect of the

disturbance^ i.e.^ trapped charge^ along the track. Thus, at

some time after the passage of the particle, a radius of disturbance is established around the track. Some relaxation processes associated with charge trapping, untrapping, diffusion,

and recombination then take place as the disturbed region tends

toward a new equilibrium.

DOSE -RATE DEPENDENCE

The origin of the dose -rate dependence can be understoodby

considering the relationship between the characteristic relaxation times of the disturbance and the average time for particles

to disturb the same volume of the material (called the track

overlap time). In general, the disturbance caused by the passage of a charged particle decays with a variety of relaxation

times, ranging from the very short to the very long. If the

track overlap time in the real environment is comparable to a relaxation time in the material, particles will be traversing

material which is still disturbed by a previous particle passage, and the disturbance created will be modified by the existence of the previous track. Thus, lacking any knowledge of the physical processes occurring in the material, the track

overlap time of the real environment must be reproduced by the

simulator. This requirement can be relaxed somewhat if the re- laxation times in the material are known. Then if the track

overlap time for the real environment should fall between two relaxation times in the material, the track overlap time of the

Numbers in parentheses refer to the list of references ap- pended to this paper.

526
simulated environment can fall anywhere in the time window

"bounded by these two relaxation times.

If enough is known ahout the relationship of the disturbance to the relevant effect_, then some relaxation processes may

not be important and the size of the time window for track over- lap can be enlarged. For example^ if transient optical absorption measurements following pulsed irradiation of a spacecraft

thermal control material should reveal that the color centers

are formed in a few micros econds_, any longer-term relaxation

processes could be neglected.

The relaxation processes often take the form of "annealing"

or "self healing/' which simply means that the radiation effect

tends to disappear with time. The rate at which annealing takes place often depends on the average level of ionization in the

material and^ thus_, on the dose rate. The effects of the total

environment—i.e.^ particle irradiation, photon irradiation,

temperature, and ambient gas—^must be considered in evaluating

the effect of annealing, since these are all important to an- nealing rates in various systems. This implies that it is just

as important to study the annealing rates as it is to study the

damage rates in any simulation test.

DOSE DEPENDENCE

Regardless of dose rate, for low doses a negligible fraction of the tracks overlap. In this case, the fundamental physical effects must be linear in dose—i.e^, superposition must

hold. This follows from the fact that, if the tracks do not

overlap, they are independent of one another. For conductivity,

the fundam.ental physical effect is the generation of free carriers, and the effects of carriers contributed by each track

are additive if the tracks do not influence each other. This

does not mean that all physical quantities which can be used to represent the effect are linear—e.g., resistivity is not—but

only that the physical effect is linear in dose. As the dose

accumulates in the material, a larger fraction of the volume

has been previously disturbed by particle tracks, and no general statement can be made about the dose dependence. The relevant quantity at higher doses is the dose delivered in a time

comparable to the relevant relaxation time; this dose is coupled to the dose-rate dependence.

SPECIFIC IONIZATION DEPENDENCE

The specific ionization of an irradiating particle is re- lated to the density of ionization events along a track. Usually, it is defined as a dimensionless number (s) numerically

equal to 1 for a minimum ionizing particle—e.g., a relativistic electron. Thus,

527
S = number of ionization events per cm of path for particle ^ number of ionization events per cm of path for relativistic

electron.

We can make no a priori assertions about the linearity of an effect as the value of S varies, since this quantity determines

the distribution of charge within the disturbed volume. We can make a few general assertions:

1. The overlap dose must increase at least linearly with

increase in the specific ionization, since the radius

of the disturbance does not increase but the density

of tracks is inversely proportional to specific ionization at a given dose.

2. The effect of an external electric field applied at an angle to the track will decrease with increased S due

to the increased screening within the track. (1)

SIZE OF THE TRACK

The radius of the disturbance around a particle track is

generally in the range 100 to 2000 depending on the material

and the particle type. For a few special cases, such as very

pure crystalline anthracene, the disturbance may be of larger

dimension, but these special cases arise because of the migration of a long-lived exciton state (electron and hole move to- gether). (2) In most materials, the outward propagation of the

disturbance is due to the secondary electrons ejected by the

primary event. Then the size of the track can be deduced from

the range-energy relation for very low-energy electrons . (l)

Unfortunately, the exact range -energy relation for these very

low-energy electrons in solids is unknown, and we must rely on experimental inference to deduce the size of the track.

For the aromatic hydrocarbons, the radius of the track is estimated to lie between 100 and 300 1.(3^^) In water, (5) the

ejected electron travels about 100 A before being captured by a hydrogen ion or a water molecule, while in liquid hydrocarbons,

the range is nearer 300 A»(6) For insulating solids which have

some ionic bonding character, the range is longer (near 1000 A

in good single crystals ), (7) while in semiconductors a range of

2000 A seems reasonable .( 2) Thus, once the class of material

is known, range can be estimated.

NUMERICAL EXAMPLES

Scanning Simulators : Applying these concepts to a real simulation problem, let us discuss the simulation of proton (solar

wind) damage to optical and thermal control materials. From

the discussion, it can be asserted that the electrons ejected

from a particular proton track will not influence more than a cylinder of radius of ~103 A.

528
It is well known that the carrier trapping process (color

center formation) can be a strong function of the carrier generation rate. The underlying cause of this dependence is that

the simultaneous presence of a higher concentration of free

carriers (electrons, holes) influences the dynamics of trapping

and recombination by changing the average charge states of impurity centers that act as traps and recombination sites. The

importance of this potential nonlinearity is that it depends on the microscopic carrier density, not on the spatially averaged

value if the two are different. If one simultaneously exposes

a sample to N proton tracks that do not interact (average spacing greater than 2000 1)^ "the processes occurring within each

track are unaware of, and unaffected by, the presence of the

other tracks. They are strongly influenced by the ionization

density within the track, but this is a property determined by

the proton's energy, not by how many tracks there are. There are important nonlinearities once the tracks overlap. A proton penetrating a previously ionized region may find

the longer-lived traps in a different charge state than in virgin material. Furthermore, the charge state depends on a com- parison of average time between overlaps and the trap relaxation time. In very impure or heavily displaced materials, there

is almost a continuous spectrum of trap relaxation times; hence,

there are likely to be some trap relaxation times in any range

of interest.

This discussion suggests that important simulation require- ments are to reproduce the proton specific ionization (same

proton energy) and the average time between production of overlapping tracks. This latter requirement is particularly relevant to the evaluation of non-CW simulators (including RF mass

separators or beam-scanning techniques). At the maximum ex- pected solar flux (0 = 2 x 10^ protons/ cm^-s ), the mean overlap time, tQ, for a track of Rq = 1000 A radius is

tQ = = -0.02 s (1)

It is important to reproduce this overlap time, or any longer

overlap time, at lower fluxes or electron ranges. This implies

that for a scanning system the peak flux must not create overlap during a single pass, and that the pass repetition rate

must be fast enough to create overlap between passes within

0.02 s. Both these criteria are met if the pass repetition

rate is faster than 100 Hz. They are obviously met for an RF

mass filter that operates in the 10^-Hz regime.

Critical Dose ; The dose at which significant track overlap

occurs is called the critical dose. The exact choice of the

critical dose is somewhat arbitrary, since tracks are produced

randomly throughout the material.. Harrity(7) has discussed

this problem, and we will follow his analysis.

529
The rate at which the volume that has been irradiated at least once increases with fluence as dV^ V - I s I

u V V

3o

(2)

(3) or

= (1 - e-"^/^) , (4)

where u = vol-ume ionized by a single incident particle^

Vj = total irradiated volume^

Vg = total sample volume^

$ = irradiating electron fluence_,

d = sample thickness.

If we consider u as a cylinder with length d and radius R and realize that both Vj and Vg can be written as d times the

equivalent areas^ Eq. 4 may be rewritten as

= (1 - e-"K^*)

, (5)

where Aj is the irradiated area of the total sample area Ag

In recent work at this laboratory Harrity(7) found that

in Cz-grown single -crystal sapphire_, the conductivity is linear

in the dose for doses below 10 rads and above 500 rads . Be- tween these doses is a nonlinear transition region which is interpreted as the result of particles traversing regions which

have been previously irradiated. He uses this data to derive a value for Rq 10^0 A

We may use this data to find the value of Aj/Aq for which

significant overlap occurs. Deviations from linear behavior

were observed at 10 rads (4 x 10 ^> 30 MeV e/cm^), and the value

of Aj/Aq for this point is 0.126 from Eq. 5« Thus_, significant

overlap occurs when about l/8 of the total area has been influenced by a radiation track.

From a practical viewpoint^ the track overlap is sufficient to produce nonlinearities only if enough carriers are generated within the track to fill a significant fraction of the traps. A minimum condition for a significant perturbation of a track would be that enough carriers have been generated in the

vicinity of the track to fill 50 percent of the traps ^ assuming

all the carriers to be captured. That this is a reasonable

criterion can be seen by applying it to Harrity's data. Har- rity(7) estimates the trap density in his Cz-grown_, singlecrystal sapphire to be 6 x

10-^^ cm~3. At a generation rate of

3 X 10^3 carriers/cm3-s, (8) 10 rads are required to fulfill

530
this condition and Harrity's data show deviations from linearity just above 10 rads. For most insulators in technological

use. a more reasonable trap density is of the order of 10l3 to

10^^I cm~^ , and a dose greater than 105 rads is required to achieve critical dose.

Vacancy Motion in Silicon ; So far we have considered only effects which are the results of ionization. However accelerated production of displacement events can also be evaluated.

Consider the displacement of atoms in silicon. At recoil

energy transfers just greater than the threshold for displacement_, it is likely that the displaced atom is still within the

strain field of the vacancy and it is most probable that it

will return to its lattice position. A very small fraction of

these displaced atoms may be favorably directed away from the

track and escape as interstitials moving through the lattice.

The vacancy moves also, and each will continue to move until it

undergoes some type of immobilizing reaction at another defect.

Since there is very little information on the interstitial, the

remainder of this discussion will concentrate on the vacancy.

The questions are, what are the mechanisms by which the

final defects are formed from diffusion of vacancies from the

defect track, and how large can we allow the vacancy concentration to become—i.e., how much can we accelerate the test without perturbing the results?

As a vacancy diffuses away from its formation site, it has

a choice of combining with various kinds of defects to form a stable complex. In a typical quartz crucible (QC) grown silicon crystal, these include:

1. Mobile radiation-produced interstitial atoms and

vacancies (density 10^2 to lO-'-^ cm'^),

2. Interstitial oxygen atoms (density lO^T to lO-'-^ cm'^)^

3. Impurities such as donor or acceptors (density 10^^ to lOlS cm-3)^

Dislocations (density 10^ cm-2);if we assume a strain

field affecting a cross section of five atoms, the dislocations yield an effective density of approximately

10l2 cm" 3^

5. Surfaces.

It is clear from the foregoing that dislocations can normally be neglected. For equal capture cross sections, the interstitial oxygen atoms will dominate vacancy capture to form

A-centers. Since the interstitial oxygen atoms are neutral in

both n- and p-type silicon and the vacancy can be charged, it

is possible that a charged impurity atom may compete with the

oxygen atoms for the vacancies in heavily doped material, or in

float-zone (FZ) material in which the oxygen concentration is

much less. The effect of the surface can only be estimated by

evaluating the mean free path for capture of a vacancy within

the bulk of the crystal. It can be assumed that a neutral

531
capture cross section is 10"!^ cm^ and an attractive charged

1 it P cross section is 10"-^^ cm^. The average net distance, d, trav- eled by a vacancy prior to its being captured by a defect -can be estimated from the mean free path for capture, X^, and the mean free path for scattering, \q. Let us assume that the mean

free path for scattering, Xg, is 5 X 10"° cm or approximately

equal to the interatomic spacing. Considering the random-walk

nature of the process, and assuming the scattering to be random

and isotropic at each site.

Now

^c " Na

where N is the density of capture centert and q- is the capture

cross section. For A-center formation in QC silicon, let us take an oxygen atom density of 3 X cm"-^. Then 3 X 10"^ cm and d «i 4 x 10~5 cm. For formation of E-centers

(vacancy-phosphorus complexes) in n-type FZ silicon with an

ionized donor concentration of cm"3, Xq 10""! and d s« 7 X 10"5 cm. In either case, d is small compared to the size

of typical samples, and s-arfaces are not likely to be sinks for

mobile defects.

Addressing the second question above, the length of time

elapsing before the stable defect is formed is

where u is the Jump frequency. For a defect which is not found

at each lattice site, u is approximately equal to the lattice

frequency and t 10~ s. In the actual case, the vacancy mobility is dependent on the charge state of the vacancy, and is an activated motion describable by a relation of the type

-E/kT

u = u e ' , o

' where Uq E are functions of the charge in the vacancy. (9)

Assuming first-order formation kinetics, for which there is some evidence, (lO) the equilibrium vacancy concentration is

[V] ^ [V] t , where [V] is the production rate for free vacancies and t is the time before capture

Now, let us apply these calculations to a 1-MeV electron

irradiation, and compute the equilibrium vacancy concentration

in a steady 10-mA irradiation .of a l-cm^ solar cell. This cur- rent is equivalent to 6 x

10-^^ e/cm s which, at an introduction rate of 0.2 cm"-'- of free vacancies, yields [V] 1.2 x 10^^ (cm3 s

)"-'-. The longest-lived vacancy occurs in p-type

532
silicon; therefore^ this will be considered as a worst case.

The time before capture for a vacancy in p-type silicon at 300 K is approximately 2 x 10"3 s.(9) This implies an equilibrium

vacancy density of approximately 2.4 x 10^3 cm"3_, which is much

smaller than either the oxygen concentration in QC material or most impurity concentrations of interest in FZ material. Thus^

the irradiation rate used does not create enough vacancies to perturb the impurity vacancy interactions. The effect of the

increased ionization caused by the increased dose rate in the test compared to the actual environment may change the charge

state of the vacancy. However,, this shortens the time during

which the vacancy is free by increasing its mobility. Therefore,

the enhanced dose rate tends to lower the steady-state vacancy

concentration, which decreases further any perturbation of the

vacancy- impurity interaction. Thus, the accelerated dose rate used in this example can be used to simulate displacement damage in silicon produced by all lower dose rates.

CONCLUSIONS

In assessing a radiation effects simulation experiment,

seme knowledge of the mechanisms of the relevant effects is

necessary. The more that is known about the mechanism of the

effects, the more confidence there is in the results. For that

important class of radiation effects which depend on the disturbance created by the transport and trapping of charge around

a particle track (ionization),

1. The average track overlap time in the simulated environment must appropriately represent that in the real

environment;

2. Extrapolation to higher dose levels must take into

account the potential nonlinearities which can occur when the critical dose is exceeded;

3. The influence of "self healing" or annealing must be

taken into account.

In addition, consideration of the mechanisms of displacement defect formation, coupled with the ionization track model,

can be used to evaluate the effects of accelerated testing on

the damage formed.

533
\ REFERENCES

1. Kallman, H., "Mechanisms of Energy Transfer in Dielectrics^"

Technical Report EC0M-00129-F;, January 196^0

2. Nichols, D. K., and Vc A. J„ van Lint, IEEE Trans on Nucl.

Sci. NS-13, p. 119;, December 196b.

3. Watanabe, Y., N. Saito, and Y. Inuishi, "Carrier Production and Transport in Anthracene," Osaka University Technology Reports , Vol. I8, Nos . 801-322, p. 143, March 1968.

4. Kepler, Ro G., and F. N. Coppage, Phys. Rev. 151^ 61O

(1966).

5. Kupperman, A., in The Chemical and Biological Action of Radiation, Vol. V, M. Hassinsky (Ed. ), p. 93^ Academic

Press, 1961.

6. Hummel, A., and A. 0. Allen in Radiation Chemistry, Dobo

and Hedvig (Eds.), p. 295^ Akademiai Kiado, Budapest, I967.

7. Harrity, J. W. (to be published in IEEE Trans, on Nucl.

Sci. , December 1970).

8. TREE Handbook, DASA 1420, September I969.

9. Cesena, R., T. M. Flanagan, R. E. Leadon, Co E. Mallon, J. A. Naber and E. Wenaas, "Transient Radiation Effects," Summary Report on contract DAAG39-68-C-0060, Gulf General

Atomic document GA-IOO98, p. 33 , April 1970.

10. Harrity, J. W., and C. E. Mallon (to be published in IEEE

Trans, on Nucl. Sci. , December 1970).

534
Paner No. 37

A SIMULATION TO VALIDATE FUTURE SPACE MISSIONS

T. P. Foley} A. W. Metz^and A. W, Yonda"^

ABSTRACT: A simulation tailored to be responsive to

the needs of aerospace systems analysts has been developed. The simulation is written in FORTRAN so that it can easily be transferred between different

computers. The simulation is modularly structured

so that new functions can be simulated and easily

added. The level of detail of the simulation of

model elements can vary widely. When macroscopic

system aspects have been studied and optimal configurations established, then sub-systems of interest

can be modeled in greater detail and embedded in the

macroscopic simulation for m.ore detailed study. Thus

design engineers can study the performance of their

design in a total system environment before hardware

is fabricated.

The parameters of the simulation permit

the configuration of models rapidly. These can be

run against a scenario of input parameters which

when varied provide information on system performance. Thus, many different projected system configurations can be quickly exercised and discarded until only a few feasible system configurations remain. These then can be exercised with model elements being more detailed enabling a more comprehensive analysis. Provision is made for entry of errors into the system so that performance characteristics of the system in a degraded state can be

studied. The exercising of Executive Program strategies is provided for. Thus, a tool which is oriented toward supporting the systems analysis require

ments for studying dynamic system performance has

been developed.

'" Aerospace Systems, Equipm.ent Division,

Raytheon Company, Sudbury, Mass.

535
KEY WORDS: Simulation, space systems, system test

bed

1. INTRODUCTION

The reasons for developing this simulation are: To provide a design tool for design engineers and to

provide a means to support systems analysts in the

analysis of the operation of complex space systems.

Due to limited resources and the need to quickly furnish support to engineers designing a data bus, it was decided to provide a data bus simulation which

was general enough so that it could easily be expanded to simulate more comprehensive space systems.

The goals of the simulation are to provide design

engineers and systems analysts with a tool which is

easy to use and can quickly provide performance measures of the operation of asynchronously functioning

interrelated elements of the systems being studied.

To attain the goal of ease-of-use, a user oriented language for describing the input parameters

for simulation scenarios is provided. This input

language minimizes the users scenario definition effort by providing stylized English statements for ex- ercising control over the simulation process, and

symbolic type inputs to specify the system parameters

The goal of quickly providing relevant data in a timely m.anner is attained by designing the simulation

so that it accepts more than one level of detail.

That is, a sub-system designer can embed a detailed

scenario of his sub-system in a macroscopically defined scenario of the entire system. Thus he will be

able to run detailed experiments for his sub-system

in a total system environment with considerably less

delay than if he had to specify the total systems in

fine detail. An additional advantage accrues from

savings effected because of shorter computer runs. The sim.ulation is providing a viable test bed

which economically exercises experiments in a control

led environment. Thus the systems analyst and design

engineer are able to test out and validate concepts

before building hardware.

2. METHODOLOGY OF THE SIMULATION

The simulation is modularly structured and table

driven so that additional program modules are easy to

add. That is, the professional programmers who wrote

the simulation can quickly add new program modules to

respond to new requirements . In addition to this

type of modular construction, the simulation provides

536
for the user to write FORTFIAN programs which implement the interpretation of any event the user wishes

to specify for his simulation runs. Thus the simulation is flexibly structured so as to permit response

to a wide variety of requirements. The primary advantage in this type of organization lies in the inherent capability to respond to new rules unthought

of during -the initial simulation design.

The simulation is written in FORTRAN in order to

provide a tool which is easily transferred from one make or model computer to another.

A major requirement is to provide the systems

analysts with the means to easily vary simulation

parameters and thus be able to quickly reconfigure

systems which are under consideration. Hence, by

providing an easily varied system test bed at minimal cost, the systems analysts will be encouraged to

try bolder and more innovative system configurations.

How successful we will be in this endeavor has yet to

be determined.

3, SIMULATION STRUCTURE

The purpose of this section is to describe the

structure of the simulation program. Figure 1 is a block diagram of the simulation structure.

The simulation is divided into four parts. These

are

File Generation Processor

Preprocessor

Run Processor

Post-Processor

Their relationships are shown in Figure 1. The simulation starts with the hardware configuration, the

software configuration, and the scenario in the form

of punched card inputs. In addition, if the user

wishes to define an entirely new type of element to

augment the simulation, there may be a definition of

that new element in punched card form.

The file generation program takes the punched

card data and creates three magnetic tape files.

These are: Model Configuration File

Simulation System File

Scenario File

The model configuration file contains data on the

particular hardware and software configuration for

this run. The simulation system file contains the

definitions of the various units (e.g., processors,

memories, buses, and so forth) which are the building blocks of the simulation. The scenario file contains the definition of the scenario developed from

the input cards

537
The preprocessor, using the data on the three

files listed above, creates tables in the memory describing the configuration and the elements (e.g.,

processors, memories, buses, and so forth) for the

simulation run. It also creates an input event file,

which is a listing of the events based on data coming

from the scenario file. The scenario data can be the

statistics of these events, an actual listing of desired events, or a mix of the two. There is no particular limitation on the allowable granularity of

event times on the input file, and, in fact, the way

that the run processor operates, additional granularity does not result in an increase in running time.

It should be noted that the input events remain

on the input event file; therefore, if the analyst

wishes, he can run the same input event file while

making changes in the hardware configuration and soft- ware configuration. This means that if the performance of a configuration is sensitive to the exact

sequence of input events, rather than simply to

their statistical distribution, the analyst can re- . peat the identical sequence of input events , rather

than a different sequence of input events having the same statistical properties.

The run processor executes the actual simulation.

Our approach to the run processor is to make it

"event driven" , a technique which permits us to handle the simulation run in a fashion which is indendent of real time. This technique is similar to that

used in languages like GPSS and SIMSCRIPT. It does

so by using an event table in which are stored a list

of all events scheduled to occur in the future, in

order of their time of occurrence. Some of these

events in the event table come directly from the input event file. These include such things as incoming messages, requests for service, and external

interrupts. Other events in the event table are a result of system action taken due to input events.

For example, if a particular interrupt causes an application program to be started, the run processor

will compute the time in the future at which that

application program will terminate and will insert

that time of termination into the event list. Each

time the run processor "processes" an event, it

determines some number (possibly zero) of future

events, with their times, and puts these events into

the event table. Each time the run processor goes to

the next event in the event table, the "real time"

becomes the time of that next event. The actual run- ning time of the simulation is determined by the num- ber of events which must be processed, rather than by

538
the granularity of the measurement of time.

The run processor creates an output event file,

which is a record of every event that has occurred

during the simulation. The post-processor extracts

from that file the events and statistics which have

been requested by the user. The reason for creating

a separate output file, rather than creating the output data in the run processor, is that the simulation

run will be the most time-consuming part of the simulation and, hence, the most expensive in elapsed

computer time. After the completion of a run, the

analyst may wish to examine the output data and then

request other data for the same run. Rather than

execute the run processor again, the user can simply

change the input cards to the post-processor, which

will then extract the new output data from the output events file.

Running time is, of course, heavily dependent on

the particular scenario and on the size and complexity of the system configuration being simulated.

4. SIMULATION OUTPUTS

The reason for performing the simulation is to

track the operation of the simulated system under the

postulated scenario. The results of the run processor are much too detailed to be easily digested

by the analyst; hence, a post-processor is required

which can accumulate and reduce the data to a form

that is more useful.

As described above, the post-processor operates

somewhat as a report program generator, permitting

the extraction of any data at all which is on the

output file and the formating of that data in the

form desirable by the analyst.

There are essentially four kinds of reports which

can be extracted by the post-processor. These are: Event Rei; orts

Time Line Reports

Snapshot Report

Summary Data Reports

Event reports are simply listings of specified

events. Examples include listing of all interrupts

from a particular external sub-system, listing of all

executions of a particular application program, and

so forth. The analyst can get a listing for the

whole run, or he can have a listing of specified

events from time Tl to time T2.

Time line data is a report of a continuous variable. For example, the percent loading of one of the

CPU's as a function of time or the loading status of

a specified bus or data channel would be a time line

report. Of particular interest may be the timie line

539
of certain operational programs with an identification of the processor used, or a time line for a specified device.

The snapshot is a third type of output report.

The snapshot will consist of a complete report of

the status of some area of interest for example,

a complete map of a computer memory or a complete

listing of the operational programs running in each

of the processors. A snapshot may be taken at any

point during the run, including the end. Memory map

and processor task snapshots are particularly effec"

tive in comparison of Executive Program strategies.

The fourth type of output data is summary data,

which requires that some processing or filtering of

the event record be performed. Examples of this

type of data are a record of percent device loading

during a specified period of the simulation, the

executive program overhead add factor, frequency of use of specified programs, or frequency of use of

specified devices.

The four types of output data listed above provide the analyst with a complete and flexible tool

for determining the performance of the simulated

system.

5. SYSTEM ELEMENTS

The elements of a simulated system can range from

individual logic elements such as flip flops to

whole computers depending on the level of detail used

and the needs of the user. With event driven simulators, it is a simple matter to combine many different levels of detail into a single system model.

In our initial application, a relatively simple

data communication system was implemented. This

system is modelled at the level of the major functions

required to move messages from a central computer to

remote device controllers by means of a data bus.

The basic functional elements represented in the

data bus communication system are described in the

following sections.

5.1 COMPUTER

The computer exercises overall control over the

entire system. It determines the order and frequency of the transfer of data to and from all remote

devices. As the data transfer requirements build up

the computer maintains the queues of data messages

awaiting transmission and applies various scheduling

algorithms to select the next message to be transmitted . For multiple computer access to the data bus, a master computer, in addition to its normal message

540
processing also controls the data bus access of the

remaining computers in the system.

5.2 MASTER DATA TRANSMISSION UNIT

The Master Data Transmission Unit (MDTU) provides

the interface between a computer and the data bus.

It controls the movement of individual messages via

the bus to or from the remote controllers . The MDTU

initiates synchronization, performs parallel to serial

and serial to parallel data conversion, and performs

a clock time-out as a check against lost messages.

5.3 REMOTE DATA TRANSMISSION UNIT

The Remote Data Transmission Unit (RDTU) controls

up to thirty-two devices which can range from sensors

to other computers. Each RDTU can be addressed separately and devices at an RDTU can be addressed separately or in groups. The functions performed by the

RDTU include, address and message type decoding,

serial to parallel and parallel to serial data con- version , device addressing and the control of all

data transfers between itself and its devices.

6. SAMPLE MODELS

A candidate system model can be designed and

studied at various detail levels.

For example, consider the model shown in Figure

2. This model, when exercised for a range of traffic loads provides data on: the load on the bus

message delays

traffic scheduling effects

Next, the model is amplified so as to include more

detail to the extent of introducing errors into the

messages. For this configuration the model is as shown in Figure 3. Now, the simulation provides

data on: the effects of errors on bus

the effect on the load on the BUS because

of repeat messages

the effect on system performance because

messages are received by the wrong

addressee.

For this same configuration, the same model can be

further amplified so that the model as shown in

Figure 4 now includes stored programs and executive

programs. Here the simulation provides data on: the CPU performance for a variety of

Executive program strategies

the effect on system performance when

erroneous messages are received by the CPU's

the effect on system performance when the

compute power of any CPU is varied.

The detail level of each model can be further

541
amplified by putting more details in the program

functions. In this manner, systems can be studied at

various detail levels and when one model has areas that require greater depth of analysis, then the

model detail level can be expanded,

7. SIMULATION PAR?^METERS

The list of system variables which are parametric

to the simulation and can be easily varied by the

systems analyst follows:

CONNECTIVITY

Processors

Memories

• Data Busses

Sensors

Controllers

, PROCESSOR CYCLE SPEEDS

COMPUTER ASSIGNMENT OF MISSION PROGRAMS

DATA BUS

Bandwidth

Word Size

Block Size

Error Detection/Correction Techniques

Acknowledgement Techniques

NOISE ENVIRONMENT

MASS MEMORY SPEED

EXECUTIVE CONTROL FUNCTIONS

Data Bus Control Strategy

Priority Handling Techniques

Data Management Techniques

' Checkout Techniques

Corrective Procedures in Event of

Failure

Additionally, the following types of failures can

be simulated during a simulation run: SIMULATED FAILURE OCCURRENCES

Processor Failure

Memory Failure

Data Bus Failure

Out of Limit Sensor Reading

8. SIMULATION RESULTS

Results from a simulation run provide the following information:

PEAK AND AVERAGE BUFFER QUEUES

PEAK AND AVERAGE DATA BUS TRAFFIC

FLAGGING OF POSSIBLE SYSTEM FAILURES

DETAILED AND GROSS TRACES OF SYSTEM INTER^

ACTIONS (ON REQUEST)

. , PEAK AND AVERAGE CENTRAL PROCESSOR LOADINGS

AS A FUNCTION OF TIME

PERCENTAGE OF DATA BUS AND COMPUTATIONAL

BANDWIDTHS REQUIRED FOR OVERHEAD (EXECUTIVE)

FUNCTIONS

542
9. CONCLUSION

The simulation provides a viable tool to the

systems analyst which by virtue of being problem

oriented is highly responsive to his needs. The

building block structure of the simulation makes it

easy to expand the details of functions being simulated as well as to add new simulated functions un- thought of in the initial design.

By judiciously varying the simulation parameters,

it is possible for the systems analyst to derive

enough information so as to be able to do the following : Evaluate the performance of the specified

system configuration

Optimize system control procedures

Minimize redundant elements in the system

Establish limits of system performance

Validate data bus scheduling strategies

Evaluate many projected system

configurations

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Paper No. 38

GEKERAL PURPOSE CENTRALIZED AUTOMATIC TESTING FOR IMPROVEMENT

OF SHIPBOARD SYSTEM PERFORMANCE

F. J. Zupan-*-

REFERENCE: Zupan, F. J. "General Purpose Centralized Automatic

Testing for Improvement of Shipboard System Performance,"

ASTM/iES/AIAA Space Simulation Conference, l^-l6 September 1970. ABSTRACT: Increased operational effectiveness of Navy shipboard systems is the primary goal of the Navy automatic testing

program. This paper discusses the rationale and techniques

utilized in developing an experimental model which will help

update and modernize traditional methods and approaches to

systems monitoring and maintenance. Developments in acquisition^ processing and dissemination of information as related

to a centralized on-line performance monitoring system are

covered. Both laboratory and shipboard functional tests were

performed;, with a laboratory model system monitoring typical

Navy electronic and non-electronic systems. Test results, a discussion of problem areas, and a prediction of the impact of

centralized testing on the shipboard maintenance and operating

organization are included. KEY WORDS: operational readiness, performance monitoring,

testing, data acquisition, data processing, data dissemination,

maintenance

The Navy Electronics Laboratory Center in San Diego has

been engaged in exploratory development aimed at improvement of

ship systems operational readiness. Efforts of a technical

group which I represent are in response to a specific operational requirement which directs development of a Centralized

"^Performance Monitoring Systems Division, Naval Electronics

Laboratory Center, San Diego, California

549
Automatic Test System (CATS). This system is intended to monitor the performance of complex and militarily essential shipboard systems and the environmental conditions under which

these systems and personnel operate. A further requirement for

CATS is that it shall have the capability to predict and locate

failures in these shipboard systems.

A typical Navy Shipboard installation is comprised of many

complex systems; all of these are necessary and essential to perform critical functions. The proposed monitoring system

must somehow be able to converse with these ships' systems

while they are operating, without adversely affecting this

operation, A cursory consideration of these requirements might

lead one to conclude that the magnitude of the task in terms of

management problems, hardware complexity and the necessity for

software would make this a difficult endeavor. Why the" does

the Navy elect to establish such a series of requirements?

One reason is that in terms of potential long tenn payoff,

automation is a most promising approach to the Navy's need to improve or replace its outdated methods for performing shipboard

maintenance. In the modern Navy, reaction time to detect and

evaluate threats in combat ship operation has decreased, re- quiring more complex systems, increased automation and continuous optimum performance. Ship's comomnd requires timely information regarding status and level of performance of systems

in order to select from possible options such as alternate

modes of operation or use of alternate systems to assure mission

effectiveness. Extensive studies have shown that traditional

manual methods of performance monitoring and fault detection

are too cumbersome and time consuming. New techniques and

methods must be adopted or developed to meet the requirement

for quick reaction.

APPROACH:

Efforts at the Naval Electronics Laboratory Center with

which we have been associated have been directed toward improving the monitoring and fault location tools and methods of the

Navy technician, who is assigned responsibility for improving

the performance of shipboard systems. His present aids which

include manual portable test equipment, instruction manuals,

built-in indicators, and localized testers, are insufficient in

themselves to meet contemporary and advanced requirements for

up to the minute system performance monitoring information.

Since the technician is the ultimate source for systems status

inforatation which is eventually the basis for tactical and

strategic decisions of command, improvements in his tools and

methods will result in better status reporting and improved

command decision and action.

In order to determine the effectiveness of Centralized

Automatic Test Equipment with respect to General Purpose Manual,

550
Built-in^ and Localized Automatic Test Equipment ^ the Naval

Ship Systems Command directed that a special study be performed.

The Naval Applied Science Laboratory conducted the study(l)

which involved consideration of mixes of these support equipments. Mission simulation and mathematical analysis techniques were utilized. It was concluded that in the setting of

typical situaticns studied;, support mixes utilizing up to 8ofo centralized automatic test equipment were the most effective

in terms of monitored systems availability and overall cost effectiveness. This study and other supporting evidence docu- ment the feasibility and effectiveness of centralized automatic

testing aboard ship.

DESIGN PHILOSOPHY:

One of the most publicized criticisms of automatic testing

systems developed in the past has been the enormous complexity

and cost of software associated with these systems. In organizing and planning an automatic testing approach, means have

been sought to make efficient use of automation. Since the

earlier period^ favorable advances in programming techniques

as well as in hardware state of the art;, have been helpful in

achieving this goal. A substantial reduction in software re- quirements for our system has been achieved by application of

the fundamental principle of retention of the man in the testing loop. Beginning with the conceptual model and continuing

through the development period of a laboratory model system,

the design philosophy has emphasized the man and his capability

to exercise technical judgment based on a combination of the

information presentation, his technical training, and his

technical experience.

OATS is a computer-centered system which utilizes adaptive

man-machine interfacing as an effective approach to the Navy's

need to update ship's systems operational readiness. Displays,

controls, and readouts are being planned and specified to ex- ercise the man's capabilities and to eliminate or reduce routine, time-consuming tasks which increase maintenance and

reporting times. System philosophy is such that a wide range

of operator or machine control is possible, depending on the

monitoring problem and/or the capability of the man. Tracing performance data in monitored systems from the

source to the ultimate shipboard user, capabilities may be

grouped under the following broad categories: Information

processing, presentation, recording, and system control and

operation. These capabilities have been incorporated into a system that is designed to perform services for the ship's

system, operators, the ship's maintenance people and the ship's

The number in parenthesis refers to the list of references

appended to this paper.

551
captain. In order to be effective the test and monitoring

facility must supply this information when it is needed and in a form that is understandable to each of the various users.

DESCRIPTION OF SYSTEM STRUCTURE AND OPERATION:

The laboratory model of a centralized automatic test system has been tested under laboratory conditions while interfaced with several typical Navy equipments. Additional similar tests were also conducted aboard the USS NORTON SOUND.

Figure 1 is a simplified functional block diagram of the model

which was developed at the Naval Electronics Laboratory Center. Data utilized in the system is acquired from sensors (l) installed within the monitored prime equipment. These sensors (actually a part of the prime equipment) convert the parameters

of interest into proportional d.c. voltages. These d.c. volt'- ages are multiplexed in the Data Acquisition Unit (2) under

control of the Central Processor (8) and are converted into a pulse duration. The pulse thus developed is transmitted to the

Distribution and Data Conversion Unit (5). Conversion to digital form is accomplished in that unit to facilitate inputting to the Central Processor. The Central Processor formats, treats and compares this raw data, referring to the Mass Memory (t)

for stored data pertaining to tests. Processed data is sent to the Operator Control Console (9) for display in a continuously updated manner and/or to the teletypewriter (h) for print

out, and periodically is sent to the Status Panel (6) for use of controllers. The operator at the Control Console can initiate new data acquisition sequences, directing the Central

Processor to address data to the various Data Acquisition Units

via the Distribution and Data Conversion Unit. In the event of

casualty to the Central portion of CATS, the Auxiliary Test Unit

(3) can be used to access test parameters from sensors and to convert them to numbers which can be compared to pre -printed

limits. This brief description is merely intended to convey

the fundamentals of operation. The test system is capable of

handling a wide variety of input parameters under actual operating conditions. Figures 2-k show various equipments which

comprise the CATS hardware.

INFORMATION GATHERING:

With this description as background let us examine the

functional capabilities listed earlier, in terms of hardware

and software development effort, conducted at the Naval

Electronics Laboratory Center.

Some of the most challenging work performed during the

entire program was the development of the data acquisition system. Work in this area spanned problems in both the systems

monitored and in the CATS data gathering and monitoring faciLrty .

552
Because of the critical nature of this effort, various data

acquisition techniques were considered. Criteria used to gauge

the effectiveness and suitability of various elements of the

data acquisition system include the following: Cost, reliability, simplicity, speed, size, maintainability and repeatability.

To emphasize the effectiveness of automatic data acquisition it

might be instructive to compare the process with the manual

method. In the latter case when equipment fails a technician

commences a prescribed or improvised trouble -shooting routine

fitted to his capabilities and those of the available manual

test equipment ( multimeters oscilloscopes, signal generators,

etc.). As he begins to probe for the defective area in a bewildering array of tightly packed circuitry he faces thefraquent

possibility of inadvertently shorting critical circuit terminals.

His ability to acquire and evaluate data is such that several

minutes per measurement are required. Automatic data acquisition circuitry developed for CATS supersedes these manual techniques with a substantial increase in data rate;, and no damage

from manual probing. Testing is accomplished at rates of 200

to ^00 per second. Acquisition in the laboratory model is

tailored to the unique requirements of automatic perfonmnce

monitoring;, comparing favorably with other candidate schemes

which were evaluated against design criteria listed earlier.

However, flexibility of CATS is such that advantages of state

of the art developments in data acquisition can readily be incorporated. For example ;, if tradeoff studies indicate accuracy

is miore important than cost for a particular application, a potential candidate scheme might be a highly accurate analog

to digital conversion facility located close to the monitored

equipment. Specific details concerning data acquisition developments at the Naval Electronics Laboratory Center are available . (2)

Successful monitoring of shipboard systems is a function

of both the test system and of the test logic design and sensor implementation in the tested system. Recognizing the need for

some elements of standardization in test logic design and sensor implementation for shipboard systems, the Naval Ship Systems

Command directed preparation and issue of Military Standard

1326 (Navy).(3) This document is intended to provide manufacturers of automatic test systems and shipboard operating systems

with a standard guide to insure effective and compatible interfacing.

INFOEMATION PROCESSING:

In CATS, raw data acquired from monitored systems is processed against stored data pertaining to the various tests being

The numbers in parenthesis refer to the list of references

appended to this paper.

553
performed. Routine processing functions are performed automatically. In keeping -with the overall system philosophy which

attempts to make optimum use of the man^ selection of displays

on the central display device, and selection of other peripherals;, as well as interpretation and analysis of the information

presented, are manual fmictions.

lEFORMATION PRESENTATION:

Information is disseminated to various users in a variety

of display formats. Operational users are concerned with over- all availability and performance of systems; therefore, readouts^ status indicators and status type displays are the chief means of information presentation. Users directly concerned

with maintenance are provided with several levels of information detail on the cathode ray tube display. Four general

forrnats are presently available in the system (refer to Fig.

5-8). These illustrations were developed by photographing

the CATS Central Display Device imder operating conditions.

Typical shipboard radar, communication and mechanical type sys- tems were monitored. Identification of equipment as noted is readily understood by Navy operating and technical personnel. For example, AN/wRT-2 is the nomenclature for a generally used

communication transmitter.

Informational: This type is static and is primarily used

to convey static data of any type to the operator, Lists,

tables and indices are examples of this type of display.

Bar Chart: This type combines static information, such as shown on the informational type display, with dynamic information resulting from performing tests on equipment being monitored. Test results are shown in the form of horizontal

vectors, the length of the vector being proportional to the

percentage that the measurement departs from the normal toward

either the upper or lower limit for the test. The program can branch on the result of any test so that if a test is within

tolerance one sequence of tests will be conducted, while if

that test is out of tolerance, a different series of tests will be conducted. As vectors in the bar chart reach the

length that represent 100^ deviation from the norm, the vectors begin to flash. The 100^ perm:*ssible deviation point is shown

by the vertical cursor line.

Block Diagram: This type is similar to the bar chart ex- cept that in place of a list of tests a block diagram of the

equipment or portion thereof is shown. Vectors are used as in

the bar chart in order to show result of tests. Both the block

diagram and the bar chart may have light pen targets for calling up other displays or tests and may have other displays

referenced by number, indicating to the operator which displays to call up for continuation of signals or for more detailed information on areas of possible interest.

m
Diagnostic Display: This type is valmble for isolating

faults when it can be predetermined that a series of out of

tolerance tests infallibly indicates a particular fault. The

diagnostic display shows a number of test numbers with dashed

lines connecting them vertically and horizontally. As a test

is performed, its vector is updated and if the test is in toler- ance a solid line appears connecting the test number to the next

test number to the right. As the next test is performed its

vector is updated and a solid line is drawn to the next test to

be performed. If a test is out of tolerance a solid line will

appear connecting the test number to the number vertically below it. In this way tests are performed and the path through

the diagnostic "tree" is indicated to the operator. Following

the last test the solid line will extend to a block indicating

the probable cause of failure or the fact that the equipment is

in a stand-by, off, or no-fault condition. The operator has

the option of calling for a rapid fault diagnosis sequence or

one that progresses one step at a time with the capability of

overriding the program's decision. Display formats described

in this paper are typical; however, system flexibility permits

a variety of others, - for example, logic diagrams.

Il^ORMATTON RECORDING:

During the normal operation of CATS a large amount of data

is gathered from various monitored systems. Properly used, the

time history recording of this data is valuable in supplementing the human capabilities. In designing the system, attention

has been given to implementation of this function. Comparison

of current readings with past readings taken at various time

increments is made possible by incorporation of the capability

to retain desired sensed data in the memory store. How this

data is used will be covered in detail under the discussion

on Trend Analysis.

CONTROL OEERATION:

Functional controls in CATS are designed to complement

machine functions and operator needs. The system can operate

in both the unattended mode and the attended mode. In the

unattended mode only discrete information such as "good", "bad"

"marginal", or "indeterminate" is available. Alarms, both

visual and aural are activated on the basis of this discrete

machine assessment. In the attended mode more detail is available, including level of performance information for each test,

combination of tests, and system tests which the operator

interprets and analyzes. For example, a typical bar chart

shows the immediate general status of the system it represents

by a pictorial representation of the system in terms of dynamic

555
vectors shoving test results with respect to nominal values,

high limit values, low limit values and allowable deviation

from nominal. Controls in the form of function buttons, keyset,

typewriter keyboard and light pen provide a versatile array of access means to the computer program. Thus the operator is integrated with the machine in a manner which places emphasis

on his ability to think and to exercise technical judgment.

TREKD MAiYSIS: . - Prediction of system behavior is a complex and many

faceted subject. What has been accomplished as a result of

technical efforts in our automatic testing program is merely an initial approach to a very broad problem. However, con- trasted to present methods, it appears that several techniques

not yet fully developed and implemented, have attractive fea- tures with the potential for practical application and long

term benefits. Past methods are based on failure mechanisms

and physics of failure studies which are costly and time con- suming to implement and perhaps are not completely related to the particular system installation and its actual operational

environment

The prediction process is based on the capability to autoimtically and rapidly acquire, process and disseminate

parametric data concerning the monitored system. Systems behavior is enhanced by the fact that on line monitoring may

detect equipment and systems degradation before incipient fail- ures become serious or catastrophic. This feature permits

maintenance personnel to plan ahead. Repair or adjustment can be accomplished during scheduled preventive maintenance periods

and therefore major casualties can be avoided, or at least reduced. In a broad sense, command and maintenance supervisory

personnel are provided a gross behavior prediction capability

by observation and study of the up to the minute display

formats available on the cathode ray tube.

Several operator-oriented trend-analysis display formats are used in CATS. The operational program provides this

capability by the technique of mark, store and recall of dynamic displays, including bar charts and block diagrams. By

noting the overall change in the display or specific portions

of the display on the basis of time, an estimate of probable

equipment behavior can be projected. Figure 9 exemplifies

this feature. A second method of trend analysis depends upon

the use of the historical recording subroutine. Presently the

program is written so that each hour and for the ensuing 15 hours all test measurem^ents are stored in memory. This historical data can be recalled in various display formats and

analyzed by the operator as an aid to trend prediction.

Figure 10 is an example of this type display.

m
Actual experience in regard to systems behavior prediction was relatively limited during development assist tests with CATS

aboard ship. However, sufficient data was acquired to conclude

that further expansion of present techniques is warranted.

Principal recommendations which resulted from this study are

as follows:

(1) Increase the flexibility in rate and/or amount of

data recording. For example program modification to increase

the sampling rate from a given point following detection of an intermittent fault should provide better information to judge

the nature of the cause and prediction of possible complete

failure

(2) Optimize selection of data which is most useful in

coirelation and analysis. For example in the automatic mode

a program routine to collect and show the correlation of interrelated test point information would be helpful, once a most probable fault has been indicated. Putting it differently, a refinement of the program would permit concentrating both data

and assessment automatically in the test point areas related to

the difficulty, to provide a better basis for predicting the

probable behavior trend.

(3) Increase use of the computer in automation of the

analysis process. This of course involves a comprehensive

analysis of failure modes and the earliest available indications of them, subject to automatic treatment.

This is a potentially fertile field. So far we have shied

away from complete automatic analysis, as we mentioned previously, because of the size of this job of determining failures and most probable causes in an all encompassing manner.

If the job of predicting faults and their causes in advance

of experience were a straight forward process, we would agree

that we should correct the deficiencies in advance rather than

test for them. Some techniques related to computer aided design are undoubtedly relevant to computer aided fault analysis

and we would like to explore this area more extensively in

the future. In considering the potential of system behavior prediction

the value of information assessed from analysis of a number of

systems would result in increased assurance that a particular

trend is typical of a family rather than an isolated case. Recorded information could be analyzed to determine the patterns

or trends. With appropriate statistical treatment, specific

mathematical functions could be derived. When programmed into

CATS automatic recognition of behavior patterns and trends

would be achievable.

557
PRESENT STATUS AND FUTUEE TRENDS: Introduction of on-line automatic testing into the shipboard

environment must be considered in terms of two general time

frames. The present is called the retrofit time frame, one in

which existing systems are not necessarily fully amenable to interfacing with an automatic test system. The CATS laboratory

model has been interfaced with typical shipboard equipments in

this time frame requiring an appendage -type data acquisition

scheme. Tests indicate results are favorable despite the relatively unfavorable retrofit conditions.

Full exploitation of automation can be achieved during the new equipment time frame because automatic testing and new

equipment interfacing problems can be effectively solved as a part of the design life cycle ^ rather than as a retrofit. At that point in time an optimized interface between tested and

testing systems can be expected. System planning for automatic

testing will be more integrated and correlated by the application of design principles discussed in this paper. One cannot minimize the technical problems which impact on effective implementation of centralized automatic monitoring. Questions

like the following need to be studied in greater detail:

(1) What . is the optimum configuration of monitoring and

testing hardware for various type ships?

(2) Who needs displays and what type?

(3) Should the system use dedicated processing or is

shared processing more desirable?

Requirements of various shipboard maintenance departments and

other user organizations need to be examined in terms of overlap and redundancy.

SUMMARY AND CONCLUSION:

This paper outlines the requirement for shipboard automatic

testing; describes the development effort and philosophy of

design for a centralized automatic test system; and points out salient features of CATS. Predicted long term worth of centralized automatic testing indicates that it will reduce down time

for monitored systems; speed up the maintenance process; provide an increased system mean time before failure; reduce proliferation of test and monitoring facilities aboard ship; and

generally increase the overall readiness of the ship.

Because of the revolutionary change to traditional ap- proaches and methods, both technical and management problems

must be solved. The laboratory model has demonstrated effectiveness during the difficult retrofit time frame. Application

during the new equipment time frame will exploit the full cap- ability and effectiveness of this concept.

558
REP^ERENCES

(1) System Formulation for a Centralized Automatic Test System (CAIS) Vol. 1 (Sk6-23) July 1, I968.

(2) Development of Centralized Automatic Test Systems at

KELC, Vol. 3, Data Acquisition, 3 Feb. I969.

(3) Military Standard I326 Navy Test Points, Test Point

Selection and Interface Requirements for Equipments Monitored

by Shipboard On-Line Automatic Test Equipment.

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PERCENT or ALLOWABLE DEVIATION

0 100X 200X

TEST NO. 003-6 114

PREVIOUS DISPLAYD

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H «^.«420

FIGURE 10 568
Paper No. 39

EXPLOSIVE OXIDATIONS INITIATED BY SIMULATED

METEOROID PENETRATION INTO SPACECRAFT ATMOSPHERES

William H. Carden

Assistant Manager, Aeroballistics Branch, VKF

ARO, Inc., Arnold AF Station, Tenn.

ABSTRACT: Hypervelocity impact tests were conducted

in an evacuated range on target samples which simupLated a NASA S-IVB wall configuration since this

btage is planned as the primary structure in the NASA

Orbiting Workshop program. The samples formed part

of the wall of a large tank which contained an oxygenrich atmosphere. The explosive oxidations which oc- curred inside the tank as a result of perforation

were observed and the results were analyzed. The bare

thermal insulation on the inside of the wall further

enhanced the otherwise severe reaction which occurs

with a metallic wall in the presence of enriched

oxygen

KEY WORDS: hypervelocity impact, meteoroid penetration, spacecraft, combustion front, oxygen atmosphere,

oxidative detonation, oxidative flash, blast over- pressure

INTRODUCTION

The perforation of the wall of a vessel by a hypervelocity projectile will produce an explosivelike reaction inside the tank if the tank contains an oxygen-rich atmosphere. Studies (1-10) have shown

that the extremely hot fragments of projectile and

wall materials rapidly undergo a chemical reaction

with the oxygen in the immediate vicinity of the per- - foration, producing a combustion front which penetrates into the tank. This violent reaction is ac- companied by a brilliant flash of light.

The process just described is similar to the

situation which could exist if the wall of a manned

space vehicle were to be perforated by a meteoroid.

There is ample reason for concern regarding the effect of such an encounter on the human occupants of

^The numbers in parentheses refer to the list of

references appended to this paper.

569
the space vehicle. Among the phenomena which could

present a serious pathological hazard to human occu- pants are the blast overpressure, the intense light

[flash, and the high temperatures associated with the

icombustion front. The possibility that a spacecraft

ifire would be initiated by the combustion front and

jaccompanying hot particles is a secondary, but perhaps

equally serious, potential hazard. Ways to alleviate

these problems are discussed in Refs. 9 and 11.

Roth (8) has presented a literature review and

jsurvey of this flash oxidation hazard from the biomedical point of view. Much of his discussion is

focused on the work of Gell and his colleagues at

Jjing-Temco-Vought (1-3) . These workers exposed white

irats to this oxidative detonation inside a vessel

Ivhich was perforated by a high-velocity projectile.

These animals suffered various degrees of injury and

incapacitation, and under some test conditions mor- tality occurred. It was impossible to determine

whether death occurred primarily as a result of the

blast effects or as a result of burns received by the

animal

Several recent studies (4-7) have produced addiItional information regarding the mechanism of this

'oxidative reaction and its potential hazards to spacecraft structures, systems, and personnel. However,

sail questions have not been answered, and any addiItional experimental information which becomes availiable should be welcomed by those concerned with spacecraft design.

In a recent test program (9) at AEDC, hypervelocLty impact tests were conducted on target samples

which simulated a wall configuration of the Saturn

S-IVB stage. The S-IVB stage is planned as the primary structure in the NASA Orbiting Workshop program.

On many of these tests, the target sample was attached to a large tank which permitted the rear of the

wall sample to be exposed to an oxygen-rich atmosphere during perforation. The main purpose of these

tests was to evaluate methods of preventing or minimizing combustion of the polyurethane foam which

covered the inside of the wall sample. As a result

of these tests, the decision was made to install a 3-mil aluminum foil f ire-retardant liner to cover the

insulation and to install a micrometeoroid bumper on the vehicle (11). At the same time, however, these

tests provided an opportunity to observe the blast

effects inside the test tank at the time of perfora-

^i<Mi. The combustion front was recorded photographi570
cally on a large number of these tests. In addition,

on a limited number of these tests, transducers were

placed inside the test tank to allow further observation of the phenomena associated with the combustion

front. The purpose of this paper is to describe

these observations and to relate them to previously

available information on the subject. The present

results, although limited in extent, provide information which applies specifically to an uncovered polyurethane insulation and unbumpered wall configuration

TEST CONFIGURATION

LAUNCHER AND RANGE

These tests were conducted in the Hypervelocity

Impact Range S-2 at AEDC . The range includes: (lO a two-stage launcher; (2) a blast tank into which muzzle gases expand and in which the projectile (1/8 in.

diam., type 2017 aluminum spheres in these tests) is

separated from the sabot; (3) a connecting tube along

which velocity measuring shadowgraphs are located;

and (4) a target tank where the test specimen is impacted by the projectile. The large size (6 ft. diam.

by 21 ft. long) of this target tank permits the impact testing of unusually large configurations in a vacuum environment. In the present test, it permitted the use of a relatively large (3 ft. diam. by

ft long) test tank whose volume was greater than that

of any vessel used in previous explosive oxidation

impact studies (1-7) . TARGET CONFIGURATION

The wall specimen used on this test was designed

to simulate the Saturn S-IVB basic wall structure. It

consisted of a 0.125 in. thick, type 2014-T6 aluminum

sheet to which a 1 in . thick layer of closed-cell

polyurethane foam was bonded. The foam was sealed

with a layer of glass fabric and a coat of polyurethane resin. A photograph of the rear of a wall specimen is shown in Fig. 1. This polyurethane insulation provides thermal insulation on the inside of the

tank to prevent liquid hydrogen boil-off for the propulsive stage.

The target specimen was attached to the upstream

end of the test tank, with the foam insulation on the

571
inside. The test tank was mounted inside the range

target tank with its axis on the range centerline.

Provisions were made for evacuating the test tank

prior to charging it with the desired gas mixture.

INSTRUMENTATION

The combustion front inside the test tank was recorded photographically on a large number of those

tests reported in Ref . 9. The tank was equipped with

a side port (approximately 10 in. diam.) immediately

behind the upstream tank flange, providing a view nor mal to the tank centerline. The uprange edge of this

field of view was about 4 in. behind the rear of the

wall specimen. In addition, another port was located

in the rear end of the test tank, providing an un- obstructed centerline view of the rear of the wall

specimen. The side port was used most often for this

high-speed camera coverage, because the rear port was usually occupied by other instrumentation systems. A

16-mm camera was used, and the framing rate during

the event was usually about 3200 frames/sec.

For the shots which are of particular interest

in this report, additional instrumentation devices were placed inside the test tank to record the passage of the combustion front. This additional instru

mentation consisted of four quartz pressure transducers and four lead-zirconate piezoelectric "timeof -arrival" gages. These devices were mounted in

four pairs at 12 in. intervals longitudinally in the

tank. Each pair consisted of one of each type of

gage. The devices were attached to a bracket which

was cantilevered from the rear end of the test tank.

The bracket was shock-mounted at its point of attachment to the tank. A sketch of this arrangement is

presented in Fig. 2. The outputs of these devices were amplified and recorded on magnetic tape.

Three tests were also conducted with iron-constantan thermocouples located inside the test tank.

The thermocouple junctions were resistance welded

using 0.002 in. wires. These thermocouples were fabricated using a standard thermocouple connector and

ceramic insulator as shown in Fig. 3, and the devices were mounted on the bracket inside the test tank with

the exposed junction facing forward. The thermocouples were located at 8 in. intervals as shown in

Fig. 2. The reference junctions were immersed in ice

baths, and the thermocouple outputs were displayed

and recorded on oscilloscopes.

572
The instant of projectile impact on the wall

specimen was monitored by means of a radiometer which

observed the initial impact flash. The radiometer

output was recorded on the magnetic tape to provide a time reference for the pressure gages and time-ofarrival gages inside the test tank. The radiometer

output was also used to trigger the oscilloscopes

which were used to display the thermocouple outputs.

TEST RESULTS

TEST CONDITIONS

For the tests of interest in the present report,

the test tank was charged with the desired gas mixture to a pressure of 5 psia. The test gases were

nitrogen and oxygen, and individual tests were con- ducted with specific volumetric mixtures of 21 percent oxygen (air) , 45 percent oxygen, 59 percent oxygen, 70 percent oxygen, as well as 100 percent oxygen

or nitrogen. All tests of interest were conducted

using 1/8 in. diam. aluminum spherical projectiles at

velocities in the 20,000-26,000 ft/sec regime.

The target configuration previously discussed

was used on all tests of interest in this report. Impact usually occurred within the target area covered

on the rear by the polyurethane foam. However, on a few shots, including two of the shots made with the

additional instrumentation devices inside the tank,

target perforation occurred outside of the area covered by the foam. These shots provided a fortunate

opportunity for comparing the test results for an insulated wall with the corresponding behavior for a simple aluminum wall.

PHOTOGRAPHIC OBSERVATIONS OF OXIDATIVE FLASH

Typical film sequences of the oxidative flash

occurring inside the test tank at impact are presented in Figure 4. Film exposure resulted from luminosity inside the tank produced by the rapid oxidation

of projectile and wall materials. The film clip at

the left depicts perforation of the 1/8 in. thick

aluminum wall without foam insulation. The duration

of the event recorded on the film is about 6 msec.

The second, third, and fourth film clips depict perforation of a typical wall specimen for the case of

573
45, 70, and 100 percent oxygen mixtures, respectively.

The complete duration of the events on these tests,

of which only the initial portions are shown in the

figure, were 34, 44, and 53 msec for the 45, 70, and

|L00 percent oxygen mixtures, respectively. The exi)osure of the pictures indicates qualitatively that

the intensity of the combustion front increases with

increasing oxygen concentration. On other tests with

5 psia of air in the test tank, film exposure was too

^eak for printing. On tests using 100 percent nitrogen, no film exposure occurred.

The fifth film clip in Fig. 4 was obtained

through the rear port of the test tank instead of the

pide port. The camei^a was focused on the rear of the

itest specimen, the outline of which is visible in the

first frame. The sharpness of focus is reduced as the flame front propagates into the tank. Duration

of the recorded event was about the same as for the

adjacent film clip (100 percent oxygen). This view,

unrestricted by port size, indicates that the flame

front penetrates throughout the test tank.

These film clips demonstrate clearly the effect

of the polyurethane foam on the oxidative flash. In

the absence of the foam, the metallic wall and projectile provided the only particles available for

oxidation. Furthermore, production of these particles

is complete within a relatively short time interval

.after impact. However, with the addition of polyurethane foam to the wall, a large volume of lowdensity, highly combustible material is ejected into

the oxidizing atmosphere inside the tank at impact.

Also, the foam continues to be ejected for a long

period of time compared with the single metallic wall

condition. The overall effect of the presence of the

foam is to significantly increase (by an order of

magnitude) the duration of the oxidative flash inside

the tank.

VELOCITY OF COMBUSTION FRONT

The "time-of-arrival" gages inside the test tank

responded to the phenomena inside the tank. Typical

oscillograph playback traces of the recorded data are presented in Fig. 5. The results shown in Fig. 5a are typical for those tests for which impact occurred

within the area covered by the foam. The results for one of the two shots which missed the foam are shown

in Fig. 5b. In the latter case, the type of response

574
obtained from these gages indicates that the leading

edge of the disturbance is well defined as it passes

through the tank. In the case of impact within the

foam-covered area, however, the leading edge is somewhat "smeared out" by the interaction between the

foam and the high-speed metallic particles.

The time interval from projectile impact to ar- rival of the disturbance at each gage location inside

the tank was read from the oscillograph traces as indicated in Fig. 5. In the case of impacts through

the foam, an attempt was made to define the instant

of arrival of the strong part of the front instead of

its weak leading edge. A degree of judgement was re- quired in the definition of this ^'instant because of

the smearing of the front. These results are presented in Fig. 6. Although there is some scatter in

these results, the data clearly indicate slopes on

the distance-time plots corresponding to a propagation velocity of about 8000 ft/sec for impacts which

perforated the foam and about 22,000 ft/sec for impacts which missed the foam. These results appear to

be independent of the oxygen concentration in the

test tank.

It is evident that the experimental distancetime measurements shown in Fig. 6 do not extrapolate

back to the origin of the plot. This is largely be- cause the bracket which held the instrumentation devices was mounted approximately 8 in. below the tank

centerline. Since the disturbance front which propjagated through the tank was non-planar and was, to a first approximation, hemispherical in shape, its ar- rival at a gage position could occur significantly

later than its arrival at the corresponding centerline position. The general shifting in time of the

results from shot to shot probably resulted from tht

random dispersion of the projectile impact location

with respect to the tank centerline. It is believed

that the large upward shift in the results for shot

300 (Fig. 6e) was the result of a premature impact

flash signal from a small particle preceeding the

projectile rather than the effect of a change in gas

composition.

The velocity of the disturbance which produced

the time-of -arrival gage response does not correspond

ivith the velocity of cloud motions as observed on the

16-mm film (Fig. 4) . In the later case, for impacts

which perforated the bare foam, the leading edge of

the luminous cloud which initially passes across the

575
camera field of view has a velocity of about 800 ft/

sec, a full order of magnitude less than that indicated by the gage response. Therefore, there is reason to conclude that the billowing luminous clouds

which are visible in the 16-mm camera pictures represent a stage of the impact process which is fundamentally different from the initial disturbance stage

These differences are worthy of further study.

PRESSURE MEASUREMENTS

When a thin wall is perforated by a hypervelocity

projectile, the expanding debris cloud immediately

behind the point of impact will contain fragmented,

melted, and vaporized materials. It is convenient to

visualize that a "pressure" exists within this ex- panding cloud. More precisely, the cloud contains

momentum which produces an impulse acting on any sur- face the cloud strikes. If the cloud is completely

vaporized, thereby consisting of a very large number

of very small particles, the usual concept of pressure may be valid for time scales and space scales of

interest. However, appreciable vaporization cannot

be produced in most materials, including those of the

present test, at velocities attainable with light-gas

guns. Therefore, the debris cloud in these tests

consisted of discrete solid and molten fragments instead of a" "continuous" expanding vapor. Nonetheless,

a piezoelectric pressure transducer located within a few projectile diameters of the impact location in

the present tests would likely observe an apparent

stagnation pressure on the order of a few kilobars

(12,13).

The stagnation pressure in the debris cloud de- creases as the cloud expands away from the target

sheet. Furthermore, the radial pressure profile in

the cloud falls off very rapidly with increasing distance from the axial centerline. The instrumentation

bracket inside the tank (Fig. 2) on the present test was located below the tank centerline not only to prevent transducer damage from fragments within the

debris cloud but also to minimize any contribution to

the measured pressure which could result from the debris cloud rather than from the combustion front. It was hoped that any pressure observed by the transducers would result entirely from the explosive oxidation of the wall and projectile materials, the effects of which were expected to propagate throughout

the tank.

576
None of the pressure transducers inside the tank

on these tests recorded an overpressure exceeding the

noise level generated by mechanical shock at impact.

This mechanical noise level never exceeded the equivalent of 5 psi and was less than 3 psi for most tests.

Therefore, these results imply that any overpressure

which may have existed at the transducer locations

never exceeded 5 psi

This result is somewhat surprising in view of

reported maximum overpressures of 38 psi (3) , 100 psi

(4), 75 psi (5), and 150 psi (6) obtained in previous

investigations. It should be noted, however, that

all of the investigations of this type have involved

widely differing test configurations (projectile, target, tank volume, transducer location, etc.). Furthermore, the use of a piezoelectric transducer in a me- chanical shock environment is no easy task since the

crystal responds to acceleration as well as pressure.

Finally, it is possible that some investigations may

have indicated abnormally high overpressures because

of the interaction of the expanding debris cloud with

the blast overpressure. Nonetheless, it is not possible to explain with certainty the large difference

in blast overpressure between the present measurements

and those of other investigators. Nor is it possible

to identify which particular sets of measurements, if

any, are in error. The large differences between observations illustrate the need for additional testing

to resolve the present uncertainty.

TEMPERATURE MEASUREMENTS

Temperature measurements were obtained on two

tests using a 70 percent oxygen atmosphere and one

test using air. An oscilloscope trace of the output

of one thermocouple on a test with 70 percent oxygen

is shown in Fig. 7. A rapid temperature increase is

observed to begin about 5 msec after projectile impact. The signal then reaches a plateau corresponding to about 1600 F within 15 msec. The indicated

temperature remains near this level for about 35 msec

and then decreases steadily toward ambient conditions.

This type of response was typical for all thermocouples used on these tests.

The maximum indicated temperatures recorded on these three tests are summarized in Table 1. The

second thermocouple consistently recorded the higher

temperature, indicating a spatial variation in maximum

577
temperature inside the tank. It is impossible to determine whether the impact velocity difference or the

spatial variation in temperature was primarily re- sponsible for the lower maximum temperature which oc- curred on the second test with 70% oxygen. It is

clear that the indicated temperature is strongly dependent upon the oxygen concentration inside the test

tank. The implication is that most (if not all) of

the heating which occurs after impact is a direct result of the rapid combustion of impact debris. The

contribution of the individual components (insulation

and aluminum plate) to the heating effect was not

determined

DISCUSSION

The phenomena which occurred when the test speci- men was perforated on these tests can be separated into two categories. The first category includes those

phenomena which are associated with the impact and

perforation events only, i.e., those which would occur even in the absence of an oxidizing atmosphere. The

second category includes the additional phenomena

which occur because of the presence of oxygen inside

the test tank.

The first category includes the formation of the

debris cloud of fragmented, melted, and vaporized pro-^

jectile and wall materials which expand into the test

vessel. The leading edge of this debris cloud is

thought to have produced the initial response of the

time-of-arrival gages on these tests. The presence

of oxygen in the test tank apparently had little effect on the initial character of the debris cloud

since the propagation velocity of the disturbance was

the same with or without oxygen. On the other hand,

the character of the debris cloud was obviously influenced by the presence of the polyurethane foam insula^

tion which was bonded to the inside of the aluminum

wall

Although the oxygen concentration did not affect

the initial character of the debris cloud, the quantity and state of the materials contained within the

debris cloud strongly influenced the duration and

severity of the oxidation process. In particular,

the large quantity of finely divided, low-density,

combustible material ejected from the insulated wall

resulted in an order of magnitude increase in the

578
duration of the oxidative flash compared with the

jingle metallic wall condition. The billowing luminous clouds visible in the 16-mm film records are probably the gaseous products of combustion of the

foam with oxygen. The effects of these burning pro-

<3ucts on other spececraft materials is not included

in this paper. A limited study of secondary combustion is reported in Ref . 9.

The results presented herein have demonstrated

the possibility that an exposed and unprotected internal thermal insulation could aggravate the reaction

which would occur upon meteoroid perforation of an unbumpered spacecraft hull without a fire retardant

liner. The phenomena treated in the present paper

would occur only in the event of meteoroid perforation

pf the basic workshop wall. No attempt has been made

in this investigation to assess the probability of

such an occurrence or to evaluate methods of reducing

this probability. The probability of this event can Clearly be eliminated for all practical purposes by

the use of meteoroid shielding techniques. As a result of tests reported (9) , NASA elected to install

an external micrometeoroid bumper and an internal

fire retardant liner on the Saturn Workshop to alleviate these problems (11)

REFERENCES

1. Thompson, A. B. and Gell,.C. F. "Meteoroids as

a Hazard in Space Flight - A Survey of Present

Information." American Rocket Society Space

Flight Report to the Nation, October 9-15,

1961.

2. Gell, C. F., Thompson, A. B. and Stembridge, V.

"Biological Effects of Simulated Micrometeoroid Penetration of a Sealed Chamber Containing

Animal Specimens." Aerospace Medicine , Vol.

33, February 1962, pp. 156-161.

3. McKinney, R. and Stembridge, V. A. "Study of

Optimum Environmental Protection against Meteoroid Penetration." NASA CR- 50899, February

1963.

4. Caron, A. P. "Oxidative Detonations Initiated

by High Velocity Impacts." AFFDL-TR-65-41

May 196 5.

579
5. Gehring, J. W., Lathrop, B. L., Long, L. L., and

Hammitt, R. L. "Meteoroid Hazards as Related

to Internal Cabin Effects." Seventh Annual

Symposium on Space Environmental Simulation,

Seattle, May 23-24, 1966. Also, Long, L. L.

and Hammitt, R. L. "Meteoroid Perforation

Effects on Space Cabin Design." AIAA Paper

No. 69-365, AIAA Hypervelocity Impact Confe;rence, Cincinnati, Ohio, April 30-May 2,

1969.

6. Burch, G. T. "Hypervelocity Particle Penetration

into Manned Spacecraft Atmospheres." Boeing

Document D2-24149-1, May 1966.

7. Friend, W. H., Caron, A. P., (^t al. "An Investigation of Explosive Oxidations Initiated by

Hypervelocity Impacts." AFFDL-TR-67-92

August 1967.

8. Roth, E. M. "Space-Cabin Atmospheres, Part II - Fire and Blast Hazards." NASA SP-48, 1964.

9. Garden, William H. and Payne, J.J. "Simulated

Meteoroid Impact Testing on Wall Configurations for the NASA S-IVB Orbital Workshop."

AEDC-TR-68-246 (AD845231) , December 1968.

10. Garden, William H. "Simulated Meteoroid Impact

Testing on a Composite Expandable Structure

for Spacecraft Airlock Application." AEDCTR-69-14 (AD850512), April 1969.

11. "Developments Leading to Material Modifications

in the Saturn I Workshop Grewbay." NASA TM

X-53732, April 1968.

12. Friend, W. H., Murphy, C. L., and Shanfield, I. "Review of Meteoroid-Bumper Interaction

Studies at McGill University." NASA-GR54857, August 1966.

13. Tillotson, J. H. "A Theoretical and Experimental

Correlation of Hypervelocity Impact on Layered

Tragets." Edgerton, Germeshausen, and Grier,

Inc. Technical Report S-425-R, May 1968.

580
Fig. 1 —Wall Specimen (Inside View)

Fig. 3 —Thermocouple

581
g. 4 —Impact-Initiated Combustion Front

582

_l 1 1_

1 2 3 Longitudinal Distance, ft a. 100-percent Nitrogen

_l \ L

1 2 3 Longitudinal Distance, ft d. 59-percent Oxygen

Longitudinal Distance, ft Longitudinal Distance, ft c. 45-percent Oxygen f. Air, without Insulation Fig. 6--Time-of -Arrival Results

584
Fig. 7 —Thermocouple Output

TABLE 1--Temperature Measurements

Shot

Velocity

(ft/sec) Gas Composition Thermocouple

No. 1 (F)

Thermocouple

No. 2 (F)

1 26,300 70% O2, 30% N2 1075 1600

2 24,300 70% O2, 30% 825 1325

3 25,500 Air 200 300

The work reported in this paper was performed for the

George C. Marshall Space Flight Center and sponsored

by the Arnold Engineering Development Center (AEDC)

Air Force Systems Command (AFSC) , under Contract

F40600-70-C-0002 with ARO, Inc. Further reproduction

is authorized to satisfy the needs of the U. S. Government.

585

Paper No. UO

THE DYNAMIC ENVIRONMENT SIMULATOR - A MULT IENVIRONMENTAL MANRATED CENTRIFUGE

Dana B. Rogers, Captain, USAF^, Michael McCally, M.D.I ^nd

Klaus L. Cappel, B.S.2

REFERENCE: Rogers, D. B. , M. McCally and K. L. Cappel, "The

Dynamic Environment Simulator - A Multienvironmental ManRated Centrifuge", ASTM/IES/AIAA Space Simulation Conference,

14-16 September 1970.

ABSTRACT: The U.S. Air Force Dynamic Environment Simulator is

a man-rated centrifuge with multiple environment simulation

capabilities. This facility can generate combinations, either

simultaneous or sequential, of acceleration, vibration,

temperature, barometric pressure, gas composition, noise and

related environmental stresses. Simulation of a broad variety

of aerospace flight mission environments is now possible

using this facility. The man rating of this facility was

carried out in the last six months of 1969 and resulted in

the first manned run in December of that year. Although

presently programmed to meet Air Force data requirements for

the design of planned and proposed aircraft systems, the

Dynamic Environment Simulator represents a national research

capability capable of supporting the requirements of many R&D programs including manned space flight, high speed surface

transportation and other systems or functions where man is

exposed to physical environmental stresses in combination.

KEY WORDS: Dynamic Environment Simulator, man-rated

centrifuge, combined environmental stress, man-machine system,

simulation research

^Aerospace Medical Research Laboratory, Wright-Patterson Air

Force Base, Ohio.

'^Franklin Institute Research Laboratories, Philadelphia,

Pennsylvania.

587
The ultimate effectiveness of aerospace systems can be

measured only by criteria or parameters taken from the total

man-machine system. With the increasing complexity of present

and future aerospace systems, the need is manifest for ground

simulation of environments. The interactions of multiple

concurrent sequential stresses is yet undefined; and the

requirement for repeated exposure to mission environments

demands that a centrifuge designed for simulation research

should be capable of producing vibration, gas composition,

pressure, acoustic, thermal and acceleration variables which

may occur in aerospace vehicles during normal flight and

emergency periods. The necessity to man rate aerospace

simulation facilities cannot be avoided even though incorporation of man rating procedures adds significantly to the cost

and complexity of the facility. Safety, of course, is the

paramount consideration, and the well established medical

ethics of experimentation with normal human volunteers demands

complete fail-safe capabilities. In addition, the design

requirements for a safe man rated device include consideration

of mechanical, life support, control, communications, structures, emergency systems and associated medical facilities.

In 1960, after 18 years of continuous operation, the

United States Air Force human centrifuge facility at the

Aerospace Medical Research Laboratory, Wright-Patterson Air

Force Base, developed excessive vibration and deflections. A

study of alternative courses of action was conducted and

three options became evident: (1) structural reinforcement of

the existing quasi-operational device, (2) replacement of the current machine with an "off-the-shelf" centrifuge, or (3)

design and installation of a specific motion simulator clearly

based upon Air Force requirements evolving from aerospace

operational needs foreseeable over the next two decades.

Comparison of estimated costs, time, and, above all, the

ability to solve Air Force scientific requirements directed

the selection of the third alternative course of action. To meet this need, a complex multimode human centrifuge has been

constructed at the Aerospace Medical Research Laboratory,

Wright-Patterson Air Force Base, to replace a single mode

simple centrifuge. The new facility called the Dynamic

Environment Simulator (DBS) is capable of providing acceleration, vibration, gas composition, pressure, acoustic and

thermal environments as they may actually occur in an aerospace vehicle during normal flight or during emergency

and escape conditions.

Simulation of a broad variety of aerospace flight mission

environments is now possible using this facility. The man

rating of this facility was carried out in the last six

months of 1969 and resulted in the first manned run in

December of that year. PRIMARY SUBSYSTEMS

The Dynamic Environment Simulator has been called many

588
things. A schematic diagram of the major hardware assembly

is seen in Figure 1. It is essentially a stiff arm human

centrifuge and complex motion device with a multiple

environment capability. The system includes the following

major components or functions: (1) rotation capabilities,

main arm, fork and cab axes; (2) vibration table; (3) power

train; (4) hydraulics and bearing systems; (5) environmental

control; (6) command and controls; (7) standard operating

procedures and operator positions including the medical

monitor's room, machine operating room and pit and (8) data

acquisition systems. The subject is positioned in a 10 foot

spherical chamber or cab located at the end of the main arm. The cab is hydraulic powered and can rotate at 150 rpm. The

outer gimbal or fork which moves in a horizontal plane is

used primarily for vector alignment, however, it is capable

of reaching 30 rpm. The main arm length is 19-1/2 feet from

the center of rotation to the center of the cab axis. The

major structure can reach 56 revolutions per minute or 20G

units at the cab. The main arm of the centrifuge with its

two gimbals weighs about 170 tons. The structure is supported

by externally pressurized "hydrostatic" bearings and rides

on an oil film 0.005 of an inch thick on the thrust bearings.

Similar radial bearings support the main shaft from all sides.

This type of bearing was chosen because of its smooth operating characteristics and minimum friction. A solution of water

and propolene glycol which is not only nonflammable but

harmless to personnel coming into contact with it was chosen

for the hydraulic fluid. The glycol fluid required special

deaerators because of its capacity for retaining air bubbles.

The arm is driven by three electric motors acting

through a common gear system. Together they deliver 330

horsepower. The main drive system is augmented by six flywheels deriving their power from small hydraulic motors. Energy is built up gradually in the flywheels over a period

of 15 minutes or more and then transferred to the centrifuge

by a set of aircraft disc brakes used as clutch. The flywheel drive adds about 15,000 horsepower for quick thrust

during the brief period that it is in action. Although this

capability is in place, it is not in use. The design

calculations predict onset rates of lOG per second for the

main arm using this "hi-drive system". The electric motors maintain the high velocity obtained by the rapid onset of

acceleration from the fl5^heels.

In order to prevent large bearing forces due to unbalanced weight distribution, good main arm balance is

essential. Conventional methods used for balancing centrifuges could not be used since they depend on measuring minute

changes in length of the centrifuge arm under centrifugal

force. The main arm of the Dynamic Environment Simulator is so stiff that its elongation under load cannot be measured. It

589
was, therefore, necessary to develop a special method of

balancing that makes use of small changes in the hydrostatic

bearing pressures as the arm is rotated at low speeds. In spite of the large weight of the arm (over 300,000 lb), the

balancing method, which utilizes a digital computer, is able

to detect an imbalance of less than 100 lb in the capsule.

When immediate medical attention is necessary, the

machine must be brought to a stop in minimum time. Even unforeseen events, such as power failures, must not interfere

with emergency procedures. The centrifuge can be brought to a stop from rotations at its maximum speed of 56 rpm, corresponding to 20G, in 35 seconds, or at the rate of about

1-1/2 rpm per second, even if electrical power used for

dynamic braking is not available. This is achieved by three

emergency brakes that are engaged automatically upon loss of

power, or manually by the operator of the machine. At the same time, the bearings can remain pressurized by means of

fluid stored in accumulators which maintain the oil film

between the bearings and the journal of the main arm so that

rotation could be continued for as long as 70 seconds, or twice the shutdown time, without damage to the machine.

In combination with the motion of the arm, the two outer gimbals make possible complex rotary movements about

the three axes of pitch, roll and yaw. The maximum acceleration force at some portions of the gondola when all three

systems are operated at top speed is 85G. Forces due to gyroscopic precession would probably limit such three mode

rotations but the limitations of the machine have neither

been calculated nor demonstrated in these modes.

One of the most important features of the new centrifuge

is the cab mounted shake table capable of vibrations in all

six degrees of freedom. It has a maximum vertical trans- lation of 12 inches and rotations of +30° and +19° in pitch

and yaw respectively. Smaller displacements are obtained

in yaw and the other two lateral translations. Peak shake

table accelerations will be 15G with a combined payload and

shake table weight of 1500 pounds.

To simulate the internal environment of aerospace

systems, the capsule can be evacuated to an absolute pressure

of about 3 pounds per square inch or approximately 45,000

feet. Over pressure of approximately 1 atmosphere is also

possible. In addition, the equipment is in place to vary the

temperature within the cab between 40 and 120° F and relative

humidity from 10 to 95%. In practice, however, it is likely

that thermal stress will be applied to the subject with the use of a small cab-type enclosure or thermally controlled

clothing. For efficiency in programming experiments, interchangeable gondola or cab capsules or shells are provided,

thus one experiment may be set up in advance while another is

in progress on the DES. A separate signal patchbox is

available so that the entire experiment instrumentation can

590
be set up in the off line gondola and complete test checkout

and nondynamic experimentation performed.

All motions of the Dynamic Environment Simulator are controlled by a digital computer which also monitors the status of the main subsystems and automatically initiates

shutdown in case of malfunction or emergency. Emergency stops can be called for either by the medical monitor or the machine

operator or the subject inside the capsule. The medical

monitor is able to watch the subject by means of two closed

circuit TV cameras and can also gauge the magnitude and

direction of the G forces applied to the subject by means of

a special display.

The standard operating procedures for the function of

this device are built around three major operating areas. The operating areas are further composed of integrated teams.

For example, in the medical monitor's room area, there are

four stations: (1) test director, (2) medical monitor,

(3) instrumentation, and (4) experimenter. Other areas are

the machine operator's room and the pit.

The data acquisition and processing system provides for

the collection, processing and display of experimental data.

Approximately four hundred signal channels are available for

either physiological signals , performance measures or power

transmission. A hybrid data processing system composed

primarily of an Adage Ambilog 200 system serves two functions:

(1) The on line new real time processing of biomedical

signals to be displayed to the medical monitor for medical

monitoring of safety purposes. For example, EKG and

respiration signals can be transformed into appropriate rates,

heart and respiratory rates or rate of change of rate and

displayed on a sequential time based display in front of the

medical monitor. (2) The preprocessing of experimental data

with the preparation of digital tape output of experimental

data. At this point, it is readily seen that the system is

designed with the concept of man rating safety and ease of function. It is necessary to formalize the safety concept

and to provide a succinctly reviewable package describing

the system and insuring that there are no omissions of

safety requirements. This led to development of the man

rating approach adopted for the DES. MM RATING THE SYSTEM

To satisfactorily resolve any unforeseen problems and meet the requirements for man rating a simulations facility,

such as is described in the first part of this report, a man-machine team system philosophy was developed and utilized. A document which includes all subsets of the system was written and provides the backbone of the safety analysis.

The document is intended to collect, analyze and collate all

that material pertinent to man rating the DES. Safety

considerations as applied to the DES operation are broken

into three related considerations: subject safety, operator

591
safety including all industrial/occupational safety and

environmental health considerations and facility safety.

The man rating procedure includes : (1) a detailed review

of the safety considerations designed into all subsystems by

reviewing as-built drawings of the major subsystems; (2) a detailed safety analysis conducted using the "worst possible

case" or "failure mode" method. The pertinent components of

each subsystem are identified as potential failure sources. The nature of the failure of a system or component is

specified and its consequences established. The policies,

methods and procedures to prevent such failure are listed.

Systems studied include: structures, power/electrical,

hydraulics /bearings, electronics, communications /control, data

handling and air conditioning. Selected structural,

electrical, control and operational failure modes are listed.

Safety considerations are also implicit in operating procedures

as well as the physical facility. Typical medical emergencies

are manually inserted to develop reaction capabilities of

the operating team. An ideal communication system for this

facility was developed and implemented. The emergency system

includes two types of planned shutdown, emergency air, power

and fire protection systems . The man rating study provides

logical organization and continuing analysis to achieve

absolute or fail-safe safety standards for the operation of a manned environment simulator. The organization of the

collected analyses provides a compendium of system information

in the form of drawing files, specific safety analysis,

industrial safety surveys, operating and administrative

procedures.

The operation of this facility requires a well

integrated team of approximately 12 people. Specific

functions include test director, machine operator, pit

operator, floor monitor, computer operator, medical

instrumentation, investigator and medical monitor. The

machine analysis currently reviewed is limited to the following configuration of the DES : main arm - rotation up to 35 rpm (approximately 8G) at 0.5-lG per second onset rate; fork - position holding capability for simple centrifuge runs with simple rotation up to 30 rpm with arm and cab locked;

cab - position holding capability for simple centrifuge runs simple rotation up to 90 rpm as a short radius centrifuge;

shells - perform as wind shield only; and restraint - seat design rated to 12G and direction perpendicular to platform

with 500 lb subject payload. Subject restraint is designed

to function in motion at G levels described above.

The heart of the analysis is developed in a tabular

format listing Structures and Functions. The format forces a recognition of a potential system fault and a documentation

of preventive measures. Table 1 shows selected portions of

the analysis.

592
PES RESEARCH PR0GRAI4 The DES is presently programmed to support Air Force

research efforts in two technical areas: biodynamics and

combined environmental stress. A summary of the work units

that the DES is to support in the next five years has been

prepared. Specific support will be given to the development

of the B-1 and F-15 systems in particular.

Although programmed to meet Air Force data requirements

for the design of planned and proposed aircraft systems, the

Dynamic Environment Simulator represents a national research

capability capable of supporting the requirements of many

R&D programs including manned space flight, high speed

surface transportation and other systems or functions where

man is exposed to physical environmental stresses in

combination.

EPILOGUE AND ACKNOI^^LEDGE>IENTS The development of a complex state-of-the-art facility

such as the DES includes personnel as well as technical

considerations. The DES cost 6.5 million dollars and six

years to build. That it exists is tribute to the foresight

and dedication of a large group of scientists and managers.

All of the problems of advanced systems development

technical administration and fiscal were present, and

ultimately resolved. The list of credits is long. Well over

200 people devoted significant amounts of personal effort

and energy over prolonged periods. Of the authors, Mr. Cappel of the Franklin Institute Research Laboratories was

the principle investigator and design engineer, M. McCally,

M.D. , Chief of the Environmental Medicine Division of the

Aerospace Medical Research Laboratory was the Air Force

project engineer and monitor and Captain Rogers, Chief of

the Dynamic Simulations Branch, accepted the system from the contractor and set up the systems operation program including

final system testing, man rating and operation. Alvin S. Hyde, M.D. , Ph.D., 11457 Washington Plaza West, Reston,

Virginia, wrote the initial specifications and began the

program, with Colonel A. S. Swan. Laboratory Commanders,

Colonels A. Karstens, J. Quashnock, R. Yerg and C. Kratochvil

provided the vital administrative support. The Raytheon

Corporation designed the controls and the major subsystems

and now support the maintenance and operations activities.

Subcontractors selected for the construction of major

components and subsystems deserve the highest credit for the

excellence of their workmanship, and include: B. G. Danis

Company, Dayton, Ohio; Digital Equipment Corporation, Dayton,

Ohio; Foremans, Dayton, Ohio; General Electric Company,

Cincinnati, Ohio; Goodrich Company, Troy, Ohio; HelldoerferCastellini Incorporated, Dayton, Ohio; Honey^^ell, Dayton,

Ohio; Hughes Bechtol Incorporated, Dayton, Ohio; Walter M. Litsey Incorporated, Dayton, Ohio; Muth Brothers, Dayton, Ohio;

Mutual Electric Company, Dayton, Ohio; Ohio Valley Painting

593
Company, Dayton, Ohio; Rayt leon Company, Wayland,

Massachusetts; Stevens Company, Fairborn, Ohio; Tibbetts

Plumbing and Heating Company, Dayton, Ohio; Timons, Butt and

Head Incorporated, Dayton, Ohio; Wiley Company, Huntsville,

Alabama and finally, Mrs. Carole Smith who prepared this

report and many others during development of the DBS.

594
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Paper No. 41

USE OF THE BEN FRANKLIN SUBMERSIBLE AS A SPACE STATION SIMULATOR

1 Matthew J. Ferguson

Chester B. May^

REFERENCE: Ferguson, Matthew J. "Use of the Ben Franklin Submersible as a Space Station Simulator, " ASTM/IES/AIAA Space

Simulation Conference, 14-16 September 1970. ABSTRACT: This paper presents the NASA-sponsored effort using

the Ben Franklin submersible as a space station analog during the 30-day drift mission in the Gulf Stream, starting 14 July and ending

14 August 1969. The areas of investigation include: (1) Psychological and physiological measurements during the pre- mission, mission, and post-mission phases were related to observed crew behavior. The results reveal that detailed consideration must be given

early in the design to those aspects which could cause crew annoyance and frustration, and which could be further aggravated by long

confinement. Selection of crew pairs for compatibility of personality will help reduce stress in small closed systems. (2) Habitability in a closed ecosystem was investigated. The objective was

to determine the suitability of BEN FRANKLIN submersible habitability data for providing guidelines for future spacecraft design.

These include provision for privacy, control of temperature and

humidity, and adequate facilities for personal hygiene. (3) Microbiological investigation as a study of the effect of total biological

isolation upon the flora of the crew, environmental and life support

subsystems. The continuing shift and simplification of microbial

flora on the 30-day mission indicates a need for investigation of the problem in association with longer space missions. (4) The maintainability experiment objective was to obtain detailed information on

the frequency, duration, type, and complications of onboard maintenance during the mission.

;Ocean Systems Dept. , Grumman Aerospace Corp. , Bethpage, N.Y.

Advanced Systems Office, NASA, Huntsville, Ala.

599
KEY WORDS: space station simulator, submersible, closed ecological system, psychology, physiology, habitability, microbiology,

microflora shift, maintainability prediction.

INTRODUCTION ^ Background. The 30-day Gulf Stream Drift Mission (GSDM) was

conceived by Dr. Jacques Piccard in 1965 to explore the Gulf Stream

from Florida to Nova Scotia using visual observations, bottom photography, biological surveys, and acoustical surveys. Early in 1967

the Grumman Corporation agreed to undertake the mission and established a program for the design, development, and construction

of the BEN FRANKLIN (Figure 1).

During the design and development phase, the similarities between

the GSDM and space missions became apparent. At the same time

the NASA Office of Manned Space Flight awarded Contract NASW1965 to Grumman to study the feasibility of using undersea facilities as space mission analogs. Final Report OSR-68-6, 11 March 1968,

"Feasibility Study - Use of Submersibles as Space Mission Analogs, " presents the results of the contractual effort. It was concluded that submersible missions would be reasonable analogs of space missions. Both types of missions would be manned by scientific and en- gineering crews motivated by a scientific purpose to work under

operational hazards. Therefore the submersible could provide

space programs with data on crew reactions, the man-machine in- terface, habitability, and the effects of complete biological isolation

during a long mission. In support of these conclusions, NASA

awarded Grumman a contract to study these factors during the GSDM.

The Naval Oceanographic Office (NAVOCEANO) agreed to support

the ocean mission by providing a surface vessel and two BEN

FRANKLIN crew members to perform ocean experiments. The re- mainder of the crew of six consisted of two pilots (including Dr.

Piccard), a relief pilot and oceanographer, and a NASA crew member responsible for the NASA effort.

The GSDM began on 15 July 1969 when the BEN FRANKLIN submerged into the Gulf Stream off West Palm Beach, Florida. It ter- minated 30 days, 11 hours later when the BEN FRANKLIN surfaced

360 miles south of Nova Scotia. The drift covered 1444 n mi at an average depth of 650 ft. Ten excursions were made to depths between 1200 and 1800 ft. The mission was supported by two oceanographic ships, two land bases, and a mobile support van.

600
Submersible Advantages. The Gulf Stream Drift Mission (GSDM)

generated data and information applicable to space missions which

could not be acquired through other forms of ground-based simulators. The mission tasks, the diversified crew, the sealed en- vironment, and complete physical separation from the outside

world, produced unique problems. For example:

• The diversified crew and their interaction with the support

team indicated potential command (organizational) problems.

• Microflora of the crew and within the vehicle tended to simplify and to move toward a microbial imbalance which might

prove harmful over extended periods of time in isolation. • Limitations of communications caused the crew to feel cut off from the world. In addition, a lack of private communications with the surface prevented personal discussions with

families and physicians.

• Complete isolation and separation forced the initial provisioning of all food. Foods were pre-mission tested and accepted,

but under mission conditions were found to be unsatisfactory.

The submersible provides a means for evaluating the operational

effectiveness of crew habitability factors (food, clothing, accommodations), crew skill mix, command structure, and selected spacecraft subsystems checkout and maintenance during a real ocean mission.

To date, manned space operations, habitability provisions, and life support hardware have been tested in ground-based static simulators or chambers. However, as test durations have increased in manned testing, motivational problems have resulted due to a lack

of meaningful work activities. The GSDM showed that the submersible overcomes this problem since a variety of meaningful scientific tasks are performed by the crew in support of the mission. At the same time, the effectiveness of candidate long-duration spacecraft

hardware, operational procedures, and crew interactions can be

evaluated in a closed, stressful environment.

Submersibles and ground-based chambers are complementary.

Chambers are required for thermal, vibration, radiation, and life tests under simulated environmental conditions. The submersible

offers a means for expanding on such tests by providing men and

equipment an analogous mission environment. Figure 2 lists the many similarities between spacecraft and submersibles and compares them with chamber studies to date. GSDM Inputs to Manned Spacecraft Systems Design. The objective

of the NASA contract was to explore the areas of psychology and

601
physiology, habitability, microbiology, and maintainability to obtain

spacecraft design criteria. The GSDM schedule precluded the acquisition and installation of space subsystems. BEN FRANKLIN subsystems and operations were, therefore, the basis for the spacerelated studies, and only incidental space equipment was provided.

The following are a few examples of the t3^es of spacecraft design

guidelines supported by the data from the NASA contract: Crew Accommodations

• Design and locate bunks for conversion to private lounges for

relaxation, reading, and writing.

• Soundproof the crew quarters and locate them as far as possible from operating equipment and work areas.

• Separate work and living/recreation areas. Where limited

volume prevents this, activities must be scheduled to avoid

overlap and interference.

• Provide for private communications with family and friends.

• Provide privacy, recreation, and storage for personal belongmgs.

Food

• Provide home-type food and preparation facilities. • The crew should not be forced to accept a monotonous diet and

disagreeable foods. • Variety and individual preference should be considered.

Clothing

• Daily underwear change is essential for comfort

• Provide for internal environment off- design temperature conditions. • Provide two-piece garments rather than jumpsuits for ease of personal hygiene functions.

Biotechnology

• Develop automated on-line contamination monitoring and provide simple means for decontaminating the vehicle surfaces,

internal atmosphere and water management subsystem.

• Compartmentize the spacecraft to reduce spread of contamination. • Provide negative pressure in hygiene areas to prevent issue of contaminants into living areas.

• Choose materials and designs to minimize microbial nutrients

and breeding grounds.

• Provide microbiological screening of crew to eliminate pathogen carriers.

602
• Monitor individual crew microbial makeup to detect potential

spread of infection. • Provide for microbe incineration, in addition to filters, to assist in atmosphere decontamination.

• Provide safe means for final disposal of microbially contaminated items.

• Provide an on-board microbiological laboratory to facilitate prompt analysis.

Maintainability

• Develop specific maintenance skill levels, experience and

training requirements for crew members on long duration

space missions.

• Develop the facilities required for on-board off-line repair of space equipment.

A study of the command structure during the GSDM was not part of

the NASA contract, but it is evident that such a study would have

provided valuable data for space missions. The BEN FRANKLIN

crew comprised a "mini" crew of scientists and operational personnel similar in composition to spacecraft crew mixes. During the GSDM, operations problems arose related to command structure,

scientific/operations personnel skill mix, and the mission control

team decisions made on the support ship, PRIVATEER. These

problems are analogous to those anticipated in space station operations. An investigation into the causes and resolution of such problems in submersible missions could help provide insight into the handling and prevention of similar situations on future space missions.

Recommendations - During BEN FRANKLIN ocean missions in the

future, the space investigations of the GSDM should be expanded to further develop general space technology and to support the skylab,

the space station, and the space base programs.

Psychology and Physiology

• Crew performance evaluation

• Biomedical instrumentation

• Crew selection test verification

• Work task unit Habitability

• Personal hygiene provisions

• Food management

• Clothing

603
• Recreation provisions

• Noise control

• Personal accommodations

Biotechnology

• 60/90 day mission with crew rotation

• H2O and atmosphere contaminant measurement and control

• Spacecraft and subsystems decontamination

Subsystems Operation

• Water management

• Waste management

• Atmosphere storage/supply

Maintainability (of space equipment)

• Failure prediction techniques

• Scheduled and unscheduled task analysis

• Spares and tools requirements

• Repair techniques/operations

Mission Operations

• Command Structure evaluation

PSYCHOLOGY AND PHYSIOLOGY

Objectives. The objectives of this portion of the study were:

• To relate observed crew behavior to variables (engineering

design, choice and training of crew) which designers and mission planners can use to influence space systems performance.

• To investigate the physiological aspects of the 30- day confine- ment. Approach. Pre-mission data were obtained to establish a personality profile of the crewmen, to establish a physical fitness index, and

to develop a baseline on a motor skills test.

Data were obtained during the mission by means of a daily questionnaire or log that included, in addition to items related to the operation, the mission and the environment, the Cornell Medical Index, a Mood Scale Check List, a Subjective Stress Scale, a Sleep Recall

Questionnaire, a number of tests to evaluate fitness, and daily tests

of proficiency on the Langley Research Center Complex Coordinator.

604
Time- lapse photographs were obtained of most of the vehicle with 3 cameras, and approximately 1 hour of conversation each day, was

recorded on tape.

Upon their return to Bethpage, N. Y. , the crewmembers were given

the following to complete for the debriefing with the psychologists:

• The Group Confinement Inventory (Retrospective)

• The Isolation Symptomatology Questionnaire (Retrospective)

• The Hostility Symptomatology (Retrospective)

• The Primary Affect Scale

Five of the men returned them immediately, and the sixth about a month later.

Results. Although selected only for special skills and desire to participate, the crewmen assigned to work together were reasonably

compatible. Predictions of crew behavior based on pre- mission

psychological tests, clinical interpretation and observation of the crew were proven reasonably accurate in the mission. However, the number of personality tests might be considerably reduced and still provide the same degree of insight.

Predicted annoyance and psychological stress were produced by the austere BEN FRANKLIN characteristics. These included the bunks,

their location, food, the small galley, people noise, odors, lack of privacy, clothing, inaccessibility of equipment, limited personal

hygiene facilities, and environmental control.

As time increased, the men showed a general trend toward withdrawal and an increased need for privacy. This was evidenced in part by the crewmen's tendency to eat more and more meals alone

as the mission progressed. None of the crewmen reported psychosomatic or hypochrondrical symptomatology. Depression was

greatest and the sense of personal well-being lowest at the midpoint of the mission (scores for these factors were extracted by

Naval Medical Research Institute's factor analysis computer program) .

None of the crew suffered serious deterioration in proficiency. However, as judged from measurement of complex coordination, changes

in proficiency in one instance could be related to a mood of depression. Potentially serious problems resulted from failures and misunderstandings in communications with the surface crew. This was

especially true when expected personal news was inexplicably lacking. 605
Events in the GSDM indicate that expected communication of personal news at regular intervals probably is unwise because, if communications are delayed or interrupted, crewmen tend to feel anguish and

concern. On the other hand, a lack of private communications to the surface was a source of annoyance during the mission.

Food provided a topic of conversation and possibly allowed for at least limited sublimation of psychological stress. Other topics

such as daily questionnaires and the interactions with support personnel, accomplished the same result. These are not, however, ap

propriate avenues for the release of tension. Investigation is recom- mended to develop more acceptable techniques to relieve psychosocial tension.

The physiological investigation included analyzing data obtained prior

to, during, and after the mission. These covered a physical fitness index, wrist and forearm strength, recovery pulse, blood pressure, oxygen utilization, and weight. None of the crewmen showed signs of physical deconditioning, although some had a weight loss. All were

declared by the Grumman physicians to be medically fit subsequent

to completion of the drift mission.

Inputs to Space. Detailed consideration must be given early in a design to those aspects which could cause crew annoyance and frustration, and which could be further aggravated by the long duration confinement. Particularly important are:

• Environmental control/life support system

• Privacy areas

• Illumination

• Noise

• Food/preparation/clean-up facilities • Recreation

• Work areas

• Multiple use of spaces

• Personal hygiene

Selection of crew pairs for compatibility of personality characteristics will help reduce psychological stress in small closed systems.

In addition, the mixed crews of future space missions should have a voice in the selection of teammates. A better understanding of the importance of this input is required and further investigations are recommended.

606
Performance rating both psychological (clinical) and operational can

be obtained by self- reporting of the crew if they believe that the in- formation they provide will be held in confidence. The development

of improved techniques is recommended for obtaining self-evaluations by the crew with the intent of reducing the number of questions,

eliciting observations not specifically called for, solicitations of reports about others and for reporting during a mission. An objective measure of skill by a device like the Langley Research Coordinator

appears to be predictive and was accepted by the crew. Further in- vestigation to develop its utility is recommended.

HABITABILITY

Objective. The objective was to determine the suitability of BEN

FRANKLIN habitability data for providing guidelines for future spacecraft design. Factors considered in the study were food, clothing,

control of environmental conditions, hygiene provision, crew equipment items, and crew reactions to these provisions.

Approach. The procedures and records used in the study included:

• Time-lapse cameras located at three places, set to function

every two minutes.

• Environmental measurements (light, noise, temperature,

etc . )

• Counters to measure use of toilet facilities, etc. • Ship's log. • Crewmen's personal logs and questionnaires.

• Debriefing

• Comparison of actual activities with planned activities in the Mission Plan.

Results . Analysis of the camera photos coupled with a study of the

logs established a record of each man's location and activity throughout the mission. From these studies of area utilization and deviations from planned activities, it was determined that generally, half

of the crew followed their plan and half did not. Figure 3 illustrates

this type of analysis on day 1 of the mission, crewman #1 deviating

from the plan and #6 following the plan. Also shown are the planned

and actual hours spent at each location. These deviations are due in part to lack of pre- mission training establishing each man's role and

to changes in work/rest cycles. The significant factor from this habitability standpoint is the overlap of work and recreation activities in a given area.

607
The ten top complants on questionnaires and volunteered complaints are shown in Figure 4. These indicate items which demand man/

machine consideration both in submersible and spacecraft design.

Food complaints stemmed principally from the difficulties in preparation. Cooking was ruled out since it would contaminate the atmosphere. Canned and freeze dried food was provided and was satisfactory from a nutritional and storage viewpoint. However, complaints about food increased with time.

The crew complaints on privacy and free space are presented in Figure 5. A maximum of four complaints were made on Day 15. The complaints decreased to one on Day 22 and started to increase

to Day 29. It is interesting to note that no complaints were volunteered throughout the mission. The principal complaint was that each crew member needed a place other than his bunk.

Although the bunks were oversized, the crew complained that it was

not possible to sit up or bend knees without hitting the pressure hull.

Atmospheric conditions were monitored and recorded throughout the mission. Variations were readily maintained within allowable physiological limits. Atmospheric constituents and trace contaminants were monitored with Drager Tubes and a gas chromatograph. The

Drager tubes identified a continuing rise in CO throughout the mission to a maximum of 40 PPM, identifying the need for greater

capacity in the CO removal apparatus.

On the basis of the habitability study during the drift mission the following guidelines are recommended for future spacecraft design:

• A separate area with soundproofing, adequate lighting and

comfortable chairs, is needed for reading and writing. • Sleeping quarters should be isolated from the noise of the work area.

• Food preparation devices and techniques should be simple.

• Environmental monitoring should be automatic to free the crew for more useful activity, • Clothing and bedding for space stations should be evaluated at off-design conditions, to determine their adequacy.

• Illumination levels should be adequate for the task to be performed.

608
• The crew's use of the vehicle, crew activity, crew time lines, crew living and working areas require detailed consideration

and integration.

MICROBIOLOGY

Objective. The 30-day submergence during the GSDM produced a unique internal environment which, with regard to biological isolation, closely resembles that of future spacecraft. The objective of the microbiological investigation was to study the effects of total biological isolation upon the flora of the crew, environment and life support subsystems.

Approach. For comparison purposes, pre-mission, mission, and

post-mission sampling were scheduled. To make the bacterial

counts, a variety of sampling media and devices were taken aboard

including:

• Andersen samplers, (atmosphere)

• Swabs and agar plates (human and surface environmental

sampling)

• Field-tjTDC water monitoring unit (field monitor kit). Food was tested for bacterial count pre- and post-mission„ Waste,

garments and linen were stored on board until the completion of the mission and were then returned to the biotechnology lab for bacterial

coimts and identification of the types of bacteria present. Approximately 15, 000 separate culturing steps were required to identify to Genus the 2230 isolates obtained.

Results . There appears to have been a general simplification and

shift towards gram -negative organisms, (Figures 6 and 7), particularly Pseudomonas and Aerobactor. This has been postulated as an

effect of long term isolation, but could have been biased by on-board

sample incubation at low temperature, followed by long-term storage, or by the use of antimicrobial soap.

With the exception of one crew member (who had a historj^ of Staph

infections), there were only transitory occurrences of Staph Aureus

in the crew and the environment (Figure 8c) Beta Hemolytic Streptococci were isolated from 5 of the 6 crew members (Figure 9) = These 5 also developed upper respiratory infections during the early

mission phase.

Contamination levels of the foods were well within acceptable limits

both pre and post missionc

609
Garments and linen had generally low levels of contamination when

cultured after use and onboard storage. It would appear that anti- microbial treatment of garments and linen were effective in suppressing bacterial proliferation and odor.

Iodine treatment was planned for control of the cold water system

microbial contamination. However, the crew objected to the iodine

taste in the water and the reconstituted food. In addition, the iodine treatment was difficult to perform and was not implemented. This

contributed to cold water contamination.

Pseudomonas, which had been a problem during much of the pre-mission attempts at cleaning the system reappeared as a consistent contaminant. Later in the mission a variety of human associated organisms were recovered, including the fecal E coli. Several filters (head, galley, shower- sink) were found to be contaminated with

Pseudomonas when cultured post-mission. Serious discrepancies

existed between the on-board readings and base laboratory analysis

of the same sets of samples. (Many samples read as "sterile" on- board were discovered by base lab analysis to be contaminated.

This was attributed to the limited facilities on board the BEN

FRANKLIN:

• Low-temperature incubation

• Poor lighting ; • Lack of experience

Provisions were made for dispensing germicides, replacing odor

control canisters, and adding antimicrobial agents to the waste stor- age tanks o During the mission, macerator electrical problems prevented proper mixing of germicide with waste. This resulted in noticeable odor levels.

The effectiveness of odor control in the waste management contamination control system was not evaluated during the mission. At postmission sampling, all waste tanks were found to be contaminated with

between 106 to 107 micro organisms/ml, most of which were of in- testinal origin.

As the mission progressed, the environmental flora reflected more

and more of those organisms found on the men. The pattern of a

shift towards gram negative organisms with respect to number of genera isolated was also similar to that of the human flora. Cleaning

appeared generally effective in attaining a transitory reduction of

total microbial level, although a general rise in contamination persisted as the mission progressed. The initial clean-up appeared to

610
lower the microbial load considerably, because on Day 2, all counts were low. The dips in contamination levels can be related to cleaning

procedures with transient drops in microbial counts on the walls and

floors noted at the general cleanups. However, after cleaning, there was a rapid rise of the contamination level on the floors and walls. The contamination of the table tops increased with time at a slower

rate because of the daily washings.

Even with the limitations imposed by schedules and funding, the microbiology study produced guidelines for spacecraft design. Many

unknowns exist in the area of space microbial technology, and this

test should be considered only as a basis for much additional work,

i. e.

• The continuing shift and simplification of microbial flora on the 30- day mission indicates a need for investigation of the problem

association with longer space missions.

• The personal hygiene areas, and humid areas in general, are

fertile microbe breeding grounds and require microbial control. • The water and waste management systems are particularly

fertile areas and require suitable contamination monitoring and

simple decontamination provisions.

• The use of anti- microbials offers temporary advantages but the overall effect may be to create an undesirable microbial imbalance. Additional work is required in this area.

• The stored hot water system was effective in controlling contamination and should be a candidate for spacecraft. It could

eliminate the need for biocides such as iodine or chlorine which

are disagreeable in food and drink. • Further testing is required on the effects of crew rotation, i. e.

putting a new crew member into an altered environment.

MAINTAINABILITY

Objective. The objective of the maintainability experiment was to obtain detailed information on the frequency, duration, type and

complications of the onboard maintenance performed during the Gulf

Stream Drift Missiono This would permit evaluation of existing

maintainability techniques for application to space vehicle missions.

Approach. The maintainability study covered:

• Analysis of the systems and equipment in the BEN FRANKLIN

to establish spares, tools, test equipment, and estimated work

loads.

611
• Preparation of maintenance procedure and data sheets, crew

training, and dock side maintenance time trials, etc. • Maintenance recording during the mission

• Reducing and evaluating data

This maintainability experiment did not encompass all of the equipment aboard the vessel. Systems and equipment were selected on the basis of criticality and available information on which analysis could

be performed. This became the "controlled" portion of the study.

Actual mission data was collected for all maintenance performed.

Logistics preparation consisted of crew training and of maintenance

procedures, troubleshooting information, checklists, computation

charts, spares, tools, and test equipment for the equipment in the controlled maintenance portion of the experiment. This was successful in satisfying the maintainability objectives. Crew comments in- dicated that this preparation was precisely what they needed for ap- proaching the mission with confidence.

Results. The crew performed 1354 individual maintenance tasks, an average of 45 per day. Figure 10 shows the percent of total available

manpower expended on maintenance during the mission. The maintenance workload actually required from 12 to 31% of the crew's total available duty time each day. On the average maintenance occupied

the equivalent of one man full time throughout the mission.

Scheduled maintenance accounted for 1312 of the 1354 maintenance

tasks (Figure 11). Successful completion of the remaining 42 unscheduled repair actions, however, assured mission success. Two crew members performed 58% of all the maintenance work, but more significantly they accomplished 96% of the unscheduled repair

actions primarily because of their highly skilled and maintenanceoriented background. Figure 12 illustrates the maintenance workload

assumed by the one crew member who was the prime mover in all of the unscheduled repairs. His skill contributed to the mission's suc- cess and attest to the need for this type of crew member on all such

missions.

The statistical analysis of mission data indicated that maintenance

tasks time predictions by Method II of MIL Handbook 472 were rea- sonably effective in determining task times. Figure 13 shows a comparison of the results when regression analysis was applied to these

predictions in mission action dock-side time trials, and an aircraft

program as a control case. In view of the results, we concluded that

612
Method n was considerably better than Method III since it is more

closely associated with actual hardware configuration and limitations.

Mission data analysis indicated that there was no discernible time

differential for the performance of maintenance in the stress of this mission environment versus the relatively unstressed environment

of the dock-side time trials. All crew members however, did admit

to feeling the effects of stress at various points during the mission.

Since the GSDM maintainability studies were performed on submersible systems, no recommendations can be made concerning specific

space systems. However, the study showed that the submersible

could be used productively for testing maintainability techniques on space hardware under analogous mission conditions.

It was revealing to find that the equivalent of one man out of the six was required to perform maintenance tasks. Space stations will require many more highly complex subsystems, and mission duration

will be measured in months and years rather than days. Hence, it appears that sophisticated analysis, training, and automatic failure

detection methods will be required.

Since sophisticated training of many crewmen is expensive, it is ap- parent that means must be developed to reduce future space crew

training. The submersible offers a facility to develop and evaluate

alternative crew training procedures, with the aim of reducing crew

training requirements and associated special skill needs.,

613
GENERAL SPECIFICATION DISPLACEMENT LENGTH BEAM (OVER MOTOR GUARDS)

HEIGHT OPERATIONAL DEPTH COLLAPSE DEPTH SUBMERGED SPEED (MAXIMUM)

LIFE SUPPORT PAYLOAD

TOTAL POWER

VIEWPORTS

130 TONS

48 FEET, 9 INCHES

21 FEET, 6 INCHES

20 FEET

2000 FEET

4000 FEET

4 KNOTS

6 MEN FOR 6 WEEKS

5 TONS

756 KWH

29 FIGURE 1 BEN FRANKLIN CHARACTERISTICS SPACECRAFT SUBMERSIBLE CHAMBER

CONFINEMENT X X X SOCIAL ISOLATION X X X DEPRIVATION X X X CLOSE QUARTERS X X X MEANINGFUL MISSION X X SUSTAINED MOTIVATION X X

HOSTILE ENVIRONMENT X X OPERATIONAL STRESS X X REMOTE OPERATIONS X X ABORT DIFFICULTY X X REQUIRE REAL NAVIGATION X X

SCIENTIFIC CREW X X DATA TRANSMISSION DIFFICULTIES X X ON-BOARD MAINTENANCE PROVISIONS X X COMPLETE BIOLOGICAL ISOLATION X X COMMAND STRUCTURE X X FIGURE 2 SIMILARITIES -SPACECRAFT/SUBMERSIBLE/CHAMBER

614
AFT SCIENTIFIC AREA

PRIVATE AREA

PILOT AREA

GALLEY AREA WARD RM/ OBSERVATION AREA

CREW MEMBER 1 HOURS

PLANNED ACTUAL

9 M 2 0

3 0

10 15 PLANNED LOCATION ACTUAL LOCATION -A A10 12 14 16 18 20 22 2400 DAY 1, HOUR

AFT SCIENTIFIC AREA

PRIVATE AREA

PILOT AREA

GALLEY AREA WARD RM/ OBSERVATION AREA

CREW MEMBER 6 HOURS

PLANNED ACTUAL

7 10 1

k

'

2 1 rr \. .. 1 2 5 1? 8 PLANNED LOCATION -•- -•- ACTUAL LOCATION -A A—A0 2 4 6 8 10 12 14 16 18 20 22 2400 DAY 1,H0UR FIGURE 3 ACTUAL VERSUS PLANNED CREW MEMBER TIME LINES, DAY 1

615
COMPLAINTS REQUESTED VOLUNTEERED

1. SURFACE COMMUNICATION

2. FOOD

3. FORWARD SEATS x

4. CLOTHING

5. FORWARD TABLE

6. BUNKS

7. TEMPERATURE CONTROL

8. ACCESSIBILITY

9. HOT WATER

10. GALLEY SPACE

11. TIME LINE

12. NOISE

0 6 12 18 24 30 36 42 48

y////////////m

w//////m

mmm

FIGURE 4 MAJOR HABITABILITY COMPLAINTS

o e zt 5

2O 3

NOTE: NO VOLUNTEERED COMPLAINTS

REQUESTED COMMENTS

15 MISSION DAYS

22 24 29

FIGURE 5 PRIVACY AND FREE SPACE COMPLAINTS

616
FIGURE 6 TREND OF TOTAL BODY SIMPLIFICATION

FIGURE? TREND OF TOTAL BODY SHIFT 617
m. J2Z2 EZZl YTTh

TABLE TOP

NOTE: DAILY WASHINGS

jzza WA EZZa YZA WA ^^vm

8 11 14 17 21 24 27 +8 WALLS

1 1 -5 2

:i^2

A

11 14 17 21 24 27 +S A A

FLOORSi

i

I

SYMBOL OPERATION CLEANING

11 14 17 A

MISSION DAYS

21 24 27 £^AAAA A A

+8

FIGURE 8 ENVIRONMENTAL CONTAMINATION

618
TIME DAYS

FIGURE 9 POTENTIAL PATHOGEN INCIDENCE

619
15 20 MISSION DAY

25 30

FIGURE 11 MAINTENANCE ACTION SUMMARY

10.0-, 9.0 8.0H

ccI

Z 7.0 <

uj 6.0 u

< 5.0- Z

LUH 4.0 1 3.0

2.0

1.0-0

H 4.0 -^J

.:

UNSCHEDULED MAINTENANCE

SCHEDULED MAINTENANCE'

MISSION DAY

FIGURE 12 CREWMEMBER NO. 4 MAINTENANCE WORKLOAD BY DAY

620
ESTIMATED TIME, MIN FIGURE 13 COMPARISON OF PREDICTION CHARACTERISTICS

621

Paper No. 42

DEVELOPMENT OF m OPEN CIRCUIT CRYOGENIC LIFE SUPPORT SYSTEM

FOR USE IN NEUTRAL BUOYANCY SPACE SIMULATION.

1 2 C. M. McClure and C. W. Williams

REFERENCE: McClure, C. M. and Williams, C. W., "Development

of an Open Circuit Cryogenic Life Support System for use in

Neutral Buoyancy Space Simulation," ASTM/IES/AIAA SPACE

SIMULATION CONFERENCE, 14-16 September 1970.

ABSTRACT: Marshall Space Flight Center has a need for a self

contained underwater life support system which can supply a breathing mixture to the neutral buoyancy test subject,

without the need for a surface tethered supply line, A prototype system has been developed which utilizes a mixture

of liquid oxygen and liquid nitrogen contained in a single

supply tank. Development and testing has shown that this

system will supply the quantity and quality of gas required

under completely controlled and predictable conditions to depths of at least 60 feet.

KEY WORDS: Neutral buoyancy, space simulation, cryogenics,

liquid oxygen, liquid nitrogen, liquid air.

Senior Engineering Specialist, Mayes International

Corporation, Missile and Space Support Division, lluntsville

Alabama

Scientist, Hayes International Corporation Missile and

Space Support Division, Huntsville, Alabama

623
INTRODUCTION

Present neutral buoyancy space simulation techniques

still model the tethered astronaut type of mission. Recent lunar surface operations and many future space missions under

zero and sub-gravity conditions show the necessity for space

simulation techniques which can model, with high fidelity,

the umbilical -free astronaut missions. The present umbilical

is used as the suit gas sunply line. Therefore, a major type

of equinment which is required to meet umbilical-free

operations is a self-contained underwater breathing unit

(SCUBA)

The Manufacturing Engineering Laboratory of the Marshall

Space Flight Center has just completed the development of such

a Drototype system.

In order to develop the simplest system possible, it was decided to develop this backpack to be compatible with present

suit systems in use in neutral buoyancy testing. This requirement results is retaining both the open circuit suit gas system and present suit through-flow gas rates. This

combination could not be met without excessive weight and

volume in a backpack using high pressure gas storage.

Cryogenic liquid gas storage at subcritical pressure and high

compressive potentials offered high storage efficiency

without excessive weight or volume penalties. Further

operational simplification was obtained by dictating that the

diluent gas constituent be contained within a single liquid

cryogen mixture. Neutral buoyancy depths easily allow choice

of nitrogen as the diluent gas. Therefore the original

design objective was for a single liquid mixture of liquid

oxygen (LOX) and liquid nitrogen (LN2) to be stored in a cryogenic system to feed the open circuit neutral buoyancy

suit

DESIGN REQUIREMENTS

Several design boundary conditions were established prior

to starting this project; for principal importance were: (1)

safety of operation, and (2) simplicity. Since the proposed

apparatus must eventually be man-rated, the first condition

follows as a natural consequence. The second requirement is merely a logical step towards achieving an operational state

as soon as possible.

624
From these two basic conditions were develoned certain

specific design requirements:

1. The prototype was to be a breathing system of the

"open circuit" type. "Open circuit" refers to the

fact that the supply gas is Jumped overboard after

only one pass through the pressure suit. This system

is currently used in Marshall Space Flight Center

(MSFC) neutral buoyancy operations with an umbilical

connection supplying a continuous flow of air through

the pressure suit. Physiologically, this is the

safest and simplest system; however, in addition to compromising simulation fidelity it is highly

wasteful of the gases being supplied.

2. One hour duration was established as a target

operational time for a single charging of the

cryogenic tank.

3. Present suit through-flow rates of 5.0 ACFM (actual

cubic feet per minute) were to be maintained at all

depths. This means that as depth is increased, flow

rate must also be increased to maintain 5.0 ACFM

through the suit. If 45 ft. is established as the

maximum operational depth, then the maximum standard

cubic feed per minute (SCFM) rate on the bottom will

be 2.36 times the surface rate, or 11.8 SCFM, In addition, it was felt that flow rates should be

adjustable by the diver. From these considerations,

a maximum flow capability of 16 to 18 SCFM was finally

selected,

4. A cryogenic storage system was to be employed using a mixture of liquid oxygen (LOX) and liquid nitrogen

(LN2) in a single dewar. Cryogenic storage is necessary in order to supply the relatively large

quantity of gas required.

5. Human safety and comfort parameters (breathing gas temperature, oxygen partial pressure, suit cooling,

etc.) were to be maintained within the following

limits:

(a) Air temperature should be maintained at a comfortable level between 68°F and 80°F,

(b) Oxygen partial pressure (PO2) should be held

within acceptable physiological limits. The

minimum PO2 of 0.20 atmosphere (152 mm Hg) is

625
the lowest allowable. The maximum PO2 is less

clearly defined; however, the Marine Technology

Society has set a PO2 of 1.33 atmospheres (1010.8 mm Hg) as the upper limit for continuous

exposure. Other equally well qualified sources

have set this figure at 1.25 atmospheres (950 mm

Hg) , In general, the spaceflight environmental

control system (ECS) limits for maximun PO2 have

been set lower than either of these figures.

The problem of maximum PO2 is simpler in cabins

sealed for spaceflight. These cabins have a total pressure which is 1,0 atmosphere or less.

The hyperbaric Pn2 problem found in submerged

operations is more difficult than for space

cabins. With preset gas ratios, such as with air

or any single liquid cryogen, the minimum

(hypoxic) surface limit selected also presents

the depth PO2 curve. Therefore, under conditions

of submergence v\7ith constant ratio gas mixtures

a realistic maximum may well be those chosen by

the quoted diving oriented-sources

As a design goal, the PO2 range from 0.20

atmosphere to 1.0 atmosphere was selected for

this work.

THEORETICAL CONSIDERATIONS

Initial studies showed a major problem to be the positive

control and delivery of safe oxygen partial pressures.

Several important theoretical aspects relating to the overall

problem are outlined in the following sections.

A, Oxygen Partial Pressure (PO2) Control

In a vessel containing a cryogenic mixture of liquid

oxygen and liquid nitrogen, the partial pressures of the

individual gasses over the liquid components have been shown to obey Raoult's Law. This law, which applied for all possible

concentrations, states that at any given constant temperature

the partial pressures of the components of a mixture are equal

to the mole fraction of the component multiplied by its vapor

pressure in the pure state at the temperature of the liquid.

The liquid boiling and vapor condensing curve for oxygen

and nitrogen mixtures is shown in Figure 1 and indicates that

at 1 atmosphere pure nitrogen is saturated at about -320°F

(-195°C) and pure oxygen is saturated at •.29S°F (-183°C) , In

626
a binary mixture the simple terms "equilibrium or saturation

temperature" do not exactly apply. In such a mixture, addition

of heat and resulting liquid vaporization will shift both the

mixture temperature and the e^as and liquid phase composition.

However, standard engineering terms are used throup,hout this

section in reference to binary mixtures.

As an example consider a binary liquid mixture of 22% oxygen and 78% nitrogen contained as shown in Fijjure 1. The

saturation temperature is about -317''F (-194°C) . Under these

conditions, this specific binary liquid mixture has in equalibrium with it a ^as phase which contains only 8% oxygen which

is well into the dangerous anoxic level. As this binary

mixture continues to absorb heat, its sensible heat gain and

subsequent temperature rise cause a disproportionate change in

nitrogen and oxygen compositions. More of the liquid nitrogen

boils off, resulting in a constant oxygen enrichment of the

liquid phase, until the liquid is finally 100-6 oxygen.

As can be seen from the example above, any attempt to breathe the gas phase over the liquid mixture would be

dangerous. From the analysis, it would seem possible to produce a gas phase containing 22% oxygen by starting out with

a 50-50 liquid mixture. However, as useage continued, the

partial pressure of oxygen would soon increase to a dangerous

level especially in diving conditions. Thus, the situation

in which a breathing gas phase is derived from a "pool" of a binary cryogen liquid mixture is undesirable. Such a situation will be inherently unstable and unpredictable with

respect to the partial pressures of the gases being delivered.

Because the oxygen nartial pressure is life critical in this

developmentj the instability which derives from "pool boiling"

must be completely avoided.

If a system is considered whereby a volume of liquid is removed unchanged from one container and then totally

vaporized and mixed in a separate container, the percentages of

individual gaseous components will of course be identical to

the percentages of individual components in the liquid.

Continuous total vaporization and mixing of liquid unchanged

from the tank composition can be accomplished by withdrawing

liquid through an orifice small enough so that a constant

flow of liquid in one direction is obtained followed by total

vaporization and turbulent mixing in a sufficiently long,

small diameter, heat exchanger.

627
B. FLOW CONTROL

Since pressurization and venting of present day

spacesuits require a flow rate of approximately 5.0 ACFM, a flow control mechanism is needed to regulate flow according

to changes in diving depth. To maintain 5.0 ACFM through the

suit will require a gaseous flow rate which can be increased

as depth increases. At a water depth of 40 feet (MSFC Neutral

Buoyancy Tank depth), a flow rate of 10.9 SCFM is needed to yield a 5.0 ACFM flow rate through the suit.

Regulating the a^mount of liquid cryogen which can enter

the heat exchanger can be used to regulate the output gas rate. The heat exchanger is allowed to operate at low pressure and

only the amount of liquid required to give a selected flow rate enters the heat exchanger. The rate this liquid enters can be

controlled by controlling the orifice and dewar pressure.

Since a small volume of liquid will yield a large volume of ga^

optimum flow control and PO2 control can be obtained by using

a small orifice, variable flow, control valve which meters

liquid only. Therefore, the flow control valve can satisfy

the flow control requirement discussed above and also the

oxygen control requirement discussed in Section A.

C. TEMPERATURE CONTROL

The established comfort level for neutral buoyancy

pressure suited subjects has been found to vary between 68°F

and 80° F. The heat exchanger on Prototype II is overdesigned

with respect to its heat exchange capacity so that extremely

large flow rates can be accommodated without any appreciable

drop in output temperature below ambient water temperature.

Since the Prototype II unit was designed for undenvater

purposes, a heat exchanger suitable for an air environment was

not considered.

D. STORAGE CONTAINER

The best available insulation would be a vacuum walled

container possibly including multilayered "superinsulation"

This type insulation would restrict heat loss to an insignificant value during the required one hour run time.

The only significant operational contribution of vacuum insulation would be its extended "holding time" (16 to 20 hours

might be expected) , HoV);ever, vacuum based insulation

techniques have three distinct disadvantages: (1) cost,

(2) fragility, and (3) heavy walls required as over pressure

(depth) increases.

628
Preliminary calculations on other methods of insulation

indicated that polyurethane foamed-in-place insulation would

be inexpensive, lighter in weight, and more rugged than vacuum

insulation. The only disadvantage of foamed polyurethane

appeared to be a reduction of the dewar holding time to about

three to five hours; this was considered adequate. Further,

since neutral buoyancy simulation is a centralized operation,

the combination of a standard high vacuum ambient pressure

storage dewar dIus the foam insulated diving dewar would allow

a cryogen holding time of several days plus sufficient

insulation for mission performance.

FINAL PROTOTYPE DESIGN

The final prototype design, known as Prototype IIA, is shown in Figure 2 and Figure 3. The unit can be loaded m a few minutes using a premix of the desired ratio of liquid

oxygen and liquid nitrogen. For operation the vent valve is closed and the buildup valve is opened. Cryogenic tank pressure is built up and controlled by the build up circuit. The

diver simply sets the flow control valve and plugs in to the

outlet to obtained controlled delivery.

629
EXPERIMENTAL EVALUATION

A. TEST EQUIPMENT

In order to measure pertinent characteristics of the gas being discharged form the life support system an instrumented

"analysis bar" (see Figure 4) was used in place of the pressure suited test subject. This analysis bar contained sensors

for measuring the following parameters: (1) oxygen partial

pressure, (2) system absolute pressure (3) actual flow rate

output, and (4) output gas temperature.

Analog signals representative of these parameters were

fed into a recorder so that all values could be recorded

simultaneously on chart paper for future reference.

B. TEST TECHNIQUES

Preliminary testing of the unit was accomplished under

standard atmospheric (ambient) conditions substituting liquid

nitrogen for LOX/LN2 (LON) mixtures. This phase of testing

not only afforded the opportunity for gaining operational

familiarity, but also permitted reliability testing of key

components, valves and heat exchanger. Changes in the

location of these components and changes in their operating

conditions were made as dictated by the test results.

After establishing the operating reliability of the unit, tests were conducted using mixtures of LOX and LN2 as the

filling liquid. These tests, consuming over 2000 lbs, of

cryogen, were conducted to establish PO2 control and to determine any changes in PO2 which may occur during filling, A performance chart for a typical run using LON is shown in

Figure 5. Strip chart data for another run are given in the

Appendix,

In addition to the standard 1-atmosphere tests, testing

was also conducted at smiulated diving depths up to 100 feet

in the single chamber recompression facility located at MSFC. These tests were conducted to establish the effects of

increased ambient pressure on PO2, flow rates, output

temperature, and the structural integrity of the unit. A performance chart for a typical simulated dive to 80 feet is shown in Figure 6. Strip chart data recordings for a dive

to 50 ft. are given in the Appendix,

630
C. TEST RESULTS

Actual testing of the Prototype unit yielded the

following results:

1. The basic cryogenic-gas cycle used by the Prototype

II Unit provides reliable control of all parameters.

2. The special valve components of this cycle (pressure

closing valve, relief valve and micrometer flow control valve)

perform adequately and reliably under normal and special

conditions including being submerged to a depth of 100 ft. The build up valve provides for complete internal system

nressurization within 6 minutes as shown by Figure 7. The

build UP circuit will hold the desired dewar nressure within

2% under normal flow conditions and within 10% of preset

pressure under continuous maximum flow conditions. The

pressure closing valve operates automatically to compensate

for denth overpressure and will maintain pre-set dewar

pressures over ambient pressure at depths up to 100 feet.

Figure 8 shows the performance of the micrometer flow

control valve. This valve is linear and smooth under all

conditions of cryogen flow required and can provide the flow

control and orifice necessary for PO2 control. Flow rates

from 1 through 18.5 SCFM are possible at 80 or 125 psi dewar

Dressure.

Flow in SCFM from the Prototype II unit is dependent only

on dewar ambient pressure and flow control valve setting.

Testing showed that the pressure closing valve accurately

maintained the preset differential pressure above ambient

pressure. Consequently, flowrates (in SCFM) were found to be

maintained at constant values dependent only on the flow

control valve setting. Figure 9 shows for the Prototype II unit the relation between flowrates (in ACFM) and water depth

for a given setting of the flow control valve. However,

constancy of flow rate in SCFM is in basic contradiction to the

requirements of a pressure-suited neutral buoyancy test subject; his situation demands that flowrate through the suit

in ACFM be constant and that SCFM be adjusted with resnect to depth to accommodate this requirement. Thus an important

design feature of the Prototype II unit is a dive adjustable

flowrate control which provides more than adequate gaseous

flow levels under all neutral buovancy depth-demand conditions.

3. The Prototype II unit will convert and deliver the

oxygen content with high precision control as observed during

631
all LON testing (which included more than 2000 pounds of

various LON mixtures)

Total PO2 control includes accuracy of LON mixing,

changes of LON ratio during storage, changes of LON ratio

during dewar transfers (loadinrr), and any change which might occur during oneration. It is interesting to note that during

the entire test program these cumulative factors never caused

a PO2 variation in excess of •»• 5%. If one considers a given

run cycle only, eliminating the inaccuracies of mixing,

storage and loading, PO2 variation is reduced to less than

± 1/2% (which approaches the accuracy limit of the PO2 sensor).

Figure 10 represents total PO2 variation for typical

L02oNgo and LO30N70 mixtures for surface conditions.

The PO2 depth control was also in excellent agreement

with theoretical and precomputed values. In a constant ratio

mixture such as gaseous air or liquid air (LO20N80) any other specific liquid oxygen/nitrogen (LON) ratio mixture,

hyperbaric operation will of course show PO2 changes. The

oxygen constituent must always contribute its relative

fraction of the total hyperbaric pressure. Therefore as the

ambient pressure is elevated on a constant ratio mixture the

oxygen partial pressure must also be proportionately elevated.

Figure 11 shows a typical depth vs PO2 performace curve

for the Prototype II Unit. Two typical LON ratios are shown.

The actual PO2 denth curves are as linear and flat as the PO2 curves obtained from an air dive (deviations are less than

0.1%). The PO2 agreement with respect to precomputed values

is also very good. The actual PO2 curve in Figure 11 in all cases showed a slightly lower slope than the precomputed

values. This type deviation is typical of all air dives and

is felt to be due to a slight lag in the electrochemical PO2 sensor from which all the PO2 values were obtained.

Although PO2 will vary with depth, this variance is in

accordance with Dalton's Law and is exactly the same as for an air dive. The maximum PO2 from any LON mixture used at any depth can be predicted and the Prototype II will deliver that

value. For instance, in a LO33N57 mix, an oxygen partial

pressure of 0.33 atmosphere (4,85 psia or 251 mm Hg) will be

found at the surface, whereas this same mixture will show a PO2 of 0,65 atmosphere at a depth of 40 feet.

4. Testing has shown that the Prototype II Unit will

deliver positive, safe output gas temperatures under all

submerged conditions down to depths of at least 60 feet. The

632
principle of an overdesigned heat exchanger effectively locks

the output gas temperature so that it is always within 5°F of

the water temperature in which the unit is immersed. Figures

5 and 6 show the output gas temperature and the water temperature vs. time and depth respectively. These curves show a slight decline in both the output gas temperature and water

bath temperature. This is due only to the limited volume of water surrounding the unit and would not occur if the unit

were operated in a large volume of water such as the neutral

buoyancy tank.

The unit can be operated at 5.0 SCFM at surface ambient

conditions (heat exchanger in air) only 5 minutes before the

output gas temperature begins to drop. No attempt was made to

size the heat exchanger or to provide auxiliary heat control

for such ambient air operations.

The heat exchanger operated as expected under all

conditions down to depths of 60 ft. and showed sensible

cooling in the first few feet of exchanger length. This

cooling did not produce ice at depths less than 60 feet. Ice was produced on the first few coils only when maximum flow

rates were demanded at depths below 60 feet. This ice

formation is due to the increased dwell time of the gas inside

the heat exchanger as gas density is increased. Ice formation

was never found to include a supercooled surface. Bare skin

contact with an immersed and icing heat exchanger showed no tendency for skin adherence.

Thus, in actual testing, the heat exchanger design

described here proved to be completely free of potential

physiological hazards associated with low temperatures, cold

gas or supercooled surfaces, at least for the pressures and

flow rates of interest.

CONCLUSIONS

The MSFC Prototype II Cryogenic Life Support System will

provide adequate gas for present neutral buoyancy operations.

Flow is diver adjustable for comfort. Positive oxygen partial

pressure control is inherent within the design and the

performance of the unit with respect to oxygen partial

pressures is exactly analagous to an air dive. The unit will

deliver exactly that ratio of gas contained in the liquid

phase in the tank. Selection of oxygen/nitrogen ratios other

than air may be desirable and are easily possible with this

system. Positive output gas temperature control is also

inherent within the Prototype II design. During submerged

operation, the output gas temperature is fixed directly by the

633
water temperature. Even under the highest flow conditions the

gas temDerature will not be more than a few degrees below

water temperature.

This unit is beinj^ refined to include more efficient

layout and packaf^ing and to include the all-attitude dewar

solution. It will then be nut through man rating qualification

tests for use within the MSFC neutral buoyancy facility.

Credits

The authors wish to thank Mr, W. D, Hand of the \layes

International Corporation and Mr. J D Bennight of Marshall

Space Flight Center for their continued support and aid in

accomplishing this project.
Figure 1. EQUILIBRIUM DIAGRAM FOR LIQUID OXYGEN-LIQUID NITROGEN MIXTURES

635
Figure 2. SCHEMATIC OF NBFC PROTOTYPE II UNDERWATER LIFE SUPPORT UNIT.

636
Figure 3. Prototype IIA Cryogenic Underwater

Life Support System

637

639
640
641
i

O \0 <N 00 ^

(N ^

642
643
LO «^ <N o o o o o

(WIV) ^Od

644
645
APPEITDTX

'

'

, \ .... :Ji ^JiiJ:::!, Ir . -uJ^iii Ihi iii! lili l , nii Isi lii 0 Flow (ACFM) 20 0 PO2 (Atm) 1.00 -150 Temp. (*F) Strip Chart Data For MSFC Prototype

II Life Support Unit. (Surface Conditions)

646
0 Flow (ACFM) 20 0 PO2 (ATM) 1 .00 -150 Temp (°F) +100 0 Press. (PSIA) 50 Strip Chart Data For MSFC Prototype Underwater Life Support Unit

(Dive Conditions, 25 Sept. 1969, Run #1, 24 Sept. LON Mix, Initial Water Temp 87° F)

647
0 Flow (ACFM) 20 0 PO2 (ATM) 1.00 -150 Temp ("^F)*100 0 Press ( PS I A) ' 50

Run n 24 Sept. 1969 Continued

648
|0

0 Flow (ACF^O 20 0 (ATM) 1.00 -150 Temp (°F) +100 0 Press (PSIA) 50

Run #1, 24 Sept. 1969 (Continued from previous page).

MSFC-RSA. AM

649

Paper No. 43

ADVANCED DIVING TECHNIQUES APPLICABLE TO EXTENDED

MISSION NEUTRAL BUOYANCY SPACE SIMULATION

Dr. Joseph B. Maclnnis' and CM. McClure'^

REFERENCE: Maclnnis, J.B., and McCiure, CM., "Advanced

Diving Techniques Applicable to Extended Mission Neutral Buoyancy

Space Simulation, "ASTAyiES/AIAA SPACE SIMULATION CONFERENCE,

14-16 September 1970

ABSTRACT: Diving methodology is reviewed. Nonsaturated diving is recognized as the basis for almost all neutral buoyancy testing to date.

Saturated diving methodology is reviewed and its possible special

adaptation for extended duration mission neutral buoyancy simulation

is proposed. The advantages of using the new closed circuit rebreather SCUBA gear is also reviewed and proposed as an aid to extended mission simulation

KEYWORDS: Neutral buoyancy, space, simulation, hyperbaric,

diving

INTRODUCTION

A requirement exists for high quality simulation of Space Flight.

Simulation methods used to date have included gravity balance by

underwater immersion, or "neutral buoyancy" methods.

Medical Director, Ocean Systems, Inc., Washington, D.C

^Director, Aquaspace (Undersea Equipment and Technology),

Huntsville, Alabama

651
Neutral buoyancy simulation to date has been basically a specialized

industrial/scientific diving operation. It has been dependent upon and

based on well-known and time-tested diving methodology. The depth to date has been relatively shallow, from 15 to 50 feet. The immersion

times involved have also been relatively short and are usually less than from two to four hours. Future tests may dictate an increase in simulation depths, although depth per se is not a simulation requirement, it is foreseeable that the testing times may increase to equal the single-crew

mission times for the extended duration space flights now planned. These

times can be expected to be on the order of weeks and months, based

only on the presently planned missions. Immersion times of this duration will establish a simulation diving

mode of operations very different from that mode upon which the past

simulation diving procedures have been based.

It is the objective of this paper to review certain advanced diving

techniques which may be of use when adapted to the special conditions

of extended mission neutral buoyancy space simulation.

CURRENT DIVING TECHNIQUES

The present techniques on which neutral buoyancy simulation is based

are those developed for and in use by military and commercial divers for many years. The equipment is primarily self-contained open-circuit, or surface air supplied open circuit types. The only significant difference

between the standard commercial diving equipment and that used in neutral buoyancy simulation is the substitution of the pressure suit for the hard hat diver's dress. Also, the pressure suited test subject is purposely adjusted to be neutrally buoyant in simulation of zero gravity,

whereas the hard hat diver is heavily weighted with lead shoes. There

is little significant difference between the operation and physical per- formance of the pressure suit and ordinary commercial hard hat diving

dress when both are air supplied at depths to 50 feet. The methodology of present neutral buoyancy operations and ordinary

commercial diving at similar depths is also similar. Of primary concern

in this methodology is to accomplish the return to the surface without

causing the diver clinical problems associated with decompression sickness. Briefly, these problems are usually attributed to the formation of

tissue nitrogen gas bubbles during ascent back to the surface. To allow

these bubbles to come out "silently", or without clinical problems

developing, the diver must follow a decompression schedule during his ascent to the surface. Such a schedule may require the diver to stop

652
and wait a specified time at specified depths before surfacing. It is normal to find commercial divers being required to accomplish such a decompression schedule. In present neutral buoyancy operations, ascent schedules requiring decompression stops are not normally needed.

This is only an apparent freedom from decompression procedures which

is due to the particular combinations of depth and time being used in neutral buoyancy operations. The combination of depth and time are

the critical parameters determining a possible decompression requirement.

The almost complete freedom from the necessity to follow decompression limited ascent schedules, a characteristic of neutral buoyancy

operations to date, is due only to the particular combinations of shallow

depths and short times used. Table I is an extraction of the U.S. Navy

Standard Air Decompression Table and Standard Air Decompression Table

for Exceptional Exposures. It shows that for durations less than 200

minutes at the forty foot level, no decompression stops are required.

The only decompression procedure is a direct ascent to the surface at 60 feet per minute. This time of 200 minutes is known as the maximum

no -decompress ion time for 40 foot depths. Beyond times of 200 minutes

(3.3 hours) a decompression stop is required at the 10 foot depth.

Excessive times (more than 1 hour decompression stop) at this stop level are not encountered up to immersion times as long as 6 hours. At lesser depths the no-decompression time is increased and, according to these same tables, the no-decompression time is infinite at any

depth less than 33 feet. This is shown by Curve 1 , Table 111. The sharp departure from decompression procedures at immersion

depths less than 33 feet is due to the fact that the human body can accommodate surfacing with a certain degree of tissue supersaturation

without developing problems with decompression . The 33 foot depth

limit, the critical supersaturation limit for direct surfacing without

decompression stops, has been accepted for many years as the maximum

depth for infinite time no-decompression exposures. ADVANCED DIVING TECHNIQUES

During the last few years new diving techniques and equipment have

been developed which are radical departures from the diving techniques on which almost all neutral buoyancy simulation has been based to date. The main new diving methodology is known as "saturation diving".

Saturation diving is based on a single key idea, that the body tissues will absorb gases when subjected to new higher gas pressures (depths)

only until the tissues are in equilibrium with the gases at the new

653
higher pressure. This absorption process to tissue "saturation" can be

considered to be complete within 24 hours of continuous exposure at the new higher pressure. Decompression procedures from saturation dives are different. Continuous pressure ascents are used, at the rate of 15 minutes per foot of depth, or four feet of ascent (sea water equivalent pressure) per hour. The ascent (decompression) time is the same for a depth stay of one day

or any number of additional days at any given depth

Depth stays of these extended durations are of course practical only

with the use of a habitat or submerged station. The environmental

pressure of the habitat is kept at the saturation depth pressure level

With no differential pressures, such habitats are of relatively lightweight

structure and direct access from them to the water is easily accomplished.

TVie idea of being able to surface with a critical degree of supersaturation is still retained when saturation diving methodology is employed.

However, recent experience has indicated that for saturation exposures,

the critical maximum depth when breathing air is less than the 33 foot

(sea water pressure equivalent) limit previously accepted. These experi- ences suggest that saturation dives to a maximum depth of 25 feet can

be followed safely by direct ascent to the surface. Additional diving advantages may be gained by using the technique

of "excursion diving" from the submerged habitat level. Excursion

diving can be described as dives from a submerged station to greater

depths with subsequent return to the submerged station. Table III, taken from a paper given by Larsen and Mazzone in a symposium on

Potential Advances in Deep Diving, shows these advantages. In this table the saturated divers are projected as operating from a station located at the 35 foot depth level . The no-decompression limits for direct ascent back to the station level are shown in curve 2. Considerably greater no decompression times are available at any depth when

the divers are operating as saturated divers returning to the habitat

level, than when operating as non-saturated divers returning to sea

level surface pressure. Curve 2 appears to be asmytotic to the TOO foot depth level . This implies that the maximum depth infinite time no-decompression limit

is 100 feet for saturated divers operating out of a 35-foot depth level habitat, it is obvious that divers operating from such a submerged

saturation station, even if the station is located at relatively shallow

depths, will have a significant increase in the infinite time no-decompression maximum depth. More recent publications by Schreiner and

654
Kelley give methods for computation of decompression schedules for repetitive saturation -excursion dives. These schedules also show time

advantages when compared to equal depth-time repetitive dives conducted from a surface base. Advanced diving equipment may also have some bearing on long duration shallow water immersions. Habitat environmental control systems

and closed circuit rebreathers (closed circuit SCUBA gear) have been

in use recently . These systems can be set to supply the diver with oxygen/nitrogen ratios different from that found in air diving. If the oxygen

partial pressure can be elevated, say from 0.2 atmospheres to a typical

value of 0.4 atmospheres, then for any given depth the nitrogen partial pressure (pN2) will be correspondingly lower than when using air at that depth

With the adjustment stated, the maximum safe depth limit for no decompression of 25 feet would be increased to about 44 feet. Stated in another manner, the nitrogen partial pressure at 44 feet (gauge) depth

breathing a 0.4/0.6 p02/pN2 ratio gas would have an equal nitrogen

partial pressure normally encountered with air breathing at immersion

depths of only 25 feet. NEUTRAL BUOYANCY APPLICATIONS OF ADVANCED DIVING

TECHNIQUES

One of the simplest applications of advanced diving equipment and

techniques to extended duration neutral buoyancy testing would be to utilize the new closed circuit rebreather SCUBA systems. Using these systems with a p02 elevated above the 0.2 atmosphere value available

with air would increase the diving depths and times available with nodecompression dives. Using this method, the crew station (simulated) could be located at surface pressure and infinite time no-decompression

dives to 40 or 50 feet would be possible, even with only moderate p02

elevation

Should simulation needs develop for depths and times greater than would be feasible by the simple adjustment of the diver's p02/pN2

breathing gas ratios, then selected application of saturation diving

methodology could relieve the simulation series from involvement with

limiting ascent schedules . For instance, if high fidelity simulation were

required which involved space structures of sufficient size so that depths

to 100 feet were necessary for complete mockup immersion, and if simultaneous with this depth requirement, continuous work/sleep schedules of several days were required to follow accurately the crew flight

655
program, then obviously new techniques would be required. Under simulation requirements just described, application of saturation diving techniques might yield the following typical solution: The crew quarters could be located at a saturation depth equal to or less than the infinite time no-decompression maximum depth. With air (p02/pN2 = 0.^0-8 atm), this depth could be 25 feet. With pOypN2

ratio = 0.4/0.6, the infinite time no -decompress ion depth would be in excess of 40 feet

From this simulated space station base depth, the saturated simulated space crew could use excursion diving techniques to perform their work

cycles under null gravity conditions at maximum mockup depths without

decompression limited ascent back to the base station. Using such procedures, the simulation would not have space flight-irrelevant involve- ment with decompression procedures.

The situation would have three distinct categories of divers:

1 . The non-saturated diver. These divers would be involved as safety and support divers, test engineers and directors, whose functions are in support of the simulation study. They would dive from the surface and return using standard decompression ascent procedures.

2. The habitat-saturated diver. These divers would be saturated

at the habitat pressure level and would therefore have direct ascent to the surface without decompression involvement. This category would

include simulated space crew members not required to perform work

cycles exposing them to greater (excursion) depths. It would include

those persons involved with conducting, supporting, and/or studying

the simulation, whose needs included intimate stays aboard the station

for long periods, but whose needs did not expose them to greater

(excursion) depths.

3. The saturated-excursion diver. These divers would include

those simulation crew members and other personnel whose duties required exposure to excursion depths. This category of diver would have some

requirement for decompression before return to surface pressure. Assuming that a habitable saturation station is already available at shallow

"no decompression" depths, in order that a multi-day simulation be made

possible, thenthis shallow station could be used with p02 adjustments to reduce the category 3 diver to a category 2 diver. Such decompression

might be possible without transfer to special decompression chambers

other than the already available shallow saturation station. Such decom656
pression might be accomplished throughout the lost rest and/or sleep period

before returning the simulated space crewmember to the surface. CONCLUSION

Advanced diving techniques and equipment are available which can

facilitate simulation fidelity for extended duration simulation missions. These techniques primarily will involve special adjustments of saturation diving techniques and equipment, special use of advanced closed circuit SCUBA gear, and special p02 adjusted decompression methods. These advanced techniques will be especially useful where structure size dictates greater depths than are currently being used. The adoption of such saturation methods will require extensive computations for the special case chosen. Also a thorough chamber testing pro- gram would be necessary to validate the computations, as is done for proposed new schedules.

TABLE I Depth (Ft.) Bottom Time Time at 10 Ft. Decompression

(Min .) (Hr.) Stop (Min .) 40 200 3.3 0

210 3.5 2

250 4.1 11 300 5.0 19

360 6.0 23

720 12.0 69

TABLE 11 Depth (Ft.) Bottom Time Decompression

40 Saturation Exposure 10 hours continuous ascent

(24 hours or more) decompression

657
658
REFERENCES

1. Bornmann, Robert C. Decompression after Saturation Diving. Proc . TViird Symposium on Underwater Physiology, Lambertsen, C.J. , Ed., The

Williams and Wllkens Co., 1967^

2. Hamilton, R.W. Jr., P. Fructus and X .R. Fructus: Physiological

Surveillance and Performance Tests During A Seven-Day Exposure to 9.5

Atmospheres He-02 with Exercise Excursions at 13 Atmospheres. In Proc.

International Congress Physiol. Sciences, Washington, D.C., 1968.

3. Hartman, H. and H.D. Fust: Pressure Chamber Experiments of Long

Duration at Pressures Between 4 and 23 Atm., In Drager-Heft 265, Lubeck:

Drager-Werke 1966.

4. Larsen, Reynold T. and W.F. Mazzone: Excursion Diving from Saturation Exposures at Depth. Proc. Third Symposium Underwater Physiology,

Lambertsen, C.J. ed . The Williams & Wilkens Company, 1967.

5. Schreiner, H.R. Mathematical Approaches to Decompression . Int.

J . Biometeor 11:301-310, 1967.

6. Schreiner, H.R= and P.L. Kelley. Computation Methods for Decompression from Deep Dives. In Proc. Third Symposium Underwater Physiology, C.Jo Lambertsen, ed. Baltimore: Williams and Wilkins, 1967.

7. Schreiner, H.R. and P.L. Kelley. A Pragmatic View of Decompression. In: Proceedings Fourth Symposium Underwater Physiology, C.J.

Lambertsen, ed . , New York: Academic Press, 1970.

8. Schreiner, H.R. The Design of Decompression Schedules, A Symposium Man's Progress Into the Sea, Marine Technology Society, Washington, D.C. 1969.

9. Schreiner, H.R. and P.L. Kelley. Computation of Decompression

Schedules for Repetitive Saturation-Excursion Dives. Aerospace

Medicine 4 1(5), 491-494, 1970

10. Workman, R.D.: Calculation of Decompression Schedules for Nitrogen-Oxygen and Helium-Oxygen Dives. Research Report 6-65,

Washington. U.S. Navy Experimental Diving Unit, 1965.

659

Paner No. UU

A METHOD TO OBTAIN AN ULTRACLEAN ENVIRONMENT

Maurice S. Cridlin, Vacuum Technologist, and Joseph W. O'Connor,

Aerospace Technologist, NASA Goddard Space Flight Center, Greenbelt, Maryland.

ABSTRACT: The Goddard Space Flight Center has designed and fabricated a system which offers an effective approach toward solving the problems usually encountered in creating an ultra high vacuum. The vacuum techniques described find application in an Auger Spectrometry System.

This ionization pumped system is unconventional in that; (1) In- terior chamber cryogenic shrouding is eliminated. In lieu of the

internal shrouds the method is to submerge the entire chamber exterior in a dewar of liquid nitrogen. Eliminating the shrouds within

the chamber eliminates sources of contamination and the exposing of the wall exterior to the cryogenics eliminates the typical ambient

temperatures that chamber walls normally experience, (2) The ionization pump is located within the ultrahigh vacuum chamber, rather than

outside. This eliminates the gate valve typically associated with an

externally mounted pump and permits greatly increased realization of

the conductance pump capacity. The pump is protected by use of gas

purging when pressures are increased to atmospheric.

The method of precleaning, prior to bakeout, involving high temperature gas purging is described as well as the entire systems

operation. Pressures in the order of 7 x 10"^^ torr have been observed.

KEY WORDS: vacuum, pumping, ultraclean, cold shrouding, environment, contamination, low temperature, low pressure, ion pumping,

anti-migration traps.

A METHOD TO OBTAIN AN ULTRACLEAN ENVIRONMENT

An ultraclean environment is often mandatory in space oriented

research, particularly when experimental materials must be free of

661
surface contamination. The Goddard Space Flight Center has designed

two ultra high vacuum systems using an effective approach to solve some of the problems usually encoimtered while attempting to create an oil-free environment. One of these systems is used to develop

ultra high vacuum components and techniques. The other system is used for Auger Spectrometry research. The spectrometer can detect

partial monolayers of materials and identify these materials. This

unique system is one example clearly demonstrating the need for a noncontaminating experimental environment. Since an oil free environment for this Auger Spectrometer was mandatory, an ionization

pumped system^ was developed, which will be described here.

Conventionally, ion pumped ultrahigh vacuum systems employ an

internal optically dense liquid nitrogen shroud^^^ in conjunction with a titanium pump. Figure 1(A). The ionization pump is mounted external

to the chamber and isolated by a gate valve. Typically, the gate valve

utilizes an elastomer'-^'* O-ring, While this geometry and selection of components finds wide acceptability, improvement can be realized by

locating the ionization pump inside the chamber. Figure 1(B); by eliminating the gate valve and the cyro shrouds and by slightly changing the typical pump down technique.

ULTRAHIGH VACUUM CHAMBER

The entire vacuum system. Figure 2, is constructed from metal

and all seals are metallic. Ceramic to metal feedthroughs are used.

Within the cylindrical chamber (20 -inch diameter by 36 -inch high) are

the ionization pump, titanium sublimation pump, an electron trap and

shelves for mounting experiments. Thermocouples monitor the chambers' wall at four separate elevations. The work area of the chamber is top loaded through a choice of metal sealed orifices up to 18 inches in diameter. The dome includes an observation port and a high intensity light port.

Surrounding the vacuum chamber on the sides and bottom is a metal thermos bottle. The annulus between the chamber and the thermos houses heater elements and when the heaters are activated,

this area becomes an oven for bakeout. The area is also capable of being filled with liquid nitrogen, thus insuring the walls of the chamber

being uniformly cooled to temperature. Cryogenics are not al- lowed to contact feedthroughs directly, they are cooled through conduction. LN^ can be gravity drained through a valve in the bottom of the thermos.

'^Numbers in parentheses refer to the list of references appended

to this paper.

662
Optically baffled titanium pump filaments are mounted through the center bottom of the chamber. When used, titanium is vaporized onto

the LN2 cooled walls of the chamber providing an intrinsic pumping

speed in the order of 60,000 1/sec (approximately 60 1/in^/sec).

A 250 l/sec ionization pump is located above the titanium pump

well, a position near the center of the chamber vertically. Conductance for the ion pump far exceeds its capacity and the solid angle of view it possesses is much improved over external mounting techniques.

An oven hood covers the top of the chamber and all feedthroughs

during bakeout cycles. It is planned to seal this hood to the system at a later date so the oven/liquid nitrogen area can be used for a third

beneficial function, a differently pumped area completely surrounding

the vacuum chamber and feedthroughs. Currently, this feature does

not exist.

GAGES - - Three separate gages were used to evaluate the chambers' performance during the equipment checkout phase and installation. The

location of two of these is shown in Figure 3, These included the ioni- zation pump gage which reads pressures to 1 x lO"^'^ torr, a cold

cathode-type gage and a hot cathode gage which cathode gage which

read to 1 x 10'^^ torr. The hot and cold cathode gages are mounted at the top of the chamber. As one might expect, variations in pressure

readings were observed, however these were well within an acceptable

range. Pressures presented in this report are from the hot cathode

gage.

PUMPS

Five different types of pumping mechanisms are used in this system: a mechanical pump, a sorption pump, an ionization pump, a titanium pump and liquid nitrogen cooled surfaces which pump considerable gases and vapors. The mechanical pump and the sorption pump are used for roughing

the chamber. The mechanical pump also evacuates the double walled

thermos bottle. Dry nitrogen is admitted into the chamber to maintain

a pressure of 200 microns while roughing with the 15 cfm mechanical

pump. This assures only viscous flow during the first stages of roughing. Roughing is completed by valving out the mechanical pump

and the nitrogen gas supply and then activating the sorption pump.

The ionization pump, titanium pump and liquid nitrogen cooled

walls are used to evacuate the vacuum chamber to lower pressures.

663
By location the ionization pump inside the vacuum chamber, the typical

gate valve can be eliminated.

When rapid re-cycling of the chamber from atmosphere to ultra high vacuum is required, a feature not normally expected of an ultra- high vacuum system, elimination of the gate valve can offer some

advantages. All gate valves decrease the conductance and reduce the throughput of a djTiamic system. They also present additional surfaces to be cleaned and where elastomer O- rings are used, bakeout

temperatures are often restrictive. Protection of the ion pump from

contamination is accomplished by special operating procedures which

are strictly adhered to. Venting the system to dry nitrogen while the system temperature is in the order of 125''C is necessary. The sys- tem is never allowed to see atmospheric pressure for over 15 minutes. Shorter exposures are more the rule. BAKEOUT

All surfaces exposed to the ultra high vacuum are baked with good

imiformity to 450°C. When it is desired to obtain a very low pressure

(less than 1 x lO"^"^ torr), the chamber is cooled over a period of around eight hours so that all surfaces are cooled at nearly the same

rate. Heat is added in ever decreasing increments during the cool

down cycle to achieve this goal. When the chamber temperature ap- proaches 50°C, liquid nitrogen is introduced imtil the entire chamber

is submerged. By reducing the wall temperature the vapor pressure

of all materials located thereon is reduced, consequently, even those

gases exhibiting virtually no partial pressures on a mass spectrometer

are affected. The mild bakeout, 450°C is very effective when the chamber is cooled to LN2 temperature.

Since typical cryogenic shrouds are eliminated, their problems go with them. They are difficult, if not impossible, to bake out or cool

uniformly. Reduction of surfaces exposed to the ultrahigh vacuum is of prime importance, to eliminate sources of contamination.

FORE LINE TRAPS

Two traps are used in the foreline. A bakeable activated absorbent trap is located next to the mechanical pump. This trap eliminates

much of the oil backstreaming from the mechanical pump. It elimi- nates some of the oil migration. To completely trap oil from the mechanical pump, the second trap employs a removable and throwaway section. Figure 4, so that oil cannot migrate through it when it is at ambient temperature. The second trap is activated with liquid

664
nitrogen such that oil cannot migrate through it when it is charged. In

all cases, this trap is exposed to viscous flow in the direction of the mechanical pump, a feature which inhibits oil backst reaming.

Mechanical pump preliminary evacuation to 200 microns is an important feature of the operating procedure. Mechanical pumps evacuate all gases and are not selective as are sorption pumps.

OPERATIONAL SEQUENCE

Initial roughing of the chamber is accomplished with the double

trapped mechanical pump to a controlled pressure of around 200 microns. Dry nitrogen is metered into the chamber to prevent the mechanical pump from evacuating the chamber to a lower pressure.

Additional evacuation is completed with the sorption pump while the mechanical pump and gas bleed are valved out of the system. The

entire evacuation procedure and sequence is performed after the chamber has been brought to at least 150°C. Roughing with the mechanical pump should continue for at least ten minutes. When the

chamber pressure is below 10 microns, the sorption pump is valved

out and the ionization pump is started. (The pressure can be reduced

to around 1 x IQ-"^ torr with liquid nitrogen prior to starting the ioni- zation pump if the pump seems hard to start.) Within a few minutes, the ionization pump will have pumped the chamber to a pressure of 5 x lO"^ torr or lower. To achieve lower

pressures, a bake cycle is initiated. Figure 5 illustrates a pump down

cycle achieve with the Auger Spectrometry system. The newer sys- tem currently being evaluated is expected to better the pump down

cycle exhibited by the first system,

ENVIRONMENT ACHIEVED

The initial cleaning of the system involved a long time (72 hours)

bakeout. Normally a four hour bakeout will be sufficient to enable the chamber to achieve a vacuum in the 10"-^^ torr range or lower. The

predominating gas species at 10"^^ torr are CO and He. In the

10"-^^ torr range H2 and He were the dominant constituents.

665
REFERENCES

(1) S. Dushman, J. M. Lafferty, Scientific Foundations of Vacuum

Technique, John Wiley & Sons, Lie, 1962

(2) C. M. VanAtta, Vacuum Science and Engineering, McGraw-Hill

Co., 1965

(3) H. A. Steinherz, Handbook of High Vacuxmi Engineering, Reinhold

Pub. Corp., 1963

(4) D. J. Santler, H. D. Holkerboer, D. W. Jones, F. Pagano, Vacuum

Technology and Space Simulation, NASA SP-105, 1966

666
^^^^^^^OVEN LID

Fig. 1—Typical and GSFC ion pumping techniques.
Fig. 4—Anti-migration trap.

Fig. 5-Auger-spectrometer

pump down cycle. 668
Paper No. /+5

A PROGRAMMABLE DYNAMIC THERMAL VACUUM SYSTEM FOR

SOLAR ARRAY COMPONENT TESTING

John W. Fairbanks-^ and Marshall B. Eck^

REFERENCE: Fairbanks, J. W. and Eck, M. E. , "A Programmable

Dynamic Thermal Vacuum System For Solar Array Component Testing," Fifth Space Simulation Conference, Gaithersburg, Maryland,

September, 1970. ABSTRACT: The Programmable Dynamic Thermal Vacuum System

(PDTVS) was designed to provide long-term laboratory evaluation of solar array components imdergoing thermal cyclic effects anticipated

for specific missions. This system is capable of duplicating the temperature excursions, including rate-of-change of temperature, for practically all missions currently under consideration at NASA. The

equipment can obtain test specimen temperatures between +150°C to -120°C with a rate-of-change of at least 6°C/min. The PDTVS can accommodate thirty-six 6"x6" test specimen holders and position each

one in front of a illumination port for periodic in-situ electrical measurements with a solar simulator.

Basically, the facility consists of an ion-pump vacuum system with

2 bell jars, thermal conditioning equipment, a 3 stage blower, a programmable control unit and an LN2 storage tank. The novel developments in this system Include a high-energy low-mass heater which employs radiative-convective coupling of a gas stream to an IR source. The test carrousel shaft is used for specimen positioning and as an in- strumentation penetration. Initial difficulties required modification of the blower bearings, shaft and housing to accommodate the wide range

of gas stream temperatures without affecting performance.

KEY WORDS: Solar cell, solar array, thermal cycle vacuum testing,

gaseous nitrogen heater , gas stream blower, thermal shroud, GN temperature conditioning.

AST Electrical Propulsion and Power Systems , NASA-Goddard Space

Flight Center, Greenbelt, Maryland

'Chief Engineer, H. C. Slack & Associates, Baltimore, Maryland

869
INTRODUCTION

The ability of solar arrays to provide the power requirements for missions now in the conceptual and preliminary design stages at NASA

require thermal cycling data which is not now available . Problems have

been encountered on previous projects with the solar cell-interconnector

solder joint aad/or the solar cell contact integrity that have been solved

primarily by "cut and try" techniques. Part of the problem stems from

the broad range of thermal conductivities of the conductive and nonconductive components which comprise the solar array. The array as- sembly with this inherent thermal mismatch must undergo repeated

thermal excursions typically with a A T as high as 230"* C. Current design thermal flight qualification procedure generally consists of subjecting the prototype solar panel or paddle to 2 few thermal vacuum

cycles to extremes about 10% beyond maximum and minimum flight predicted temperatures. Usually there are no more than rudimentary

steps taken to ensure that the rate-of-change of temperature during

shadow encounter and emergence is reproduced.

Consideration of some of the specific missions being planned,

should assist in defining the test equipment requirements. The Orbiting Astronomical Observatory (OAO) program is seriously considering a ten-year life requirement for the solar array, which means approximately 50, 000 thermal cycles (Ref. 1). Data on conventional solder

connected cells indicates problems can be anticipated after 3, 000 cycles (Ref. 2). The Astronomy Explorer (AE) C&D will encounter significant aerodynamic heating during a series of low perigee passes

(Ref. 3). Considerable interest is also expressed in solar array capabilities on close -approach solar missions as well as the capability to operate in the vicinity of Jupiter. Perhaps of equal importance is the need to obtain comparative long-term thermal vacuum test data on

solar array components or composites generated in Research and Development (R&D) programs which have the potential of cost saving, improved performance, weight reduction and/or increased reliability.

The current conservative philosophy of flying with proven components

has generated a significant time lag between R&D and flight utilization despite potential technical and/or economic advantages.

The Programmable Dynamic Thermal Vacuum System (PDTVS) was designed to support flight projects in the preliminary solar array

design stage and as problems are encotmtered during flight qualification. In addition, the PDTVS provides meaningful long-term thermal

cyclic effects data on R&D program developments which enhances their

implementation in flight programs. In order to achieve these design

objectives, the following system requirements were developed.

670
DESIGN REQUIREMENTS

The requirements generated for the PDTVS were as follows

1. The system must be able to reproduce thermal conditions anticipated for essentially all NASA missions under consideration for the

70's which would employ solar arrays.

a. Specifically operate between +150°C to -170° C and produce a rate-of-change of at least 6°C/min. between +150° C and -120°C.

2. Provision must be incorporated to accommodate in-situ periodic

electrical measurements of test specimens through an illumination port.

a. Each of thirty six 6"x6" test specimens must be mounted so

that it can be accurately positioned in front of the illumination

port.

b. Also must be capable of accommodating an EVIP-I solar panel.

3. Each test specimen must have a 99% view factor to thermally

controlled surfaces.

4. Test instrumentation wiring shall produce minimum interfer- ence to the thermal view factor.

5. System must be capable of programmable thermal cycling

performance.

6. System must operate automatically over long periods (on the order of 3 to 9 months)

a. Fail-safe features must be included.

b. Permit ready accessibility for operational and maintenance

functions

7. Must incorporate features that provide economical operation,

including low-level maintenance requirements.

8. A second chamber shall be incorporated such that it could be

thermally slaved to the programmed thermal cycle. This chamber

would accommodate small numbers of samples which could be re- moved without affecting the long-term testing on the larger unit.

9. A common roughing system shall service both chambers with

capability for a third chamber.

671
The above requirements have been incorporated in the PDTVS,

which was built on-contract for the Goddard Space Flight Center. OVERALL FACILITY

The PDTVS facility is shown in Figure 1. The major components which comprise the PDTVS facility are a

getter-ion pump vacuum system with two bell jars, a solar simulator, a test carrousel with automatic indexing, thermal conditioning system,

programmable controls, liquid nitrogen storage tank and associated

piping. Several innovative features were incorporated in the thermal

conditioning equipment and in the test carrousel positioning scheme.

VACUUM SYSTEM

Roughing for both bell jars is achieved through physical adsorption

of gas molecules by LN2 cooled high-capacity molecular sieves. Pumping into the high vacuum region is accomplished by sputter-ion pump

combined with titanium sublimation pumping. The sputter-ion pump

operation involves trapping gas molecules and atoms through the interaction of gas ionization, titanium sputtering, chemical combination and

ion burial. The titanium sublimation pump sublimes titanium from

heated filaments which is deposited on a "line-of-sight" transmission

to shielded non-working area. Pumping is affected as the active gases

combine chemically with this deposited titanium film to form stable compounds. This system provides pressures of 10 torr and below

under typical test conditions.

The getter-ion pumping system was chosen rather than systems

that involve introduction of long-chain hydro-carbons. The combination

sublimation and sputtering-ion pump system was considered developed

sufficiently to meet the long-term automated operational requirements.

This system is essentially immune from failures in power, water or

air-service.

The 24" diameter bell jar was chosen as it could accommodate an

Interplanetary Monitoring Platform (IMP) -I solar array panel within

the thermal shroud. The IMP-I panel will probably be the standard design for the next series of in-house spacecraft at the Goddard Space

Flight Center. This particular size bell jar was available as a standard

size in the industry. The 12" bell jar was designed to accommodate a smaller number of test specimens. This feature would permit periodic

removal of specimens for optical inspection and testing without interrupting the long-term thermal cycling test. Both bell jars are fitted with an illumination port of UV-grade fused quartz to permit tn-situ solar

672
illumination measurements. Calibrated standard cells and disc radiometers are mounted to insure the desired solar illumination intensity

level. The spectral distribution of the simulator can be checked with

a spectrophotometer in the same lab. THERMAL CONDITIONING SYSTEM

The Thermal Conditioning System (TCS) was designed to provide

temperature stabilized GN or LN2 to the cryoshroud of either of the two or both of the bell jars. The TCS supplies gaseous nitrogen to the selected bell jar shroud at any temperature between -170° C and +150°

with a maximum shroud temperature difference of ±4°C while the shroud

is under a distributed thermal load of 140 watts. High flow rates of thermally conditioned gaseous nitrogen transform the shroud into the necessary radiation heat source and/or heat sink to obtain the desired

test specimen temperature. Test specimen heating and cooling is effected by radiation coupling

between the shroud and the sample. The thermal shroud system in- cludes a movable shutter in the illumination port and an optically dense

baffle which shields the vacuum system internals. The only constraint

preventing a 100% view or thermal shape factor from any point on a

test panel to a controlled thermal surface is the carrousel shaft. The shroud wall rate of temperature change is sufficiently rapid

to produce a 6°C/min. change in specimen temperature at any level between +150°C and -120°C. The specimen specific weight and surface

emissivity have a large effect on the attainable rate-of-change of temperature. The value of 6°C/min. is obtainable with silicon solar cells having fused silica coverslides attached such that the composite has a

specific weight of 0. 83 lbs/ft or less. Modulating butterfly valves proportion the gas flow through the GN

to LN2 heat exchanger. This flow variation combined with a controlled

power input to the heaters produces a very precise means of affecting

temperature control. There is also a large potential heat source or

heat sink reserve. The response time and temperature overshoot of the system are mmimized by biasing the system source and sink

against each other. The basic components in the TCS are (1) a high-energy density low- mass heater; (2) a three-stage centrifugal blower, and; (3) a liquid

nitrogen boiler. These components, their inter-connecting piping and

associated controls (LN2 level, blower power, high and low temperature interlocks , etc . ) are mounted in a caster mounted 2.5'x4.5'x6'

673
equipment cabinet. A block diagram of the system is shown in Figure

2. The TCS components are briefly described as follows:

a. Blower : A three-stage centrifugal blower is used to circulate

approximately 220 CFM of nitrogen gas through the system.

This blower required considerable modification of a commercially available unit in order to operate with the range of gas temperatures involved. These modifications entailed the in- stallation of super precision bearings, a polished stainlesssteel shaft fitted with a specially designed fan and casing re- design to improve the heat dissipation. These modifications

limited the bearing temperature excursions to a fraction of the GN temperature range.

b. Heaters : Rapid transients require that system mass be minimized. This is particularly critical in the heater design as

this element runs at the highest temperature of any system

component. The system heater shells are fabricated from 3 inch diameter thin-wall stainless-steel tubing. Each heater

module contains a 1. 6 KW quartz IR lamp and a set of thin anodized aluminum fins . The fins are radiatively heated by the IR lamp and are convectively cooled by the GN stream. The

weight of the heater extended surface is approximately one

(1) lb.

c. Boiler : The system heat sink is a LN2 boiler. Gaseous ni- trogen is pumped through the tube side of a shell and the tube

heat exchanger. The shell side of this unit is flooded with

LN 2- The liquid level is maintained with a temperature probe

controller, and solenoid valve . Boil-off LN 2 is mixed with

the exit gas flow stream to aid in lowering the circulating gas stream temperature. A pressure relief valve is located in the main gas stream to prevent excessive pressure buildup

either from the introduction of boil-off gas or from gas ex- pansion during increasing temperature transients in the loop.

The relief valve maintains the GN loop pressure at 3psig.

d. Control : Operational control is provided by a solid controller

having proportional band, rate and reset functions and a dynamic setpoint programmer. The controller has a set point

accuracy of ±2 ° C. Control is obtained by sensing the shroud wall temperature and

comparing the measured value to either a fixed or dynamically programmed set point. The resulting error signal is amplified and trans- formed into a control voltage. This voltage is used to control the output 674
of a silicon control rectifier (SCR) and a pneumatic valve positioner.

The valve positioner operates two butterfly valves which are 90° out of phase. The amount of cooling provided is determined by the ratio of the flow through the boiler to the flow through the boiler by-pass. The

amount of heating provided is a function of the magnitude of the error

signal and the output of the SCR. There is sufficient overlap between

the heating and cooling functions such that transitional operation is esentially continuous. This variation in flow combined with a variation

in power input to the heaters produces a precise means of effecting

temperature control. There is also a large potential heat source or

heat sink reserve. The response time and temperature overshoot of the system can be minimized by biasing the system source and sink

against each other. An over-temperature safety interlock using a pyrometer and relay automatically cuts out the heater voltage when the temperature hits a 150° C upper limit and automatically cuts the heater

back in at 145 °C if required by the programmed input. A schematic of the system control logic is presented in Figure 3. TEST CARROUSEL

A carrousel, or test specimen holder, was designed to accommodate the maximum number of test specimens such that each specimen

can be accurately positioned at the illumination port. The system thermal performance objectives required the carrousel to be extremely

lightweight with minimal thermal mass. The carrousel design that emerged was duo-decagon frame that could accommodate thirty six

6"x6" test specimens. The specimens are mounted along the periphery of the carrousel in three adjacent rings. Thus, as the carrousel

is rotated and/or translated, each of the test specimen holders is positioned in front of the illumination port. Solar simulation is provided

by a filtered 2500 watt Xenon arc lamp simulator which closely approximates the solar spectrum in space. This capability permits in-situ

solar cell electrical characteristic measurements periodically during

the test. A test section is included on the carrousel to calibrate the solar simulator illumination at the test plane. The solar simulator

can be rotated such that its beam is directed at a spectrophotometer to periodically analyze the simulator's spectral distribution. CARROUSEL POSITIONING

The initial design efforts to provide positioning of the carrousel

at the illumination port involved various mechanical linkages with rotary and translatory motion feed throughs . These systems could be

characterized as being complex, involving considerable thermal mass,

imposing internal volumetric constraints, introducing conductive thermal paths and requiring more than one shroud penetration. A novel

675
approach to the positioning problem was suggested by Maurice Cridlin

of the Goddard Space Flight Center. He had devised a sample positioning scheme for a facility that required in-situ measurements of particle

radiation effects on the solar absorptance of materials (Ref. 4). This

scheme involved attaching the carrousel to a hollow vertical shaft on

the center-line. On the PDTVS the shaft penetrates the thermal shroud

and bell jar and provides a means of indexing the carrousel through

physical movement of the shaft. Additionally, the hollow shaft provides an instrumentation penetration.

The vacuum seal around this shaft is maintained by use of an ante- chamber with two viton "0" rings adjacent to the bell jar. This method

requires a roughing pump to separately maintain approximately 10"

torr vacuum in the ante-chamber.

Test specimen positioning accuracy was increased and test time was reduced by the addition of two remote controlled motors. One

motor provides the desired vertical movement to align any one of the three rings of specimen holders at the plane of the illumination port by

moving the shaft and the rotational indexing platform. The second motor, which is moimted on the rotary indexing platform, indexes each of the twelve specimen holders comprising the ring. This rotary motion

is restricted by stops to ±180° in order to minimize instrumentation

lead movement. The alternate approach of using instrumentation slip

rings was considered to introduce unwarranted problems.

Instrumentation capability is increased by the use of a stepping

switch installed in the carrousel. This switch connects the instrumentation leads to each of the three rings on command. Thus, only forty

leads are fed through the shaft. The leads are potted in the shaft to sustain vacuum. PERFORMANCE TESTING

The TCS was performance tested and found to meet the program

design goals . The actual transient performance of the system is shown in Figures 4 and 5. The initial cooldown rate from a steady state temperature of 20°C was 14°C per min. for the shroud wall and 6°C per min. for the test specimen.

The initial cooldown rate from a steady state temperature of 150° C was 25° C/min. for the shroud wall and 11° C per min. for the test specimen.

676
The initial heat up rate from -140' C was reduced somewhat by the

fact that the shroud wall and test specimen were not in thermal equilibrium at the time the test was initiated. The initial heat up rates from

-140'C were approximately lO'C/min. for the shroud wall andT'C/min.

for the test specimen.

The initial heat up rate from a steady state temperature of 20" C was 4. l°C/min. for the shroud wall and 3.4-C/min. for the test specimen.

The rapid temperature change of the shroud wall and the close

coupling between the wall and the specimen are particularly relevant. Operation in the flooded LX^ ciode was conducted during Acceptance Testing. The results of the flooded LX2 mode test are sho-^Ti in Figure 6. cose LUSIOXS

The PDT\'S provides a test capabilitj- previously unavailable for solar array components. This capabilit\^ is necessary to evaluate candidate components for the more exacting requirements of currently con- templated space missions. Specifically, it provides the following capabilities for solar array component testing:

1. PDT\^S can accurately reproduce the rate-of-change of temperatures experienced b\' solar arrays in tj^pical orbits and in a wide

range of probe missions.

2. It is programmable to reproduce desired thermal orbital conditions on a long-term basis and with a statistically significant number

of solar array components.

3. It enables in-situ solar cell measurements to be taken periodically during long-term thermal vacuum testing.

4. PDT\'S is automated for unattended operation. Safety* devices

will shut dov-TL equipment in the event of malfunction.

The PDT\,'S incorporates several novel design innovations. The

high energj^ density-low mass heater developed for this system was a novel application of radiative-convective coupling of a gas stream to an ER source. The carrousel shaft use for positioning and as an in- strumentation penetration was an adaption of a recent GSFC development. Z^Iodification of the blower to allow operation over the potentially troublesome range of GX temperatures presented a formidable

design challenge.
REFERENCES

1. Purcell, J. , "Space Astronomy Present Reality and Future Prosspects," Proceedings of the IEEE International Geoscience and

Electronics Symposium, Washington, D.C., April 1970.

2. Ralph, E. , Roger, J. , "Silicon Solar Cell Interconnectors For

Low Temperature Applications," Proceedings of the International

Colloquium On Solar Cells, Toulouse, France, July 1970.

3. Grimes, D. W. , Spencer, N. W. , "GSFC Specification For At- mosphere Explorers AE-C and -D," S-620-P-1, Greenbelt,

Maryland, April 1969.

4. Cridlin, M. S. , et al. "An In-Situ Facility For the Determination of the Effects of A Space Environment On The Solar Absorptance of Materials," X-713-69-257 , Goddard Space Flight Center,

Greenbelt, Maryland, Jime 1969.

678
679
GANGED BUTTERFLY VALVES

- 162°C

/ f-162°C / \-172^C

210°C

-173°C

BLOWER

172°C

HEATER

^^^^-172°C

-i7rc

SHROUD

172°C

NOTE:GN temperature shown for steady state operation at -170 C. Fig. 2—System block diagram

GANGED BUTTERFLY VALVES —

I I

VENT VALVES

TC3

TC Boiler liquid level control solenoids

series parallel (2 valves) redundant TC 2 Shroud liquid level control

(liquid mode) 2 valves TC 3 System over temperature control

secures system with over temperature opens heater circuit

TC 4 Liquid interlock prevents going from

liquid mode to gas mode whenever

liquid is present in the shroud TC 5 Shroud wall temperature sensor 6 System Temperature Controller

(Rl Controls Series 3000)

7 Dynamic Set Point Programmer

(Rl Controls Data Trak) Fig. 3—Control schematic

680
0 5 10 15 20 25 30 35 40 0 10 20 30 40 50 60 70 TIME (Min.) TIME (Min.) Fig. 4—System transient performance

681

Paper No, 46

THE CRTOPUMPING SYSTEM FOR THE SPACE SIMULATION CHAMBER

1 2 Hiroshi Nakagava and Ryuji Tsunoda

ABSTRACT: Recently, the space simulation facility which has distinctive

feature was completed in Japan. This facility consists of main- and sub- chamber and other necessary equipment. The main-chamber, having 1.8 m

dia. and 6 m length, is accompanied by the pumping system of very high

pumping speed. The pumping system consists of cryopump, diffusion pump, mechanical Roots booster pump and rotary vacuum pumps. The ultimate

_9 pressure of the order of 10 Torr was achieved by this system. The sub- chamber, having 60 cm dia. and 120 cm length, is accompanied by the pumping system which is completely oil free. The pumping system consists of cryopump, sputter-ion pump and cryosorption pumps. The ultimate pressure

of the order of 10~^^Torr was achieved by this system. Determination of the pumping speed of the cryopump is a very complicated procedure and

in this study we propose the new factor: "Arrival Probability" which is computed by Monte Carlo method. The dynamic performance of oil free pumping system, especially the using order of cryosorption pump was investigated, and the residual gases were analyzed by mass-filter. The concept of "Self-contamination Factor" proposed by Dr. R. Chuan is adopted to this study to determine the performance of this system, and

the factors of 0.006 and 0.0012 were obtained under the actual operating

condition. KEY WORDS: cryopumping, ion pump, space simulation

chamber, pumping speed, vacuum.

1. INTRODUCTION

As the Japanese space development project is rapidly progressing,

the demand for the space environment chambers is also escalated. Some

'Nippon Sanso K.K. (Japan Oxygen Co., Ltd.) Tokyo, Japan.

683
space chambers have been constructed and in operation in Japan for the purpose of testing satellite assembly, researching the adaptability of components and studying of fundamental properties of materials. But almost all of these chambers are rather small and simple, and are designed

only to display the high vacuum condition, not considering carefully the true meaning of the simulation engineering for the space environments,

especially from the viewpoints of "cleanness" and "contamination".

Recently, a complete simulation facility which has a distinctive

feature of the reproduction of space environment and rationalized simula- tion system was constructed and delivered to the NASDA (National Space

Development Agency of the Japanese Government) to satisfy the above mentioned requirements.

This facility consists of two chambers, main- and sub-chamber, high

pumping speed and completely oil free evacuation system for the mainchamber and sub-chamber respectively, solar radiation sources and other necessary apparatus.

The main-chamber was designed for the thermal-vacuum testing of the Japanese artificial satellite under environment of high vacuum, solar radiation and heat sink, and the sub-chamber for research on the properties and effects of the components and materials to be used in the space under environment of oil-free ultra-high vacuum and solar radiation.

The vacuum pumping system for the main-chamber consists of a series

of cryo (condensation)pump, diffusion pump, mechanical booster pump

(Roots type blower), and rotary pumps, while the pumping system for the sub-chamber is completely oil free and consists of a cryo ( condensation)

pump, sputter-ion pump and cryosorption pumps. Solar simulator for the main-chamber is a reflector type with a parabora mirror, using three of 6.5 KV xenon lamp and has an intensity

2 of 1.5 KV/m . Solar radiator for the" sub-chamber is a dirrect projection

type with a 2 K¥ xenon lamp and has an intensity of 1.8 KV/m^.

2. DESIGN OF FACILITY

a . Chambers

The main-chamber and sub-chamber are horizontal cylinders made of 18-8 stainless steel and general dimensions of these chambers are tabulated in Table 1. Interior surface of both chambers is sand-blasted and polished to diminish the outgassing from the stains or rough spots on the surface.

The main-chamber can be baked up to 150°C for degassing purpose

by circulating hot triethylene-glycol as a heat-transfer medium through

pipes attached to the outer surface of the chamber, and it can also be quickly cooled down to the room temperature by circulating the cooled heat-transfer medium through the same pipes.

The sub-chamber can be baked up to 250°C by electric heater jacket.

Varming and cooling of the chamber can be done by automatic programming.

The structures of chambers are illustrated in Fig. 1. The doors

of both ends of the main-chamber can be moved by means of electric carts. The gaskets are made of Viton-A, which may release large quantity of 684
gas and vapor, but the total rate could be neglected because of the surface area of gaskets is very small comparing to that of chamber surface. Satellite can be hung from the lod attached to the top of the chamber and rotated.

b . Shroud

The inside of both chambers are completely covered with double

layers of aluminum shrouds and their configuration with cryopanel is illustrated schematically in Fig. 2. The dimensions of shrouds are listed

in Table 1. The shroud is designed to be cooled by liquid nitrogen so as

to be a heat sink of radiation from the satellite and other test specimens

and simultaneously to be a cryopump for vapors and gases which can be condensed at this temperature, and also as a heat shade of the cryopanel.

Cryopanel is cooled by 10-15°K gaseous helium and this acts as a cryo- pump for the other gases.

The shrouds and the cryopanels are assembled from extruded aluminum

pipes with fixed fins as the detailed structure as shown in Fig. 3. Interior surface of the shrouds is coated by black epoxy paint, which

has the emmisivity of more than 95% even in the infra-red range. Welding of aluminum to stainless steel is usually very difficult

and causes many of leak troubles, but in this study we used friction- welded joint which showed the reliable performance.

The conf igulation of shroud and cryopanel affects very seriously

the pumping speed of cryopump, and this will be discussed in detail in section 3C.

c . Cryogenic System

Liquid nitrogen is supplied into the shrouds, to the helium refrigerator and to the cold baffles of pumping system from the pressurized

6,000 liter liquid nitrogen tank. A part of evaporated nitrogen can be recovered by the nitrogen liquefier. Helium refrigerator is a closed cycle, liquid nitrogen pre-cooled

type with a reciprocating expansion engine as shown in Fig. 4, the compressor is a completely leak free type with built-in motor. Leakage

of the helium from the whole cryogenic system is guaranteed for less than 6 m"^/year, therefore there is no need to supply helium for a long

time

As to ensure the utmost efficiency, the whole cryogenic system

including gas helium pipe lines to the cryopanel in the chambers is completely super-insulated. The refrigerating capacity of the system

was designed as much as 100 V at 15°K, and it was proved that more than

150 V of cold energy at 15°K could be available, and the ultimate

temperature of refrigerator itself was measured as below than 10°K. The heat loss at these pipes of gaseous helium was measured as the order of 0.025 V/m. The operating performance of helium refrigerator is shown in Table 2, wherein three cases using main-chamber, sub-chamber

and only the refrigerator are shown respectively. In the latter case,

the cold output was measured by means of a small heater as a dummy which

was attached to the outlet part of the expansion engine. 685
d. Pumping System for the Main-chamber

The pumping system for the main-chamber consists of above mentioned

6 m^ cryopump, a 36-inch diffusion pump, a Roots type mechanical booster vacuum pump and two rotary vacuum pumps. The capacities and specifications

of these pumps are tabulated in Table 3, and the whole system is shown in Fig. 5.The estimated outgassing rate from the wall and other parts in the chamber is listed in Table 4, and the total outgassing rate was believed

as in the range of 1.5xlO~^ Torr.L/sec, but this rate was mostly

depending on the outgassing from the shroud. When the shroud is cooled down by liquid nitrogen, the rate should be decreased to be neglectable,

and the total rate would be estimated at the order of IxlO"^ Torr.L/sec.

2 The surface area of cryopanel is 6 m and conductance of opening

area of cryopanel is about 6x10^ L/sec, then the expected pumping

speed of the cryopump could be presumed as l.TxiO' L/sec. at least. Therefore, the ultimate pressure could be expected as: (lxlO~-^)/( 1.7x10^) = 7xlO~^Torr, when the shroud is cooled down and cryo- pump is" operated

The outgassing rate from the actual satellite usually shows various

values, but if the surface area of components of the satellite is supposed as 10 m^, and outgassing rate is estimated as 2x10 ^ Torr.L/sec.

cm^, then the total outgassing from satellite would be 2x10"^ Torr.L/sec. Therefore the operating pressure of ( 2x10"^ )/( 1 . 7x10^ ) = 1 . 2xlO~'^Torr could be achieved in actual simulation test.

e . Pumping System for the Sub-chamber

The system is designed to diminish the self-contamination of the sample and also to assure of completely oil free "clean" environment.

The pumping system consists of the above mentioned 0.6 m cryopump which

has pumping speed of 20,000 L/sec, a sputter-ion pump of 400 L/sec and four of the liquid nitrogen-cooled cryosorption pump, each containing 1 Kg

of Molecular Sieves 5A. The adsorptive capacity of the above four sorp- tion pumps can be estimated as to evacuate the 300 L chamber to the

_3 pressure of 10 Torr from the atmospheric pressure. After the cryosorption pumps are saturated, the sputter-ion pump and cryopump will be started to operate.

The outgassing rate from the inside of the sub-chamber is listed in Table 4, and the total rate would be the order of 1.1x10"^ Torr.L/sec when the shroud is cooled and the cryopump is operated. The surface area

2 of cryopa^^el is 0.6 m , and the conductance of opening area of shroud

4 / is 5.9x10 L/sec, then the pumping speed of cryopump is expected at the order of 2x10^* L/sec. Therefore the ultimate pressure of: (I.lxl0"^)/(2xl0'^) = 6x10"''^° Torr could be expected when the shroud is cooled down and the cryopump is operated.

3. FUMPING-DOVN CHARATERISTICS

a . Main-chamber

Fig. 6 shows some typical performance curves of pressure down of the main-chamber. Curve A denotes the case of the following procedure: 686
(1) the chamber is baked out and evacuated for a long period, (2) dry

nitrogen gas is introduced into the chamber, (3) the chamber door is kept open for 30 min to be exposed to the atmosphere, then, (4) start

_g

to pump dovn. The ultimate pressure of 2x10 Torr vas observed within

4 hours. Curve B in Fig. 6 shovs the following case, i.e. (l) and (2) are

same as above, (3) is kept for 2 hours, closing the door, then, (4) start to p\imp down. There is little difference between cruve A and B at the final

period, of pumping down, when the cryopump is operated, there is a remarkable difference when the chamber is evacuated by diffusion pump

only, and this seems attributable to the stmosphereic moisture. By this

figure it is confirmed that the evacuation system using cryopump can endure the high gas load, and especially it shows a good performance even when the large quantity of water vapor and other condensable vapors

outgassed from sample.

b . Sub-chamber

Fig. 7 shows the pumping characteristics of the sub-chamber. Curve

A is in the similar case to that of curve B in Pig. 6, while curve B is in the case where it is evacuated by cryopump after 24 hours of evacuation and degassing by sputter-ion pump.

In each case, the four cryosorption pumps are operated successively

one after another, and this method shows better performance that the method of using these four pumps simultaneously, and this caracteristics

are proved by theoretical analysis. The sputter-ion pump should be used

when the shroud is cooled by liquid nitrogen to diminish the quantity

of outgassing from the shrouds and sputter-ion pump itself.

—9 -10 The ultimate pressures a-e 2x10 and 2.7x10 Torr in the cases

of A and B respectively. These pressures are measured by nitrogencalibrated nude B-A gauge.

c . Pumping Speed of Cryopump

The method of evaluating or estimating the pumping speed of cryopump has been investigated and discussed by many researchers (1), but they are devoted to deal with the phenomena concerning the sticking

probability or condensation factor.

In such case where the cryopanel is surrounded by shrouds, we would like to emphasize as the most important factor which expresses the performance of cryopump: "Arrival Probability" (Pa). This is the ratio

of the total number of molecules arriving to the cryopanel after several collisions with the shroud, which is the heat shade of cryopanel in this

case, to the number of molecules which enter through the opening of the shroud. Considering this probability, the pumping speed of such cryopump is expressed as:

j S = M.F.A.Pa.Km = C.Pa.Km (1)

II vhere F and A are open area ratio and overall area of shroud respectively, (F.A is area of opening of shroud), M is number of molecules entering through open area of shroud in the molecular flow condition expressI ed in L/sec per unit area. Km is sticking probability of molecules 687
reaching onto the cryopanel. Then, M.F.A. = C is conductance of shroud

in L/sec.

Of course, it is very difficult to analyze and define this probability theoretically, therefore it was computed by Monte Carlo method

(2), as illustrated in Fig. 8. Briefly, the molecules entering through

the opening area of shroud collide with the interior surface of shroud

and reflected. Some of them escape out through the opening, some of them collide again with another spot of the surface, while the others

arrive to the cryopanel. 3,000 molecules entering through opening of the shroud from various directions are counted, and numbers of arriving molecules to the cryopanel are summed up to the 20th collision with the shroud surface.

The result shows that this probability is strongly influenced by the slight difference of shape and dimension of shroud, as shown in Fig.

9. In the main-chamber, there are three parts of shroud of L=300 mm, 600 mm and 1,200 mm. The total of pumping speed of this cryopump shaded

by shroud is shown in Table 5, using this "Arrival Probability" and equation (1).

4. PUMPING SPEED MEASUREMENT

The pumping speed of the cryopump of main-chamber is measured over

—8 —5 the pressure range of 7x10" -lxl0~ Torr by the method shown in Fig. 10. Dry nitrogen gas is introduced through a leak valve for the direction

of sticking to the shroud of the door. The pressure is measured by two gauges: nude type B-A gauge which is hung at the center of the chamber

and ordinary type B-A gauge attached to the wall of the chamber.

The pumping speed observed is shown in Fig. 11, in which line A denotes the value of 200,000 L/sec measured by the wall B-A gauge, while

line B denotes the value of 150,000 L/sec measured by the nude B-A gauge. There are many problems which require careful attention on the method of pressure measurement in the space which is surrounded by cold

wall, but in this study only the temperature dependence of the reading

of vacuum gauge is corrected.

5. ANALYSIS OF RESIDUAL GAS

The example of composition of residual gases in the sub-chamber when evacuated by the oil free pumping system is shown in Fig. 12, which

is analyzed by the Varian-MAT Type AMP-3 Mass-filter at the pressure

of 3xlO~^^Torr. The pattern shows that the principal residual gases are

H2O, N2, CI2 and CO2. Chlorine seems to have come from the solvent which

was used for cleaning the analyzer. Gases having mass-number of over 50 are very seldom, and this is the evidence of the merit of adopting the completely oil free pumping system. Also the pattern shows that there

is no noticeable leak from the chamber and inside structures.

Fig. 13 shows the mass pattern of residual gases when the cryopump

is being operated, while Fig. 14 shows the pattern of 10 minutes after the cryopump is stopped and the temperature of the cryopanel is raised up

to 50°K. These two patterns were obtained in the same experiment, and these 688
denote that the quantities of H^, He, Ne ' , and 0^ inclease after cryopump was stopped. This confirms that some innert gases are co-condensed together with condensing gases such as O2 and , and are released

when the temperature of cryopanel is warmed up. The composition of residual gases of chamber which is quickly

evacuated after exposure to the atmosphere as in this study is quite different from that of the ordinary ultra-high vacuum system which is evacuated with a long period of baking up.

6. SELF-CONTAMINATION

Self-contamination factor which is proposed by Dr. R. Chuan (3) is expressed as:

= (2)

In this study it can be presumed that 7c = 0.8, 7m = 0.5, and Ac/Am =

20 for main-chamber encolsing artificial satellite, then Cs = 0.006

is obtained. For the sub-chamber, "Zc - 0.8, ^m = 0.5, and Ac/Am = 100

for small testing sample, then Cs = 0.0012 is obtained.

The self-contamination factor should be kept in a very low value

to realize the true similar environment of space, and this is sufficiently achieved in this study by using large area of cryopump.

7. CONCLUSION

The newest Japanese space simulation facility is described. The pumpirg systems for main- and sub-chamber showed the distinctive features

of very large pumping speed and oil free "clean" system respectively.

These are certified by the residual gas analysis and by the concept of "self-contamination"

The new concept "Arrival Probability" to clearfy the mechanism of cryopump is introduced, and the procedures of designing cryopump are described using this probability.

8. ACKNOWLEDGEMENTS

The authors wish to acknowledge the suggestions and helps to this

study by Prof. Dr. K.Ohshima, Dr. K. Nakayama , and Dr. J. Asai.

REFERENCES

(1) . J.P.Dawson, A. V. Peck, 1962 Trans. 9th Natl. Vac. Symp. Am. Vac. Soc, 57 (1962)

(2) . Y.Akiyama, K. Nakayama, N.Saito, J. Vac. Soc. Japan. 14 (1969) 135.

(3) . R.L. Chuan, The Fluid Dynamics Aspects of Space Flight. 2, (1964) 41.

689
Table i. Dimensions of Chambers and Shrouds Main-chamber Sub-chamber

Chamber

Type

Inside Diameter Length Material

Surface Area Capacity

Horizontal Cylinder

1.8 m 4.9 m 18-8 Stainless Steel

30 m^ 12,600 liter Shroud (Cylindrical)

Type, Material

Outside Diameter

Inside Diameter

Length

Surface Area Shroud (Door Part)

Type, Material Diameter

Surface Area Cryopanel

Type, Material

Surface Area

Assembled from Extruded

Al Pipe with Fins, Epoxy

1.64 m 1.44 m 3.9m •

. ' ^ 89 :

Disk Assembled from the Same Al Pipe

1.75 m 4.5 m^

2 Lines of Al Pipe

r 2 6 m

60 cm

96 cm

4 m

300 liter

50 cm

40 cm 70 cm 6.3 m^

50 cm 0.33 m'

0.6 m'

Table 2. Performance of Refrigerator Main-chamber Sub-chamber Dummy

Expansion Engine

Inlet Pressure (ata) 27 31 28 Temp. (°K) 20 24 21

" Enthalpy (Kcal/Kg) 26.8 31.5 27.7 Outlet Pressure (ata) 1.25 1.74 1.58 Temp (°K) 8.5 11.0 10.5

" Enthalpy (Kcal/Kg) 13.4 16.7 15.5 Returning Pressure (ata) 1.25 1.74 1.58

Temp. (°K) 13.2 13.9 15.2 Enthalpy (Kcal/Kg) 19.4 20,4 21.9 Capacity

Cold Out-put (Kcal/hr) 272 291 293 Efficiency of E.E. {fo) 78 79 76 Heat Load (Kcal/hr) 94.0 94.5 129.1

690
Table 3. Pumping System

For the Main-chamber

15°K He Cryopump

Oil Diffusion Pump

Roots Blower

Rotary Pump

For the Sub-chamber

15°K He Cryopump

Sputter-ion Pump

Cryosorption Pump

Specification

Area 6 m^

36", Heater 30 KV Motor 7.5 KV

Motor 11 KV

2 Area 0.4 m

4x1 Kg. M-S 5

Pumping Speed

170,000 L/sec 37,000 L/sec 34,000 L/min

6,000 L/min

10,000 L/sec

400 L/sec

Table 4. Outgassing Rate from Components

Are^ Material Outgassing Rate Total Main-chamber (cm ) (TorrX/sec . cm ) (Torri/sec Interior Vail, Chamber 3.0x10^ 18-8 S.S. 2x10-10 6.0x10"^

Exterior, Shroud 2.9x10^ Al 1x10"^ 2.9x10-"^ Interior, " 5.8x10^ Al, Epoxy 2x10"^ 1.2x10"^

Cryopanel 6.1x10^^ Al 1x10"^ 6.1x10"^

Mirror 5.1x10^ Glass, Al 2x10"^ l.OxlO-"^ Gasket 1.0x10^ Viton-A 1x10"^ 1.0x10-"^ Sub-chamber

Interior Vail, Chamber 3.9x10^^ 18-8 S.S. 2x10-10 7.8x10"^

Exterior, Shroud 2.3x10^^ Al 1x10"^ 2.3x10"^

Interior, " 3.8x10^^ Al , Epoxy 2x10"^ 7.6x10-"^ Cryopanel 6.3x10^ Al 1x10"^ 6.3x10"^

Gasket 3.0x10 Viton-A 1x10"^ 3.0x10"^

Table 5. Pumping Speed of Cryopump

Main-chamber

Dimension Pa F.A (cm^) C (L/sec) Pumping S. (L/sec)

L = 300 0.622 48,600 574,200 356,000

L = 600 0.306 16,200 191,000 58,500

L = 1,200 0.139 37,800 328,000 62.000

Total 477,000

Sub-chamber Dimension Pa F.A (cm^) C (L/sec) Pumping S. (L/sec)

c.f. Fig 9. 0.661 4,158 49,060 32,500

691
4900

3700

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1. Structure of

Main- and SubChamber .

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694
Fig. 6 Pump-dovn Curve Fig. 7 Pump-do\m Curve

Ma in-chamber Sub-chamber
Fig. 8 Monte Carlo Method

Fig. 10 Pumping Speed Measurement

696
^ 1x10^ Load Gas: Dry

Temp, of Cryopanel: 10°K

10 ^ 10 10^ Pressure (Torr)

10 10

Fig. 11 Pumping Speed of Cryopump for

Main-chamber

Fig. 12 Gas Analysis of Residual Gases in

Sub-chamber

697
P=:lxlO ^Torr

Pig. 13 Residual Gas of Pig. 14 Residual Gas of

Sub-chamber, Cryo- Sub-chamber, Cryopump is operated pump is stoped

698
Paper No. ^7

SIMULATION OF SURFACE "A^OSPHERE INTERACTION

PROCESSES FOR MARS

K. M, Foreman"^

REFERENCE: Foreman, K. M. , "Simulation of SurfaceAtmosphere Interaction Processes for Mars/' ASTM/

lES/AIAA Space Simulation Conference, 14-16 September 1970.

ABSTRACT: Landings of unmanned spacecraft on Mars

can be affected by the dynamics of the lower planetary atmosphere. In an effort to understand the

convective motion and effect of surface topography

and soil on local atmospheric motion, a laboratoryscale simulator has been designed and built to ex- amine parameters of significance systematically.

Scaling law relationships were used in sizing the

facility and appropriate test conditions within the

terrestrial gravitational environment. This facility is a 4-foot diameter by 8 foot high stainless

steel chamber capable of maintaining test conditions

to about 10 microns absolute pressure and gas temperatures between -85°cb and +65°C^. The planetary surface is simulated by an electrically heated

aluminum surface controlled to within ^O.^C^ of

the initial set conditions.

Measurements of test gas temperature above the

simulated planetary surface are made with a specially designed traversing mechanism carrying thin

'"Research Engineer, Grumman Aerospace Corporation

Bethpage, New York 11714

699
film and thermistor sensors. Temperature profiles

above the surface are similar to theoretical calculations for equatorial Mars regions at midmorning,

if the lower Martian atmosphere is assumed to have an eddy viscosity of about 5 x 10^ cm^/s. Additional data show that temperature gradients increase near the surface as the gas pressure is decreased

below atmospheric level, and that the degree of turbulence (as evidenced by temperature fluctuations)

appears to diminish with lower pressure.

Exploratory tests of thin film sensors as wind

speed indicators at simulated Mars surface pressure

of 5mb show possible difficulty in differentiating

velocity from temperature at wind speeds less than

several m/ s magnitude. However, thin films appear

quite responsive as temperature sensors to pressures

on the order of less than 1 mm Hg A.

KEY WORDS: Mars atmosphere exploration, simulation

techniques, research facility, convective processes,

space research, planetary atmosphere -surface interaction, planetary meteorology

INTRODUCTION

Man has already planted his automated laboratories, and his person, on an extraterrestrial body:

the moon. And within the decade, the Martian surface will certainly be investigated by automated

spacecraft. This involves a descent through the

Martian atmosphere, and during the final, crucial

moments just before touchdown the spacecraft will be

subject to the dynamic processes of the Martian

boundary layer.

Martian planetary meteorology has been the subject of many theoretical studies the past few years.

Among the notable contributions are the numerical

experiments of Leovy and Mintz^ and the analyses and

700
computations of Gierasch and Goody, ^ Ohring and

Mar iano

^ and Neubauer .

In all these attempts to explain the relativelyfew astronomical observations of Mars and predict

operational environments with reliability sufficient

for the spacecraft builders, there has been a free

projection of our Earth experience onto what passes

as meteorological data meaningful for Mars.

What is the nature of this information? We are fairly sure that Mars has a very high CO2 "content

atmosphere (perhaps ^ 100 percent) , with a surface

pressure somewhere between 3 and 12 mb. Its

gravity is about 1/3 the Earth's and it has characteristic light and dark regions, some of the latter

looking like oases within the light regions. The

surface temperature varies latitudinally, diurnally,

and seasonally between 140 °Kb and about 300

From a flyby spacecraft, several thousand km from

the planet, the surface, appears extraordinarily

cratered and rough, on a scale greater than what we

know exists on the moon. Mars exhibits dynamic meteorology if only by

evidence of visible cloud formations and motions.

Some clouds are white, others blue, still others

yellow. With almost negligible moisture, the Martian atmosphere's dynamics are most probably unlike

the Earth's and may be considered to be closely

linked to surface heating and radiative transfer to

CO2 molecules

Neubauer^ and Gierasch and Goody^ have noted

the potentially strong radiative coupling that can exist because of the 15 micron absorption band of

CO2 and the corresponding IR wavelength emission

of a 250 soil, acting as a blackbody.

Clearly, the meteorological theories are approximate, and short of in situ data to be obtained (at this writing) in 1976, they remain some- what conflicting and indecisive. We need some perspective to decide which theories and assumptions

should be accepted, modified, or discarded.

701
Simulation of planetary processes in a laboratory is a convenient way to examine the various

hypotheses, although there are inherent limitations

to scaled physical experiments, such as physical

size of instrumentation and its effect on the process being simulated. With this in mind, we set out

to examine a few concepts of the Mars surface -atmosphere interaction problem. To bring the facility

to manageable and economic dimensions, we limited

our inquiry to the planetary boundary layer where

geos trophic effects could reasonably be ignored.

However, to create a qua si -Martian boundary layer

environment we needed an enclosure that could maintain an approximately 5 mb CO2 atmosphere.

The solution was an atmospheric boundary layer

simulator. Previous efforts^ W have been directed

primarily toward simulating Mars d5niamic pressure

and properties of developed dust and sand storms

relative to their effects on spacecraft materials.

However, while composition, pressure, velocity, and

temperature of the Martian atmosphere were duplicated to the extent believed to be represented by

Mariner flyby data and Earth observations, there is no evidence that the previous experimenters were

concerned with complete dynamic simulation of the

atmosphere (including scale and gravitational effects) nor with the best estimates of the nature of

the Martian surface (soil and topography)

Initially we are concerned with convective processes in a still -wind field. This involves the

convection initiation mechanism where radiative

transfer may not be negligible and the participation

of particulate soil in the vertical transport of

free convection or dust-devil type formations; the

latter is believed to be the source of extensive

yellow cloud formations that, according to the past

90 years of detailed observations, have an occurrence frequency comparable to hurricanes

This paper will cover the design of the test

facility, the operating procedure used to conduct

tests, and some preliminary results of still -air

convective simulation.

702
Equipment Design

Scaling laws make possible the dimensioning of

the simulator; we considered the nondimensional

Grashof and Rayleigh numbers for free convection.

Incidentally, since two-phase flow (i.e., solids and

gas) may be involved in the dust-devil formation, we

can also use the Grashof number in the form where

the buoyancy is produced by local density relative

to the average density of the surroundings, Ap/p,

instead of AT/T, which relates the relative temperature difference.

Applying the Grashof number to likely conditions on Mars and appropriate test conditions in an Earth-bound simulator (ES) we have the relative

dimensions given by

where L is characteristic length; g, gravitational constant; kinematic viscosity; and T,

the temperature; from this we can size our facility.

For simulating a nominal 5 mb, 200°Kp , 100

percent CO2 atmosphere on Mars, we find we could

use a 1000 mb CO2 test gas at room temperature in

a 1 percent scale simulator, based on purely molecular thermal properties of the planetary and simulator gas. (More will be said on this point later in

the paper.) However, for other relevant investigations, such as the thermal conductivity of particulate soils at low atmospheric pressure 5 mb)

there is also need for a vacuum capability to duplicate appropriate Martian surface conditions.

Among the interesting Martian phenomena are the

yellow cloud formations. Initiation of these clouds

may be a surface -atmosphere process involving advection and lifting to altitude — by saltation pickup

703
or turbulence — of top soil material into convective columns similar to dust devils. Ives's^ study

of terrestrial dust devils indicates that a likely

minimum AT/T of about 0.12 is needed to create

dust devils with characteristic diameters on the

order of 20 meters. In a 1 percent scale simulator,

a AT in excess of about 40 should be appropriate if unstable stratification can be created.

Using Rayleigh number criteria for one fixed and one

free surface (the planetary atmosphere case) , centimeter-sized layer depths should yield convective

flow for AT of 5°C^ or more.

On the strength of these studies and considerations, we designed and fabricated a 4 ft diameter

by 8 ft high cylinder capable of maintaining test

conditions of 10|i.m absolute pressure as well as

2 atm positive pressure over an operating temperature range from -85 to +65°C^. The tank (see

Fig. 1) consists of three interchangeable spool sections, each 2 ft high, and two dished end sections,

all bolted together at matching flanges. To eliminate chemical corrosion problems, all major structural elements were made from type 304-SA 240 stainless steel. The spool sections were rolled and

welded from 3/16 in. thick material and the end

sections were made from standard 48-in. diameter

dished heads of 1/4 in. thickness. The sectional

approach was taken to provide maximum flexibility in

setting up tests as well as to permit future extensions to greater heights, if desired. In each of

the section flanges "0" ring installations prevent

leakage even at micrometer -range vacuum despite the

interchangeability feature. Two sections have

18 in. manholes that allow access to the interior

for setting up test apparatus and instrumentation.

These sections also have 4 -in. diameter viewing

ports with quartz windows for optical studies.

These can be interchanged with solid flanges to

carry special instrumentation or actuators. The

center section has three 3-in. diameter ports to

accommodate probe traversing apparatus and provide a connection to the pumping and gas feed system. An

auxiliary valve also is located in this section to

bleed off excess pressure rapidly or to permit

704
various gases, such as C02^ neon, or custom mixtures of possible atmospheric gases to be introduced.

At each of the end sections is an 8 -in. diameter

flanged port hole. These permit electrical vacuum

feed-through connectors to be installed for supplying

power to interior equipment or carrying instrumentation signals from sensors to exterior signal conditioning and recording equipment. An automatic pressure relief valve is installed at the top head to

prevent excessive overpressures from being inadvertently introduced.

As currently used, the tank is supported upright by 4 legs welded to the bottom dish, and

bolted to the concrete floor of the laboratory;

there is six in. clearance between the bottom 8 in.

flange and the floor. Another possible orientation

for the tank is horizontal, as the test section for

a closed-loop 8-in. diameter duct that is connected to the flanges at each end of the chamber. This

configuration could be used to simulate cross-wind

profiles arising from long fetches.

The chamber is connected by a stainless steel

manifold to a Hyvac 45 vacuum pump. Various valves

and bourdon type pressure gauges permit control of

the pressure to particular set values in the 0.1 mm Hg Absolute to 1 atm pressure range; a McLeod

gauge measures pressures between 10 micrometers and

5 mm Hg Absolute. One bourdon pressure gauge can be

used either for vacuum to 10 mm Hg Absolute or positive pressures up to 2 atm; another gauge overlaps the vacuum pressLre range of 0.1 to 20 mm Hg

Absolute

Experiment Procedure

The Martian surface is simulated by a 3 -in.

diameter area of an electrically heated hot plate

mounted centrally within and at the bottom of the

atmosphere simulator (see Fig. 2) . By means of an externally located temperature controller connected

to a temperature sensor attached to the plate, the

test surface temperature is controlled to within

705
0.4°C^ of the desired average value. A 1/2 In.

thick aluminum plate was intended to average out

local hot spots caused by proximity of the underlying heating element. The degree of success in

achieving surface temperature uniformity is evident

from the sample distribution of average temperature

shown by Fig. 3. The major profiles of air temperature above the heated surface were taken at 90 degrees (i.e., left of center) and. 270 degrees (i.e.,

right of center) orientation angles. The mean positional variation from set temperature across this

diameter is 0.7°C^, This hot plate performance is

considered satisfactory for experiments of qualitative objectives. For higher precision data, some modifications are planned to create more uniform

surface temperature profiles. It may be of interest

to note that the temperature profile of the initial

surface simulator configuration exhibits similarity

to a planetary surface heated by insolation, at

about midmorning, with the "sun" rising at an orientation angle of about 180 degrees in the experiment's set of coordinates.

The hot plate surface outside of the central

3-in. diameter test area is insulated from the test

gas by a double layer of asbestos sheet, separated

by a 3/8 in, air gap. To prevent edge leakage of

heat, a vertical metal radiation shield surrounds

the hot plate at the outer edge of the asbestos insulation (not shown in Fig. 2).

Experimental Work

Measurement of gas temperature and velocity

profiles above the heated surface is achieved with

miniature quartz -coated thin film sensors (see

Fig. 4). In addition, a less time -responsive thermistor temperature probe is mounted 7/16 in. to the

right of the hot film sensor, which are attached to a traversing mechanism that allows very fine vertical positioning (viz., within 1/200 in.) and horizontal positioning to within 1/16 in. The vertical

travel of the instrument carriage is controlled by a constant speed (Motomatic Model E-500 MG) motor

706
(speed regulation better than 2 percent) with continuous control variability possible up to about

2.5 cm/s linear motion. The motor is geared to a rotating drive tube (see Fig. 5) that runs into the

chamber through an "0" ring seal. A miter gear at

the end of this shaft engages a matching gear on a vertical threaded rod inside the chamber and causes

the instrument carriage to move up or down, depending on the motion mode activated at the controller.

Limit switches prevent the carriage from moving

beyond desired travel limits. Dynamic braking by

the drive motor stops motion within 1/400 in.

The thin film sensors are cylindrical, with an effective length of 0.080 in. and a diameter of

0.006 in. As employed in our tests, these sensors

can be used either to indicate local gas temperature

or velocity fluctuations caused by the convective

motion above the hot plate. As a temperature sensor,

the thin film acts as a very fast response linearly

varying resistance thermometer. Connected to a resistance bridge and output amplifier, the analog

signal output gives virtually instantaneous temperature values. Velocity at the same location is indicated by switching the sensor's signal processing

electronics to a velocity mode. Here, the film is

maintained at a constant temperature above the local

gas (approximately 195 greater) by a bridge

servo loop that responds to the increased heat

transfer from the film caused by gas motion. The

output signal is proportional approximately to the

1/2 power of velocity, i.e.. King's law

= [A + B(pV)*](Tg - T^)

where: E is output voltage; A and B are em- pirical constants; p is gas density; and V is

gas velocity. However, there is a need for corrections at low gas velocities (i.e., below approximately 20 cm/s), see Ref. 9.

In the case where the temperature and velocity

fluctuations occur simultaneously, the temperature

effect can be isolated from the velocity effect by

707
running two such nearby hot films at somewhat different overheat temperatures, Tg^ and "^82' greater than the average gas mass. Then, the difference in the two sensor voltage outputs can be re lated to the velocity term in the expression

- E2 = [A + B(pV)*](Tg^ - Tg^)

if the gas density is considered essentially con- stant over the relatively small differences in operating temperatures and their fluctuations.

Discussion

Some interesting results of preliminary temperature surveys across and above the simulated

heated surface are presented in the following discussion. A more detailed and complete analysis is

planned for a future paper. Figure 6 shows vertical

profiles of temperature above a 120 heated surface for two pressures of air and two temperature

instrumentation techniques. The profile at one atmosphere pressure gives evidence of greater con- vective turbulence than at 15 mm Hg Absolute. The

greater heat energy transfer by convection at the

higher pressure results in somewhat lower average

temperature in the gas above the surface . This is

further supported by Fig. 7 which shows the relative

variation of temperature with decreased pressure at

0.125- in. above the center of the circular heat

source. It is also interesting to compare in Fig. 6

the miniature fast-response thin film sensors to the

larger slower responding thermistor probes. It

should be noted that the thermistor is about 7/16 in.

to the right of the hot film; the surface directly

below it is at a negligibly (~ 0.25°cb) higher

surface temperature than at the center. However,

there is evidence with both sensors of the highly

variable temperature field within the first two

inches above the heat source, and the large temperature gradient within the first 1/8 to 1/4 in.

above the bare plate.

708
When CO2 is used as the test gas instead of

air, it would be expected that similar results would

be obtained, except that the profile surveys should

indicate slightly lower temperatures in CO2 because its thermal conductivity is somewhat less than

for air. This, in fact, is observed from the test

data which shows CO2 temperatures to be 2 to 3°C^

lower than for air at comparable positions above the

heated surface. Thus, qualitatively, the air data

presented here is representative for full Martian

simulation.

The significance of these results to the planetary problem can be estimated by inspection of

Fig. 8. There is a strong similarity between the

average temperature profile, measured at 1 atm

pressure, in the simulator with the detailed calculations made by Gierasch and Goody2 for the equatorial regions of Mars at 10 A.M. Also shown for

comparison are late afternoon (4 P.M.) calculations

for Mars and the lab measurements at 15 mm Hg Absolute pressure.

The characteristic length for the similar profiles in the laboratory and for Mars calculations

are in the ratio of about 10"^. On the basis of

the similarity criteria mentioned above in which

only molecular gas properties were used, the Grashof

number in the laboratory would be also about the same order of magnitude smaller than the planetary

condition. However, Gierasch and Goody^ used an eddy viscosity value on the order of 10^ cm^/s

(compared to the molecular viscosity value of the

order 10"i cm^/s) to compute the case for equatorial Mars at 10 A.M. An eddy viscosity of 10^

cm2/s is, perhaps, appropriate for an assumption of

very strong turbulence. However, if a more representative terrestrial value for eddy viscosity on

the order of 5 x 10^ cm^/s^^^^ is introduced for

the viscosity term in the Grashof number, we achieve

practically exact thermodynamic similarity, in the

laboratory, with the Mars conditions. Thus, 'Mars

inay not exhibit as strong a turbulent atmosphere as has been assumed. These qualitatively similar re709
suits, of course, need further physical simulation

such as: CO2 atmosphere, particulate surface cov- ering (possibly granular llmonlte) and more uniformly distributed heat source. Nevertheless, the

Initial results give some perspective on the possible lower atmospheric motions, especially turbulence,

that can be caused by convection. Comparison of the

reproductions of two strip chart readouts of the hot

film resistance thermometer, one at 1 atmosphere

air pressure (Fig. 9) and the other at 5 mm Hg Absolute (Fig. 10), further emphasizes the strong temperature fluctuations near the heat source when

close similarity (I.e., at one atmosphere pressure)

to the planet Is set up, as opposed to near-constant

temperature at low pressure.

Velocity data Is not presented here because the

available In situ linear calibration speed range

Is too low with the existing constant speed motor

drive. To remedy this restriction on test speeds,

and extend the range to more significant speeds

(I.e., at least an order of magnitude greater), we

Intend to Install a linear motor on the traversing

carriage to move the attached probes at shorter

strokes but faster speeds.

Mars Lander Instrumentation

The hot film Instrumentation used In our simulator Is of the type that has been considered for

wind measurements on Mars landers. Aside from the

question of operational survivability In the total

mission environment of a Viking -type lander, there

has been some question about the suitability of this

technique at Mars surface pressure.

To explore this area, we conducted a test at

5 mb pressure In the simulator. Two hot film sensors were mounted above a slowly Increasing temperature heat source; one operated as a temperature sensor and the other as a velocity sensor at 1.5

overheat ratio. We found the temperature application to be highly responsive to temperature changes

and we believe It could be successfully employed In

710
Martian pressures as low as 1 mb as a thermometer.

However^ we have had difficulty in discerning low

speed velocity from temperature changes and circuit

noise. So, unless improvements in signal processing

equipment are introduced, wind speed resolution below several m/s appears questionable.

Future Tests

In addition to the bare -surface temperature

surveys, we plan to conduct similar tests with the

simulated planetary surface covered with particulate

limonite to simulate the Martian soil. This hydrated iron oxide mineral has been found to exhibit

thermal and optical properties that best match Martian observations.^! There is a need for updated

thermal transfer data on particulates under subatmospheric gas pressure conditions.

Another future test of Martian relevance involves cooling the test gas with liquid nitrogen

shrouds to induce formation of solid CO2 aerosols

above the test area. It is believed that these submicrometer-sized aerosols can best explain data from

experiments on -board Mariners 6 and 7^2 and have

also been indicated by interpretation of Mie scattering calculations to explain Mars polar ime trie observations . 13

Conclusions

In anticipation of future spacecraft landings

on Mars, a new facility has been designed and con- structed to give thermodynamic simulation of atmospheric processes in the planetary boundary layer.

Initial qualitative tests of free convection temperature profiles in a still wind field show distinct similarity to detailed theoretical calculations for equatorial Mars regions at midmorning if an eddy viscosity on the order of 5 x 10^ cm^/s is

assumed for the planet. From these results it appears that the lower two km of the Martian atmosphere could prove quite turbulent under the condi711
tions described. Hot film sensors appear sufficiently responsive to indicate rapid temperature

fluctuations in a turbulence field at Mars surface

pressure conditions. However, unless extreme care

is taken in signal conditioning and processing, one may anticipate difficulty in using hot film anemome

try for wind speeds below tens of m/s.

The author thanks Mr. Frank Neumann for his

patient assistance in conducting the tests and instituting equipment modifications where necessary.

The design and checkout of the simulator were made

possible by the energy of Messrs. Suey Jue, H. Rodenburg, V. Taibbi, and J. Wagner. Finally, we acknowledge the interest and support of Dr. N. Milford and the sponsorship of the Grumman Aerospace

Corporation as part of its IR & D Program.

References

[1] Leovy, C. B. and Mintz, Y., "A Numerical General Circulation Experiment for the Atmosphere

of Mars," Rand Corporation Memorandum RM-5110-

NASA, December 1966

[2] Gierasch P. and Goody, R., "A Study of the

Thermal and D3mamical Structure of the Martian

Lower Atmosphere," Planetary Space Science , Vol. 16, 1968, pp. 615-646.

[3] Ohring, G. and Mariano, J., "Seasonal and Latitudinal Variations of the Average Surface Temperature and Vertical Temperature Profile on Mars," J. Atmospheric Sciences , Vol. 25, No. 5,

September 1968, pp. 673-681.

[4] Neubauer, F. M. , "Thermal Convection in the

Martian Atmosphere," J. Geophysical Research , Vol. 71, No. 10, May 15, 1966, pp. 2419-2425.

[5] Gierasch, P. and Goody, R., "An Approximate

Calculation of Radiative Heating and Radiative

Equilibrium in the Martian Atmosphere," Plane - tary Space Science , Vol. 15, 1967, pp. 1465-

1477.

712
[6] Hertzler, R. G., Wang, E. S. J., and Wilbers,

0. J., "Development of a Martian Environmental

Simulation Facility/' Proceedings of the AIAA/

lES/ASTM Space Simulation Conference , Houston,

Texas, September 7-9, 1966, pp. 172-178.

[7] Dyhouse, G. R., "Simulated Martian Sand and

Dust Storms and Effects on Spacecraft Coatings,"

Proceedings of the ASTM/IES/AIAA Space Simula - tion Conference , Philadelphia, Pa., September

11-13, 1967, pp. 97-103.

[8] Ives, R. L., "Behavior of Dust Devils," Bulle - tin American Meteorological Society , Vol. 28,

1947, pp. 168-174.

[9] Almquist, P. and Legnth, E., "The Hot Wire

Anemometer at Low Air Velocities," DISA In - formation , No. 2, July 1965, p. 3.

[10] Hanna, S. R., "A Method of Estimating Vertical

Eddy Transport in the Planetary Boundary Layer

Using Characteristics of the Vertical Velocity

Spectrum," J. Atmospheric Sciences , Vol. 25,

No. 6, November 1968 pp. 1026-1033.

[11] Egan, W., "Polarimetric and Photometric Simulation of the Martian Surface," Icarus , Vol. 10,

1969, p. 223.

[12] Herr, K. C. and Pimentel, G. C, "Evidence for

Solid Carbon Dioxide in the Upper Atmosphere of

Mars," Science, Vol. 167, No. 3914, January 2,

1970, pp. 47-49.

[13] Egan, W. G. and Foreman, K. M. , "Mie Scattering and the Martian Atmosphere," to be published in Proceedings of the lAU S3niiposium 40 , Marfa, Texas, October 27-31, 1969 , (also Grum- man Research Department Report RM-461J, November 1969) .

713
714
715
716
717
toQ

If) CD CO

J>, • HoQ

<D

•H

tJU +Jo

u: •

0) ^^ cr> o CT3 c

CO -P o CO X)

r—

-p

CO o c

(1) >it1c

-p

»5 o

•p >• H o Q CDc o •H • H c

-P o £ -P o

CD o O

a<iH

CDo a •H

CO -P

f-i +>c > CD +^

to C c CD oo• lO

CD

f-t CT) •H

718
Fig= 6 Temperature Profile Above a 120cSurfaceComparison with Gas Pressure and Sensor Technique

719
1.6 r N

^'^ ' 100 200 300 400 500 600 700 800 Pressure, mm Hg A

Figo 7 Variation of Air Temperature with Pressure at 0. 125 irio Above a 120 C

Bare Surface - Comparison of Two Measurement Techniques

720
Fig. 8 Comparison of Temperature Profiles in Atmosphere Simulator with Mars Calculations (Ref 2)

721
THERMISTOR TEMPERATURE

PROBE OUTPUT

0.01 VOLT

55.1°C'

TIME signal: 1 DIV = 1 SECFIG. 9 TIME VARIATION OF TEMPERATURE AT

1/8 INCH ABOVE CENTER

OF 120°C BARE HEAT SURFACE,

PRESSURE = 1 ATMOSPHERE

HOT FILM TEMPERATURE

SIGNAL

HOT FILM TEMPERATURE

SIGNAL

0.25 VOLT

1.0 VOLT

722
0.01 VOLT

THERMISTOR TEMPERATURE

PROBE OUTPUT

86.8°C

TIME signal! 1DIV=1SECTT

4-4-+

FIG. 10 TIME VARIATION OF TEMPERATURE AT

1/8 INCH ABOVE

CENTER OF 120°C BARE HEAT SOURCE, PRESSURE = 5 MB

HOT FILM TEMPERATURE

SIGNAL

0.25 VOLT

(5 VOLT SUPPRESSION)

HOT FILM TEMPERATURE

SIGNAL

1J

1

— H

—

t + 'OLT H

— —

1

I1 1

i —

T- s r h—

1 H

723

Paper No. U6

OPERATION OF A LARGE THERMAL VACmJM CHAMBER

AT MARTIAN PRESSURE LEVELS

T. Buna

Staff Engineer, Thermal Technology Section

J. R. Ratliff

Senior Engineer, Thermal Technology Section

Martin Marietta Corporation, Denver, Colorado

ABSTRACT: A test program was conducted to evaluate techniques

of insulating the cold shroud of a large thermal -vacuum chamber

from its outer shell, in order to prevent excessive cooling of

the latter during Mars environmental simulation tests. A fluidized

insulation concept using polystyrene flakes in the shroud/shell

interspaces was found to best meet both thermal and operational

requirements. With 140 R (77,8 K) shroud, convective transport

through the insulation was negligible up to 20 mm Kg chamber

pressure, and did not exceed permissible levels up to 50 mm Kg. At the higher pressure levels local heat transfer through the

insulation was found to be of a cyclic nature. Thermal data

are presented in the form of Nusselt-Rayleigh type correlations.

Unique techniques devised for rapid installation and removal of

the insulation are described; chamber shell temperature monitoring techniques (including infrared scanning) are evaluated;

and the pumpdown characteristics of the insulated chamber are

established.

KEY WORDS: Mars environmental simulation, thermal -vacuum

chambers, heat transmission, thermal insulation, infrared

scanning, evacuating, outgassingo

NOMENCLATURE

c constant in heat transfer correlation

D chamber diameter

E radiation interchange factor

f function symbol

g acceleration due to gravity

h convective heat transfer coefficient

k thermal conductivity

Nu Nusselt Number

q heat transfer rate per unit area Ra Rayleigh Number

T temperature, deg abs

K thermal diffusivity

V kinematic viscosity

725
p = density

a = Stefan-Boltzman constant

Subscripts:

av average, pertaining to a convective zone

L local

C convection

g gas m mean value .

, 0 shroud R radiation

S shell

oo ambient atmosphere

I. INTRODUCTION

It has been shown (1,2)''' that the thermal extremes associated with windless conditions on the surface of Mars can be

simulated in conventional thermal -vacuum chambers. The method

requires that the chamber be operated at Martian pressure levels

(2 to 20 mm Hg) with cold-shroud temperatures simulating the

radiation heat sink of the Martian sky (3). With proper test margins the latter may vary between 140 R (77,8 K) and 400 R

(222 K) depending on the environmental conditions simulated.

There is one predominant problem associated with maintaining

these conditions in a thermal-vacuum chamber originally designed

for space environmental simulation: convective heat transfer

between the outer shell of the chamber and the nitrogen-cooled

shroud may cause the shell to cool below permissible levels,

with potential catastrophic failure of the chamber's structural

members. In addition, the increased (as compared to high-vacuuu

Operation) external heat loads may overload the temperature

control system of the shroud.

Outside of complete redesign and extensive modifications of

the chamber, there are two possible approaches to minimizing the

jshell-cooling problem: external heating, and internal insulation.

The subject of the present paper is the latter approach as applied to a large (29 ft diameter by 45 ft high) thermal-vacuuiii

chamber intended for use in support of the early Mars programs,

^n order to preserve the high-vacuum capability of the chamber.

It was required that the insulation be removable and its out- gassing characteristics be compatible with chamber cleanliness

Requirements and available pumping capabilities. The temporary

character of the insulation system introduced operational problems

competing in significance with the thermal requirements: rapid

Installation and removal techniques had to be devised that were compatible with typical planetary program schedules; and means

Tih& msoh^x^ in pArfinthssjeis r_efer to_ the LLsJL of xfefereuces

appended to this paper.

726
had to be found to continuously monitor the thermal performance

of the insulation once installed and in operation. Thus, the

operational and thermal aspects played equal roles in the selection and evaluation of candidate insulation concepts.

Section II of the paper outlines the thermal analysis of

the chamber shell/shroud system used for the purpose of providing

a framework for planning and interpreting the test program.

Section III discusses the test program entailing three phases:

A) Full-scale problem evaluation; B) Reduced-scale development;

and C) Full-scale performance verification. It is the results

of the Series C tests upon which the case for the fluidized insulation concept rests. Section IV is an attempt at generalization

of the results in the form of Nusselt-Rayleigh type correlations.

Due to the limited nature of the supporting data, caution should

be exercised when applying these correlations to conditions other

than those of this program.

II. THERMAL CHARACTERISTICS OF THE SHELL/SHROUD SYSTEM A „ thermal -vacuum chamber, when operated with a cold shroud

in the presence of an internal atmosphere, represents a thermal

system too complex to be amenable to purely analytical treatment,

hence an empirical approach was considered mandatory in the present investigation. As an aid to planning and interpreting of the

experimental program, it was found instructive to formulate some preconceived ideas about the convective processes in the chamber,

and to examine them in the light of similarity theory. Such an analysis is presented in this subsection.

The Uninsulated Chamber

A plausible model of internal and external convective patterns is depicted on Figure 1. The convective flow field between

the shroud and the outer shell is subdivided into three zones, corresponding to the interspaces of the bottom dome, cylindrical

section, and upper dome, respectively. A fourth zone is shown

to represent external (ambient) convection, interfacing with all

three of the internal zones. The model further assumes that the

effect of coupling between adjacent internal zones is negligible

except in the vicinity of the zonal interfaces, and that local

perturbations due to internal and external structures, penetrations, and liquid nitrogen lines, have a negligible effect on the overall convective patterns, although they may drastically

influence local temperatures. These assumptions can be tested by

experiment, as it will become apparent subsequently.

Of the three internal zones, the bottom dome interspace has

the conditions most favorable for convective motion to develop:

unlike Zone III, the Zone I-temperature gradients are unstable

727
(cold shroud on top, warm shell on bottom), their vertical com- ponents being considerably larger than in Zone II, Furthermore, Zone I has a more favorable geometry (shorter flow path lengths)

than has Zone II

Following Batchelor (4), the convective motion in Zone II is expected to be of the boundary layer type, with upward flow

adjacent to the shell, and downward flow near the shorud. In the laminar range, the two boundary layers are contiguous, and

horizontal convective motion can only develop near the two ends,

i.e., at points C and E on Figure 1, whereas heat transfer within

the central portion of the zone (Point D) is by conduction. At higher Rayleigh numbers, the two boundary layers become separated

by a turbulent core, and convection is similar to that obtained

with single vertical surfaces.

The upper dome interspace has the conditions least favorable

for convection because of its stable overall temperature gradient

(warm dome on top, cold shroud on bottom) » Convective motion can nevertheless, develop near the periphery of the dome (point F on Figure 1) where the dome intersects the shroud at an angle.

The above considerations indicate that, at moderate chamber

pressures, the most likely areas affected by convective cooling

are: the bottom dome (in its entirety), the upper and lower ends

of the cylindrical portion of the shell, the periphery of the

upper dome, and potential cold spots due to structural penetrations or nitrogen lines. It is further apparent, that with the

possible exception of highly localized spots, the coldest temper- atures will occur within Zone I of the chamber, to which our attention will be focused in the remainder of this subsectioHo

Dimensional analysis indicates that the heat transfer coefficient depends on the Rayleigh number, the Prandtl number, ap- propriate length ratios, and temperature ratios. Due to the re- stricted nature of the present investigation, the last three of

these may be dropped from further consideration: the Prandtl number, because of restriction to gaseous media; the length ratios,

because of limitation of the analysis to geometrically similar

domes; and the temperature ratios, in view of the fact that both

the shroud and shell temperatures were allowed to vary but slight

ly during this program. Thus, we may expect correlations of the

form: Nu

av

Cav

f(Ra) (1)

Nu^ f(Ra) (2)

where, by definition

Nu = h D/k

av av Ig

(3)

728
Nu = h D/k (4) L L g Ra = g(T -T )D"^/k V T (5) av o m

^ Note that the Rayleigh number is a characteristic parameter of

the flow field as a whole within a zone, whereas the Nusselt

number may be used to represent either average or local conditions within a zoneo Equations (1) and (2) imply that: 1) for all thermally similar bottom dome interspaces the function f and

the constant C^^ are the same; and 2) within a single dome, the

ratio C^/C^^ is a function of location only, independent of Ra. When comparing data from two different chambers, these statements may be expected to apply approximately at those locations on the

shell sufficiently removed from anomalies due to penetrations or lines, and for geometrically similar shells that are sufficiently

thin so that one-dimensional conduction may be assumed. This

proved to be the case in the present investigation, as it will be

discussed in Section IV. The Heat Balance Equation

The determination of the Nusselt numbers from the test data

requires that the heat balance equation of the shell be solved

for the heat transfer coefficient. The procedure used in this

program is outlined below.

Referring to Figure 1, and assuming one-dimensional steady

state heat transfer, the heat balance of an elemental area dA of

the shell may be written as

q +q =q +q =q (6) ^R,oo ^C,oo ^R,o ^C,o ^net ^ ^ where the last term on the right hand side represents the net

heat load to the shroud. Furthermore,

= (7 E (T ^-T ^) (7) ^R,00 OO OD s

q = h (T -T ) (9) ^C,oo OO^ OO s^ ^ ^

q^ = h (T -T ) (10) ^C,o o ^ s o^ In equations (7-10), the temperatures are supplied by the

test data: E and E can be calculated from the surface optical

OO o 729
dependent variable, whereas hoo can be estimated by the following procedure. Following McAdams (5), the average heat transfer

coefficient in air at ordinary temperatures and atmospheric pres- sure, and for the turbulent regime, may be expressed in the form

h = C (T -T )^^^ (11)

oo , av oo , av oo s ^

^ Consequently, using a rationale similar to that in conjunction

with Equations (1) and (2), we may assume, that:

h = C , (T -T )^/^ (12)

oo ,L oo ,L oo s where Cqq l

-^^ ^ function of location only, Coo,L ^® determined experimentally for each location of interest from data

obtained with the chamber pumped down to high vacuum. In this

case, o ~ ^> and, by substituting Equations (7,8,9 and 12)

into (6), the heat balance equation may be solved for C^^ -j^ The Insulated Chamber

The application of similarity relations to the insulated

chamber would be of limited practical use, since thermal scaling

of insulations is notoriously difficult, and, if successful, would

probably result in a loss of identity of the insulation scheme

from an operational standpoint. Since the principal aim in insulating the chamber is to suppress convection within the shroud/

shell interspaces, the relative insensitivity of the shell tem- peratures to increasing chamber pressures may be taken as the

prinicpal measure of performance for any given insulation concepts

To permit direct comparison with the uninsulated chamber,

this performance criterion can be expressed in the form of a Nusselt-Rayleigh correlation^ Thus, complete suppression of con- vection means that the apparent Nusselt number as calculated from

Equation (3) or (4) shows no dependence on the apparent Rayleigh

number as determined from Equation (5) -the calculation being

carried out by ignoring the presence of the insulation in the

"convective" zone. Thus, with convection suppressed, we expect

Nu = constant (13)

Ju Nu = constant ' (14) av where the constants are generally different at different locations

and in different chambers, when using the same insulation schema.

III. TEST PROGRAM

The test program consisted of three phases, hereafter

730
referred to as Series A, B and C tests, respectively, as described

in detail below.

Series A Tests; Full-Scale Problem Evaluation

The purpose of these tests was to determine whether a poteittial problem existed, and to obtain preliminary data pertaining

to the pressurized (i.e., other than high-vacuum) operation of

the 29 ft chamber. The chamber consists of a 304 stainless steil

puter shell, 0.625 in thick, externally reinforced with carbon

isteel stiffeners (see Figure 2), Becuase of the carbon steel

structure, the recommended (by the manufacturer) lower temperature limit of the outer shell is 30 deg F. The chamber is equipped

with liquid nitrogen cooled shrouds, normally operating at -320

deg F. There were no provisions for shroud temperature control

at the time of this program.

The outer shell was equipped with eleven thermocouples, and

data were recorded at 15-min intervals, Pressurization was accom- plished by inbleeding gaseous nitrogen. Quasi-stabilized data were recorded at 120 and 500 microns, and 1, 2, and 4 ram Hq internal pressures. Since the lowest temperatures recorded were

approaching the 30 F limit at 4 mm Hg, the test was terminated

at that pressure.

The temperature data are summarized on Table 1, As anticipated (see (Section II) the lowest temperatures recordeS were oi> (the bottom dome, with only a slight temperature drop experienced

Ipy Zone II, and, for all practical purposes, none by Zone III.

trhe conclusion drawn from the data was that the 29- ft chamber

could not be operated in its normal configuration at the pressure

levels and shroud temperatures required for Mars environmental

simulation.

Series B Tests: Reduced-Scale Development

These tests were conducted in a 4-ft diameter by 8- ft high

chamber (see Figure 3), and had two main objectives: 1) to confirm and extend the uninsulated-chamber data obtained during the

Series A tests; and 2) To evaluate potential schemes of minimiz-*

ing the shell cooling problem. Initially, consideration was

given to the use of external heaters as the sole means of preventing subcooling of the outer shell; the idea was abandoned,

i however, in favor of insulation, due to constraints imposed by

available electrical power supply, the possible thermal overload

of the shroud, and complexity of the full-scale installation.

In designing the tests, the guidelines of Section II were

followed. The 4-ft chamber is made of stainless steel without

carbon steel stiffeners and has no lower temperature limit as

far as the present investigation was concerned. However, it wad

731
found impractical to subcool its outer surfaces much below -90 F because of excessive frost buildup. Consequently, the range of

Rayleigh numbers in the Series A and B tests were approximately

equal. All Series B tests were run with liquid nitrogen cooled

shrouds. The uninsulated tests were run at pressures ranging

from high vacuum to 50 mm Hg, whereas all insulated tests were

run at 50 mm Hg chamber pressure.

Candidate insulation materials were evaluated on the basis

of ease of installation and removal, thermal performance, pumpdown characteristics, mechanical integrity, outgassing properties,

and cost. Following bell-jar screening tests on a number of can- didate materials, three were selected for further investigation:

polyurethane foam, crinkled aluminized Mylar sheets, and polyistyrene flakes. It was found that there was no essential differ- ence in their respective thermal performance when properly installed; however, the installation of the ur ethane foam and Mylar was extremely difficult and time consuming as compared to polystyrene flakes. Improper installation did result in convective

Voids and development of cold spots on the chamber shell.

The temperature data are summarized on Table 2, The uninsulated data again indicate that cooling occurs predominantly in

Zone I, although some cold spots also developed in Zone II at higher pressures, near the Zone I /Zone II interface. The apparent

inconsistencies in the Zone II minimum temperatures of Runs 5 through 8 are attributed to differences in the respective exten- sions of the bottom dome insulation into Zone II, The apparent

increase in temperatures from Run 4 to 5 was due to a rearrangement of thermocouple locations.

The superiority of the polystyrene flake insulation concept was evidenced by two factors: the overall increase in Zone II temperatures during Run 10 as compared to Run 9; and by the lack

of development of local frost deposits anywhfere on the chamber's outer shell during Run 10. Because of its excellent thermal performance, fluidity (permitting rapid installation and removal by

^he use of mechanical or pneumatic conveyors), reasonable cost,

and immediate availability, the polystyrene concept was selected

for insulating the 29- ft chamber.

Series C Tests: Full-Scale Insulation Performance Verification

Objectives- The purpose of these tests was full-scale evaluation of thermal performance and operational characteristics of

the fluidized polystyrene insulation concept. Detailed objectives

included: 1) Evaluation of insulation and spot-heater effective-

>iess in maintaining the outer shell temperatures above the 30 F

jlimit, with liquid nitrogen-cooled shrouds, and chamoer pressures

up to 50 mm Hg; 2) Establishing methods and procedures for rapid

installation and removal of the insulation, consistent with typical progr^ schedule requirements; 3)- Evaluation of the pumpdown

732
characteristics of the insulated chamber; 4) Generation of heat

transfer data for predicting shell-to-shroud heat loads; and 5)

Evaluation of potential shell temperature monitoring techniques.

Setting the upper limit of chamber pressure at 50 mm Hg for the tests was somewhat arbitrary; it exceeded the maximum pressure

levels anticipated during Mars simulation tests by a factor of

2.5. The inclusion of spot heaters as a supplemental means of

protecting the shell from subcooling was a precautionary measure. Thirty 250-watt industrial heat lamps were made available for the

purpose^

Temperature Instrumentation - Fifty thermocouples were used

in addition to the permanent shroud-temperature instrumentation

of the facility. Five of these were used to measure ambient tem- peratures at various elevations, twenty- four thermocouples monitored potential cold spots on the shell, whereas tewnty-one were

used to gather engineering data for heat transfer calculations.

The potential cold spots were determined by inspection of the

interspaces between the shroud and the shell, and thermocouples

were attached in the vicinity of liquid nitrogen lines, and near

potential conduction paths between nitrogen-cooled hardware and

the outer surface of the chamber. Typical interspace areas are depicted by Figures 4 and 5, Local cooling was also expected

near the periphery of the uninsulated upper dome. The engineering

thermocouples were placed on predominantly smooth shell surfaces,

through which most of the external heat load is transferred into

the chamber.

Shell Temperature Monitoring - In addition to thermocouple

instrumentation, two other techniques were evaluated: visual inspection, and infrared scanning. Data obtained during the Series

B tests have indicated that incipient frost formation on a surface

is a definite indication that its temperature is approaching 30 F. Periodic inspection of the outer surfaces of the chamber for the

purpose of finding atmospheric moisture condensation or frost

spots has proved to be a very effective monitoring technique

during these tests. The second technique employed an infrared television camera which converts the thermal image into a visual picture displayed

on a picture tube. Permanent records of the infrared pictures

could be obtained by a Polaroid camera attached to the display

unit. Typical pictures are shown on Figure 6, with the white

areas representing low temperatures. Infrared scanning of the

lower dome (Figures 6a and 6b) has confirmed the anticipated locations of potential cold spots, although these areas remained

above 30 F throughout the duration of the tests. Excessive cooling of the upper portion of the cylindrical shell due to insulation shrinkage (to be discussed subsequently) was also detected

by the infrared camera, as shown by the bright white area on the

upper portion of Figure 6c

733
Insulation Installation and Removal - Before installation

of the insulation could be started, it was necessary to seal all

of the diffusion pump inlets, roughing ports, and internal re- pressurization lines. The sealing was accomplished with various

combinations of galvanized wire screens, polyethylene sheeting,

and polyurethane foam strips. Special handling fixtures were designed and fabricated for installation and removal of the insulation. Loading was accomplished with the help of the "elephant

trunk nozzle" depicted on Figures 7 and 8. Figure 9 illustrates

the unloading procedure by the use of a grain auger. The insulated

volume consisted of Zones I and II of the shroud/shell interspace,

comprising approximately 4300 ft^. Pumpdown and Settling - After loading of the insulation was

completed, a pre-test evacuation was made to determine the effect

of vacuum on the mechanical integrity of the insulation. Some set- tling of the insulation was observed: approximately 50 ft^ additional insulation was required to return the level back to its

initial height. The blankoff pressure of the roughing system was

2 mm Hg during this initial evacuation. Considerable contamination of the lubrication fluid of the mechanical pumps was experienced, however, no remedial action was required, since this

fluid could be returned to its original condition after the re- quired operating pressure was obtained. The final vacuum level

at the start of the test was 0.02 mm Hg, with the liquid nitrogen shroud flooded. After temperature stabilization, the chamber was pressurized with gaseous nitrogen, and stabilized data obtained at 2, 4, 10, 20, 30, 40, and 50 mm Hg. Approximately two hours after the chamber pressure was increased to 4 mm Hg, frost

spots were observed to develop at the periphery of the upper dome

and the cylindrical area extending 2 to 3 ft just below the upper

flange of the chamber o Subcooling in these areas was also indicated by infrared scanningo It was found after completion of the tests that the insulation had settled down to several feet from

the upper flange of the chamber, as depicted on Figure 10, This

settling is believed to be caused by vibrations of the shroud

during cool-down with liquid nitrogen. Consequently, in the

future, pre-test pumpdowns will be conducted with the cold shroud

operating. The final pumpdown curve of the insulated chamber is

shown on Figure 12,

Use of Spot Heaters - Excessive frost accumulation due to insulation settling was prevented by the use of the heat lamps

discussed aboveo Due to the limited number of lamps available,

it was necessary to preriodically move the lamps to prevent ex- cessive frost build-up in certain areas. The heat lamp arrangement is depicted on Figure 11, The lamps proved to be adequate

to maintain the shell temperatures in the indicated areas above ' the 30 F limit.

I,

734
Thermal Performance - The temperature data are summarized on Table 3o The performance of the insulation is shown to be pres-*

sure-dependent above 20 mm Hg, although it provides adequate pro- tection to the shell up to 50 mm Hg. Typical temperature traces

are depicted on Figures 13 and 14, Perhaps the most striking fea- ture of the thermal data is the presence of temperature fluctuations at three distinct frequency levels: 1) The diurnal temperature cycle, with a period of 24 h and amplitude in the order of

5 deg F; 2) A thermal "noise" with amplitudes of a few tenths of

a degree, and periods in the order of 1/2 h (temperature recording

was at 15-min intervals); and 3) A random low- frequency oscilla-*

tion at 50 mm Hg pressure displayed by the curve labeled "bottoiA

dome engineering data" on Figure 14, with amplitudes in the order

0f 3 to 4 F and periods in the order of 2 to 4 ho Similar oscillations are indicated by a number of bottom dome thermocouples,

as shown on Figure 15,

A possible explanation of the low- frequency random oscillations is that, at the 50 ram Hg level, a columnar-type convectiv^

flow develops through the insulation, with alternate upward floWs

originating from local hot spots, and downward flows towards th^

colder regions of the bottom dome. Due to convective heat transfer, the hot spots gradually cool down and lose their "activity",

while other spots on the dome take over. During their periods rtf inactivity, the former hot spots become warm again, and the cycle

is repeated.

IV, DATA CORRELATION

Nusselt-Rayleigh type correlations of the bottom dome data

obtained from all three series of tests are shown on Figure 16. The curves were determined according to the procedure outlined

in Section II, The curve on the left hand side applies to the

uninsulated dome, and it may be used to predict internal heat

transfer coefficients for conditions other than those of the

present test program. Such predictions can be used to determine

whether or not a significant shell-cooling problem exists when

using a relatively warm shroud (300 to 400 deg R), and to estab*

lish heat load curves for the temperature control system of the

shroud. Although there is considerable scatter in the data, the

correlation is considered excellent in view of the complex geom*

etries involved, the dissimilarities in the two chambers, and the numerous assumptions used in establishing the calculation procedure (see Section II), A log-log plot of local Nusselt numbers

vs. Raleigh numbers for both chambers yielded approximately par-r ellel lines (with a scatter similar to that on Figure 16), Hence,

it is concluded that Equations (1) and (2) are basically correct. The curve on the right hand side of Figure 16 pertains to the insulated 29- ft diameter dome. The Nusselt and Rayleigh num- bers represent "apparent" values in this case, as discussed in

735
Section II. The most striking feature of this curve is the ap- parent sharp transition at Ra = lO-'--^ (20 ram Hg), indicating the onset of convective flow. Whether or not the low- frequency oscillations depicted on Figure 15 are associated with a second tran- sition at a higher Rayleigh number is impossible to tell from

the limited data available.

V. CONCLUSIONS

1) The polystyrene flake insulation concept provides adequate protection for the outer shell of the chamber to perform

Mars environmental tests. It is reasonably priced, readily

available, and easy to handle.

2) The use of spot heaters is an effective means of cor- recting eventual cold spots due to local voids in the insulation

or insulation shrinkage.

3) Installation and removal of the insulation is consistent

with typical program schedule requirements.

4) Chamber pumpdown with the insulation installed presents

no problem. No harmful outgassing effects from the insulation

have been detected to date.

5) Shell temperature monitoring can best be accomplished

by a combination of thermocouple instrumentation, visual inspection, and infrared scanning.

6) The results of this test program can be extrapolated to other operating conditions by the use of Nusselt-Rayleigh type

correlations.

VI. REFERENCES

(1) Buna, T., "Thermal Testing Under Simulated Martian Environment",

Progress in Astronautics and Aeronautics, Vol,21,pp 391-^175

Academic Press, 19^9

(2) Buna, T, , "Simulation of Planetary Environments", paper presented at the AIAA Test Effectiveness in the 70 's Conference,

Palo Alto, Calif., April 1-3, 1970

(3) Wachter, J. P. , Journal of Spacecraft and Rockets, Vol 7,No.3» PP 350-352, 1970

(k) Batchelor, G. K. , '<^uarterly of Applied I-iathem.atics, Vol, XII,

No. 3, xop 209-233, 195^

(5) McAdams, //.H., Heat Transmission, 3rd ed, McGraw-Hill,

195^, PP 173-177

736
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(a) Bottom Dome, Left (b) Bottom Dome, Center Cc) Side, Upper Left

Fig. 6—Infrared Monitoring of 29-ft Chamber Shell Temperatures
Fig. 9—Insulation Removal

Fig. 10—Insulation Settling

Fig. 11—Heat Lamp Application

743
Fig. 12—Pumpdown Curves, 29x^5-ft Chamber

744
745

J Sap 'BynivysdWBi

747
748
Paper No. 49

A CRYOPUMP FOR STEADY-STATE TESTING OF ROCKET ENGINES UNDER

ALTITUDE CONDITIONS

M. J. Triplett and W. E. Riggs

AEF, ARO, Inc., Arnold Air Force Station, Tennessee 37389

PAPER AND ABSTRACT NOT RELEASED FOR

PUBLICATION

749

Paper No, 50

ABLATIVE HEAT SHIELDS FOR PLANETARY ENTRIES - A TECHNOLOGY

REVIEW

G. D. Walberg and E. M. Sullivan

REFERENCE: Walberg, G. D. and Sullivan, E. M., "Ablative Heat

Shields for Planetary Entries - A Technology Review/' ASTM/IES/

AIAA Space Simulation Conference, September l^-l6, 1970.

ABSTRACT: A review of ablation technology is carried out to assess the present state of the art and point out areas in

which further research is required for planetary entry heat

shields. Analyses and test techniques which have been

developed to treat heat shields for Earth entry with combined

radiative and convective heating are reviewed. With the

lessons learned from Earth -entry research in mind, the work

carried out to date for entry into various planetary atmospheres

is reviewed and technological problem areas are discussed. In

defining significant phenomena, various mechanisms and

processes are discussed and their relative importance is

illustrated by describing the analysis of a manned planetaryreturn Earth entry. In discussing the work to date on planetary entry, two broad categories of research are considered:

(l) entry into tenuous atmospheres and (2) entry into the dense

atmospheres of Venus and the giant planets . In each of these

categories, atmospheric characteristics, entry velocities and

modes, vehicle geometries, heating levels and candidate ablation materials are discussed. Present ground and flight- test capabilities are summarized and compared with planetary entry

conditions. Particular attention is paid to coupled ablativeradiative phenomena which require the radiation spectrum of the

facility to essentially duplicate that of the planetary

atmosphere under study.

'Head, Thermodynamics and Combustion Section, Applied Materials

and Physics Division, NASA Langley Research Center, Hampton, Va. "Assistant Head, Fluid Mechanics Section, Applied Materials and

Physics Division, NASA Langley Research Center, Hampton, Va.

751
KEY WORDS: ablation, heat shields, heat transfer, planetaryentry, radiation, evaluation.

In August of 1957 j> the Army Ballistic Missile Agency

successfully reentered and recovered an ablating nose cone for

a 1500-mile range IRBM, (l)^ and the practicality of the ablative heat shield vas demonstrated. In the succeeding I3 years,

the ablative heat shield has evolved from a dramatically new

concept to a proven, reliable and reasonably light-weight means

of protecting entry vehicles from aerodynamic heating. During

this time, ablation analyses have developed from simple steadystate solutions which consisted basically of transpirationcooling theory with additional heat-absoi-ption terms, to com- plex computer codes which describe the transient, coupled, indepth response of a char-forming plastic to the heating from

radiating, chemically reacting shock layers and boundary layera

This is not to say that all the significant problems in ablation analyses have been solved. We sha2J. see that there are areas in which large uncertainties still exist and in which

much work remains to be done. None the less, the past decade

has produced great advances in ablation technology. In the late 1950's, much research was carried out on subliming and glassy ablators. Most analyses were steady- state solutions since the bulk of the research was concerned with the

ballistic entry of vehicles having fairly high ballistic

coefficients (m/Cj^A) and entry heat pulses characterized by

high heating rates and short duration.

In the early 1960's the emphasis svung toward charring

ablators, and lower density materials began to be developed for

lower m/Cj)A manned entry vehicles. Much work was done on

trsLQsient, in-depth ablation analyses. Military vehicles were

developed with higher and higher m/C^A's. They penetrated

deeply into the atmosphere before decelerating and hence

their entries were characterized by extremely high

stagnation-point heating, high pressures, transition to turbulent boundary layers, and high aerodynamic shears.

Manned vehicles, on the other hand, tended to low

M/CdA's and lifting entries to keep deceleration loads within

human tolerance levels. These vehicles decelerated high in the

atmosphere and their entries were characterized by low pres- sures, low shears, laminar flow, modest heating rates, and

long duration.

The numbers in parentheses refer to the list of references

appended to this paper.

752
Since the entry environments of these two classes of

vehicles are so different, two separate and quite different

areas of heat- shield technology have developed. In the case of

high m/C^A vehicles, high-density materials (such as high

density nylon-phenolic, graphite-phenolic, carbon-carbon

composites, and molded graphites) have been employed to resist

mechanical erosion of the ablator surface, and to keep

geometry change within reasonable bounds. In the case of manned

vehicles, the trend has been to lower density materials since

these materials provide superior insulating efficiency to protect against the long heat- soak characteristic of lifting,

low m/Cj)A entries. Because of the low shears and pressures

associated with manned entries, it has been possible to develop

ablators having low densities but which produce char layers

with sufficient strength to resist mechanical failure within

the flight envelope of the vehicle.

Because of its higher entry velocity, the Apollo vehicle

encountered radiative as well as convective heating. In a sense, the Apollo project ushered in the era of radiative heating. Of course, research on shock-layer radiation preceded the

actual Apollo entries by many years (work was actively under

way in the late 1950' s) but the Apollo project served as a focal point for much of the radiative heating research. In the late 1950' s, many types of facilities were used to test ablation materials. Oxyacetylene torches, combustion

heated wind tunnels, pebble bed tunnels, rocket exhausts,

plasma torches, and arc-heated wind tunnels were used with

varying degrees of success. By 19^2, the arc-heated wind

tunnel had established itself as the most versatile and capable

of the ablation test facilities.

Intensive research was carried out to produce improved arc heaters, capable of operating over wider ranges of pressure and

enthalpy; and radiation sources were added to allow tests with

combined convective and rsidiative heating. Today's ablation

test facilities have reached a high level of refinement and have

played an invaluable role in the development of ablation technology. In spite of their refinement, however, these facilities

provide only a partial simulation of reentry flight conditions.

Because of this_, the proper interpretation of ground-test

results in terms of in-flight ablative behavior remains one of

the most important areas of ablation research.

Throughout the past 12 years, much research in polymer

chemistry has been carried out in an effort to produce improved

ablators. This work has enhanced our understanding of the

pyrolysis process and char formation mechanisms, but no "super"

materials have been produced. In general, advances in heatshield performance have come more through understanding the

753
signifncant phenomena of ablation than through development of

greatly improved materials. When the effects of differing

atmospheric compositions are accounted for, the phenomena

involved in planetary entries and earth entries are the same. Hence the technology developed for Earth entry heat shields

can "be used to, as a starting point, analyze the type of heat

shields required for planetary entry vehicles.

In the present paper, the analyses and test techniques

which have been developed to treat heat shields for Earth entry

with combined radiative and convective heating are reviewed,

and the significant phenomena are defined. The development of

convective and radiative heat-transfer analyses, ablation layer

analyses, and coupled heat transfer and ablation calculations

are reviewed and, as an example of the current state of the

art, a calculation technique developed for treating heat

shields exposed to combined convective and radiative heating

is described. With the lessons learned from Earth entry

research in mind, the work carried, out to date for entry into

various planetary atmospheres is reviewed and technological

problem areas are identified and discussed.

Symbols

A projected cross- sectional area of entry vehicle

elemental mass fraction of ablation products

drag coefficient of entry vehicle

E 4 k: a T

s p s f dimensionless stream function

g acceleration due to gravity

H density scale height

h enthalpy

h^ stagnation enthalpy

oxygen mass fraction

M mass of entry vehicle • M molecular weight m gas injection rate

754
p pressure

\

stagnation pressure

convective heat load

radiative heat load

reduction in heat flux due to convective blockage

convective heating rate ^HEM heat flux due to boundary-layer chemical reactions

radiative heating rate ^eff equivalent nose radius for convective heating

R 7 Reynolds number based on wetted length

cone base radius nose radius

T temperature

T

s post shock temperature

V velocity

z altitude r radiation loss parameter =

^ / ^ ^oo entry angle

Ah enthalpy potential

5 gas cap thickness

ec cone half angle

KP Planck mean absorption coefficient

P density

0 Stefan- Boltzmann constant

T aerodynamic shear

755
Subscripts:

e entry condition

V evaluated at the surface

00 free stream

Convective and Radiative Heating for Earth Entry

Convective Heating

Convective heating has been a source of serious concern

ever since the early days of supersonic flight. (Remember the

press having a field day talking about aircraft crossing the

sonic barrier and landing in the thermal thicket?) We will

briefly review the development of nonablatlng heating calculations. The coupling of heating and ablation analyses will be

discussed later.

Several specific areas can be pointed out in the literature on convective heating. In particular, the problems of

stagnation-point heating, laminar heating, and turbulent heating calculations for very high-speed entry were put on a firm

footing by the work of Fay and Riddell(2), Cohen(3), and

Hoshizaki(^) . More recently the work of De Rienzo and

Pallone(5) permits extension of stagnation-point convective

heating up to entry speeds of 21.33 km/sec. This last paper

also gives an excellent summary of convective heat- transfer

studies including the shock- tube experiments which formed the

basis for much of the theoretical work. The work of Boison and

Curtiss(6) and Zoby and Sullivan(7) explored the influence of

body geometry on stagnation-point heating.

Laminar heat transfer on flat plates, sharp cones, and off

the stagnation point on blunt bodies has been well established

by the work of Cohen(3) and Lees (8) . The results of Cohen

have been correlated by Zoby(9) to put them in a more usable

form.

Turbulent heating is still a subject for extensive research.

The Reentry F spacecraft has provided an extensive amount of

turbulent heating data which are still being compared with

existing correlation equations. The earlier work of Zoby and

Sullivan(lO) showed that turbulent heating for a large number

of flight vehicles, blunt and sharp, could be predicted with

fair accuracy using reference enthalpy methods.

No discussion of turbulent heating is complete without a discussion of the problem of boundary-layer transition. Transition is known or believed to be influenced by Mach number.

756
Reynolds number, leading-edge bluntness, surface roughness, and

a host of other parameters including the ablation rate for an ablating surface. To date, no good correlation of boundarylayer transition is available. Any of the currently accepted

correlations show a spread of the type shown in Figure 1 which

is taken from Reference (ll) . Note that there is as much as an

order of magnitude difference between the lowest and highest

Reynolds numbers at which transition occurs. In addition, the

influence of non-air gas mixtures on boundary-layer transition

is not well known. However, barring major chemical effects

which can significantly influence temperature and enthalpy

profiles, there is no reason to believe that a transition

correlation for air will not apply in non-air gas mixtures.

Radiative Heating

Radiative heating becomes an important design consideration when the gas- cap temperature exceeds approximately

10 000*^ K. A history of radiative heating calculations is

given by Anderson(l2) who points out that calculations using an optically thin or a gray gas model seriously overpredict the

radiative flux to the wall. JUgure 2, which is copied from

Anderson's paper, shows how the state of the art for radiation

caJLCulations has changed in the past decade. Calculation

procedures have progressed frcM the transparent, continuim

only, uncoupled calculations used by many early investigators

throtigh the nongray, coupled but continuum only calculations

of Hoshizaki and Wilson(l5) to the coupled with continuum and

lines calculations which are current today. In Figure 2, the

last category is demonstrated by results from Page et a.l.{lk-)

but could be equally well demonstrated with results from

Wilson(l5), 01stad(l6), Chin(lT) . or Rigdon(l8) . The radiation

model used in the calculation procedure described below is that

of Wilson(l9) which includes continuum and atomic lines.

Anderson further points out that due to uncertainties in gas

absorption coefficients, the radiative flux "can probably be

determined within 50 percent" and if the uncertainty in the

absorption characteristics of ablation products is also taken

into account the radiative flux "is known within a factor of two." It is obvious that if the entry environment is dominated by radiative heat transfer the heat -shield weight

requirement can be a serious uncertainty.

Most of the radiative heating calculations made to date

and referenced above apply to the stagnation region. It was recognized prior to I962 that radiative heating to conical

bodies was negligible because of the cooling of the gas as it

moved out of the stagnation region and onto the conical afterbody. However, bodies with larger cone angles have become

intersting (e.g.. Viking) and attention has again turned toward

this problem. A very recent paper by Callis(20) shows results

757
for sphere- cones with cone angles of 50*^, ^5°^ and 6o°. As anticipated, for the conditions investigated, the 50° and h'f^ cone heating rates fall to very small fractions of the

stagnation-point values. The 6o° cone, however, can experience

radiative rates which may actually be higher than the

stagnation-point values. Such results can have serious implications on both the heat- shield design and the aerodynamic

characteristics of the entry vehicle. These results point out

that radiative heating is a major unlmown for the design of a vehicle flying in air. For vehicles flying in non-air gas mixtures, the uncertainty must be larger and requires further

consideration.

Ablator Response

Ablation Analyses

During entry, an ablator responds to heat transfer and

chemical attack from the adjoining shock layer. The ablator

responds by receding, altering its internal temperature

distribution, or both. The principal mechanisms of ablative

response, as illustrated in Figure J), are:

1. Sensible heat increase in the virgin plastic and char

2. Pyrolysis of the virgin plastic to form char and

pyrolysis gases

3. Enthalpy increase in the pyrolysis gases as they flow

through the char

h. Rem.oval of the char surface by chemical attack and

sublimation

5. Mechanical removal of the char

6. Reradiation of energy from the char surface

Most present-day ablation analyses assume one-dimensional flow

of heat and mass normal to the local char surface. Such an

assum.ption is valid for cases where the curvature of the heat

shield is large compared to characteristic dimensions within

the ablator (large nose radii), but for small ground-test models and for flight vehicles with small nose radii multidimensional solutions may be required. This is true in treat- ing both temperature distributions and the internal flow of

pyrolysis gases. At the present, complete multidimensional

analyses have not reached a sta^e where they are suitable for

routine use. Solutions have, however, been obtained for the

important limiting cases of two-dimensional heat transfer with

758
one-dimensional mass transfer (subliming ablators and graphite - no char layer) (21, 22) , and two-dimensional pyrolysis gas flow

with one- dimensional heat transfer (models with rapidly varying

external pressure distributions) (23) . In the present paper, for

the purpose of illustrating basic phenomena, we shall, for the

most part, restrict our attention to the one-dimensional case. A typical one-dimensional ablation ajialysis is ccanposed of a finite-difference solution of the transient heat conduction equation throughout the char, pyrolysis zone, and virgin plastic; a set of empirical kinetic equations that describe the pyrolysis process;

a chemistry routine to determine the state of the pyrolysis gases; a

set of equations which describe char removal by chemical and

mechanical means; and a char surface energy balance(2l4-, 25, 26, 27, 28). In the early days of ablation analysis, it was often necessary

to sacrifice some rigor in the solution of the heat- conduction equation in order to keep computing times reasonably short. This problCTi

has now been overcome by improved programing techniques and increases

in computer speed so that today' s finite difference solutions are essentially exact and require reasonably short machine times. At

the present time, the accuracy of predicted internal temperature

distributions is limited mainly by the material properties required

as input data (i.e., specific heats, thermal conductivities, etc. ). Virgin plastic properties can be measured with reasonable accuracy

at temperatures where pyrolysis of the material does not occur. Since the techniques used to measure specific heats and thermal

conductivities are steady- state procedures, measurement made on

a material which is pyrolyzing must contain some uncertainties ( 29 )

Hence, material properties in the pyrolysis zone are difficult

to obtain. The largest difficulties with input properties are associated with the char layer. Most ablation analyses use a char thermal conductivity which is a lumped parameter and

represents the net conductivity of the char and the pyrolysis

gases flowing through it. This lumped conductivity cannot actually be measured and is often inferred from arc tunnel tests

of ablation models. That is, the char conductivity is adjusted

until the predicted tanperature distributions agree with those

measured in the tunnel tests(30,3l) . In spite of the foregoing

difficulties, present-day ablation analyses usually predict

reasonably accurate internal temperature distributions. In reference 30 for instance, after adjustment of the char thermal

conductivity, internal tonperatures were predicted to within

30 percent of the measured values.

Pyrolysis Kinetics and Pyrolysis Gas Chemical State

The pyrolysis kinetics, as obtained from DTA (differential

thermal analysis) and TGA (thermo- gravimetric analysis) data,

generally predict pyrolysis rates and energy absorption with

reasonable accuracy(52)

759
The state of the pyrolysis gases ronains one of the major

unsettled questions in ablation analysis. Its significance is

illustrated in Figure k. Early ablation analyses usually

treated the pyrolysis gases as frozen at the composition

at which they left the pyrolysis zone. If this assumption is

made, then the pyrolysis gases will absorb only a small sunount of energy in being heated from the pyrolysis temperature to

the char surface temperature. Later, gas phase equilibrium

computer codes were employed to treat the gases as being in

complete chemical equilibrium (including the deposition of

carbon in the hotter regions of the char) . This equilibrium

assumption results in the prediction of large heat absorption

by the pyrolysis gases. Experimental data have been obtained

which support both the equilibrium and frozen assumptions.

Data obtained by Lundel(55) which agree with the frozen assumption are presented in Figure k, A more accurate approach would

be to carry out a kinetic analysis which employed measured

rate constants for the important reactions. Such an analysis was reported in Reference (5^) and indicated that the equilibrium assumption would significantly overpredict the heat

absorption by the pyrolysis gases. The available rate data

are, however, limited to relatively low temperatures and, in

many cases, are available only for overall reactions rather

than the individual kinetic steps involved in the pyrolysis

gas reactions. Hence, a true kinetic model for the pyrolysis

gas reaction has not been established and, lacking such a

model, analyses based on low temperature data cannot be

extrapolated to the high char temperatures associated with

lunar return, planetary return, and planetary entry missions.

Char Surface Recession

Another important area of ablation analysis that is still

largely unresolved is the prediction of char surface recession,

^ost analyses for manned vehicles employ the assumption that

char recession is due entirely to chemical reactions and

sublimation. This is usually justifiable within the flight

envelope and makes the problem more tractable than if mechanical

char removal had to be accounted for> but the question of char

surface kinetics still presents formidable problems. Most

analyses treat chemical char removal as being jointly rate and

diffusion controlled. At low temperatures, where oxygen can diffuse to the char surface faster than it reacts, the reces- sion rate is governed by the kinetics of the solid-gas reaction*

At high temperatures, the rate of char recession is governed by

the rate at which oxygen (and other reactive boundary-layer

gases) diffuses to the char surface. Between these two regimes

there exists a transition region of joint rate and diffusion

control. At yet higher temperatures, sublimation causes additional char recession. Most investigators have assumed that

the rate controlled regime could be described by a single

760
Arrhenius equation with a constant activation energy. This

surface recession model has received such wide acceptance that

its validity is seldom questioned. However, data recently

obtained for various graphites strongly suggest that the

activation energy is, in fact, a decreasing function of temperature and that char surface removal may be chemically controlled

to temperatures much higher than previously thought

possible(35^ 56) . Sublimation is another area of considerable

uncertainty. Several investigators have shown that accurate

solutions require the inclusion of high molecular-weight gas phase carbon species (C]_ through Cq_q usually) . The thermodynamic properties of these large carbon species are not well

established. Depending on their treatment of these species,

various investigators have predicted significantly different

sublimation rates at high temperatures and pressures(37^ 38)

When, as in the case of high m/C;qA ballistic vehicles,

mechanical char removal must be accounted for in addition to chemical removal, the prediction of char recession rates becomes uncertain indeed. Several theoretical treatments of

char failure have appeared in the literature (39^^0) . Most of

these analyses describe char removal as a cyclic process wherein

the char builds up to a critical thickness and then is removed

by a combination of thermal and internal pressure stresses. While this type of char failure has been reported for some materials, it is not encountered very often. The type of char

failure that is most often seen in ablation tests involves the

removal of relatively small fragments from the char surface.

Figure 5 presents a typical set of test data for a carbonphenolic ablation material. This type of char failure only

occurs when the test stream contains oxygen. It is apparently

a coupled chemical-mechanical process in which the char surface

is weakened by oxidation and subsequently removed by aerodynamic

shear. The phenomenon is particularly apparent under highpressure test conditions and has been observed for graphites (4l) as well as charring ablators (42) . At the present time, this

type of char removal must be described by empirical relations

derived from ground tests

Reradiation of Energy From Char Surface

The superiority of the charring ablator (over subliming

and melting ablators) is largely due to its ability to reradiate

large quantities of energy from a high-emissivity, hightemperature char surface. As will be shown in a subsequent

section of this paper, this reradiation is one of the most

important energy accommodation mechanisms for charring ablators. In most ablation analyses, the char surface is assumed to be a gray body and to have a constant emissivity. For most

carbonaceous materials, this seems to be a reasonable assumption.

However, materials which contain large amounts of silica often

produce chars for which the gray-body assumption is significantly

761
in error(^5) • The counterpart to emissivity is char surface

absorptivity, which becomes important when the ablator is

subjected to radiative heating. As with emissivity, the

absorptivity is usually assumed constant, though its value is

usually assumed to be different from that of the emissivity to account for the fact that the shock-layer radiation and the

char surface reradiation involve different wavelength ranges.

Actually, little is known about char absorptivity for short

wavelength, high- energy (ultraviolet) radiation.

Since, reradiation from the char surface varies as the

fourth power of surface temperature, it becomes more important

as heating rates (and hence surface temperatures) increase. At

a temperature of approximately ^000° K (for pressures of the

order of 10 atmospheres), however, carbonaceous materials

sublime. Hence, an upper limit on the amount of energy that can be reradiated exists (around 1200 w/cm^). For entries

which involve heating rates of a few thousand w/cm^ (manned

planetary return earth entry, for instance), char surface

reradiation will usually be the most important energy accommodation mechanism. For entries that involve tens of thousands of

w/cm2 (which, as we shall see, may be the case for Jupiter

entry) , reradiation may be small compared to other ablative

mechanisms

Categories of Ablators

A great many different materials and formulations have

been investigated for use as ablative heat shields . It might

be expected that each of these materials would have its own special characteristics and would require a separate research

program to determine its ablative behavior. To an extent, this

is true, particularly with regard to thermophysical and thermochemical properties. In a broad sense, however, ablators may

be meaningfully categorized according to their elemental

chemical composition (i.e., the nmnber of atom^ of carbon,

hydrogen, etc.). The basis of such a categorization is the

concept of complete (gas-phase and gas-solid) equilibrium at

the char surface. If such equilibrium is obtained, then, in a diffusion limited situation, the char recession rate (which is one of the major determinants of ablator performance) is

completely determined by the vehicle geometry, the flight

conditions, and the elemental composition of the virgin plastic.

The assumption of complete char- surface equilibrium has been

studied, both analytically and experimentally by many investigators(26,^^,^5^^6) . In many cases, it has been shown to provide an accurate prediction of ablator behavior. As mentioned previously, there are data available which suggest

significant pyrolysis-gas nonequilibrium, and in such cases, the concept would not hold. As a guide to the overall

categorization of ablators, however, the equilibrium concept is

quite useful. When considered in this way, most ablators can

762
be grouped into two classes; those which contain the elements

carbon, hydrogen, nitrogen, and oxygen, and those which contain

these four elements plus silicon. These two groups differ most

significantly with respect to char chemistry. Materials which

contain C, H, IT, and 0 produce chars which are nearly 100 percent carbonaceous and whose chemical composition is virtually

independent of test condition. Materials which contain Si, on

the other hand, produce chars with surfaces composed primarily

of solid SiOp^ liquid Si02, solid SiC, or solid C depending on

the test condition. This behavior is particularly characteristic of elastomeric ablators and is illustrated by the results

presented in Figure 6. I\ote that as the heatin-g rate is

increased, changes occur in both appearance and chemical composition of the char layers. Over this same range of test conditions, an ablator containing only C, H, iM, and 0 would

everywhere produce carbonaceous char. Silicon-dominated ablators (ma-terials which behave as shown in Fig. 6) are best

s-oited for use at low" heating rates (low surface temperatures)

where the char is primarily solid Si02 and, since it is an

oxide, is virtually inert with respect to oxygen and hence is nonreceding. Carbon dominated ablators, on the other hand, are preferable at high heating rates since their carbonaceous chars

do not suffer from the liquid layers and c/Si reactions characteristic of siliceous materials.

Coupling Between Heating and Ablator Response

One of the outstanding characteristics of an ablative heat

shield is that it not only absorbs heat but, through the injection of gaseous ablation products, it also modifies the adjacent

boundary layer and greatly reduces the level of aerodynamic

heating. When the ablation products enter the boundary layer

they are heated up, react with themselves, and (usually

exothermically) with the boundary-layer and shock-layer gases>

they thicken the boundary layer and alter its profiles of

temperature, velocity, and species concentration, and they

absorb radiation from the shock layer.

The exothermic reaction of the pyrolysis gases is both

detrimental and beneficial. Because of its exotheimiic nature,

it actually increases the heat flux to the ablator. However,

these reactions also deplete the reactable oxygen in the

boundary layer, reduce the oxygen flux to the char surface, and

hence, in a diffusion limited situation, reduce the rate of

char removal by chemical reaction. This combustive heating can be expressed as an effective increase in stagnation enthalpy.

Hence, it is quite important at low flight speeds (low enthalpies) but becomes of lesser importance at high speeds. The

oxygen depletion effect is important at both high and low

speeds so long as chemical char removal is significant compared

to removal by sublimation and mechanical processes. In treating

763
the ablation product boundary-layer reactions there is, once

again, the question of reaction rates. The limiting cases of

equilibrium and frozen chemistry are, of course, available. At higher boundary-layer and shock-layer temperatures ( say above 2000° K) , equilibrium is usually assumed since reaction

rates are expected to be fast. A considerable amount of data

supporting this assumption has been published, but in many

cases the frozen assumption seems to give better results. The

analysis that was most successful in predicting the performance

of the Apollo heat shield treated the ablation products in the

boundary layer as being frozen.

The thickening of the boundary layer and the alteration of

its profiles constitutes the well-known "transpiration cooling

effect." This is an important ablation mechanism under

virtually all flight conditions, and becomes particularly

important under conditions producing high radiative heating

rates. Under these conditions, the resulting high mass injection rates "blow the boundary layer off" the char surface and

reduce the convective heating to zero. This phenomenon is usually illustrated by a plot of the ratio of convective heat

flux (or Stanton numbers) with and without blowing versus a nondimensional mass injection rate. From such a plot,

presented in Figure 7^ it is seen that the convective heat flux

tends to zero with increasing blowing rate as shown by the

solid cujTve. Experimental ablation data have been presented,

however, indicating that, with increasing blowing rate, the

convective heating approaches a finite asymptote (0.2 to 0.5 of

the no-blowing value) rather than zero(4-7) . This type of

behavior should be expected when the blowing is produced by an ablator responding to purely convective heating. If the heating were actually reduced to zero, the ablator would stop

ablating, the blowing rate would go to zero, and, with no mass

injection, the heating rate would increase to its no-blowing

value. Hence, some nonzero asymptote must be approached in

convective heating situations. This type of behavior will not

occur, however, in the presence of large radiative heating

rates. In this case, the high mass injection rates (ablation

rates) resulting from the radiative heating can, indeed, drive

the convective heating rates to extremely low values

Until a few years ago it was generally assumed that the

gaseous ablation products would not absorb a significant part

of the radiation from the hot shock layer. Then, a number of

investigators carried out analyses in which the absorption by

the ablation products could be accounted for(^8,^9) . They

found that the ablation products could actually be rather

efficient absorbers (especially at ultraviolet wavelengths)

and predictions of up to 50 percent reduction in radiative

heat flux were published. Recently analyses of this type

have been further refined and now reductions of around 25 per- cent are being predicted for typical planetary return

764
Earth entries ( 50) . Analyses of this type are discussed in the

following section of this paper.

State- of- the-Art Analysis for Planetary Return Earth Entry

In order to identify those ablative mechanisms which are most important during a high-speed entry, a typical up-to-date

analysis is described and typical results for planetary return

Earth entry are reviewed.

A state-of-the-art analysis which is being used to study

some of the problems associated with Earth entry of a blunt

body at planetary return speeds is given by Smith et al. in

Reference (50) . Figure 8 is taken from that reference to show

the logic involved in developing such an analysis

The analysis must begin with a trajectory if a transient

ablation solution is required. If only simple, point calculations are required then velocity and altitude or velocity and

density are all that are necessary to initiate the analysis.

The inviscid radiating solution is used to calculate all

the flow parameters in the subsonic portion of the flow field

behind the bow shock. In the analysis shown in Figure 8, the

method used is that of Suttles(5l) who combined the one strip

integral method(52) with the radiation model of Wilson(l9)

The output from the inviscid flow- field program is used to provide local values of pressure, temperature, density, and

velocity. This information can be used as local edge conditions with a conventional boundary-layer solution to provide initial estimates of the convective heating. In the

analysis of Reference (50), these convective heating calculations were made using the correlation equations of Zoby(9)

for equilibrium air.

The radiative flux calculated by the inviscid flow- field

program and the convective heating rate calculated by the

boundary-layer program are used as inputs to the ablation

program (developed by Kendall, Rindal, and Bartlett(26) ) which

computes the transient, one- dimensional response of a charring

material with heat conduction and in-depth pyrolysis governed

by Arrhenius type rate equations. The pyrolysis gases are assumed to be in chemical equilibrium throughout the char, and

two-phase chemical equilibrium is assumed at the char surface.

All char recession is assumed to be caused by rate and diffusionlimited chemical reaction and sublimation.

When the ablation rates have been calculated, it is possible to compute the velocity profiles in the boundary layers

using the technique described in Reference (50) . By testing

765
the resulting stream function against a limiting value, the

analysis decides whether or not the ablation rate is large

enough to move the air-ablation products boundary layer away

from the wall and replace it with an ablation products layer

with an almost constant, low-value shear at the wall. If the

ablation rate is sufficiently large the strong injection

solution is used. This solution, described in detail in

Reference (50), assumes that the boundary layer consists of a layer of ablation products next to the wall and a layer

within which the viscous effects adjust the flow variables

from the inner layer values to the inviscid outer layer values.

In the layer next to the body, constant shear and negligible

conduction are assumed. In the viscous layer, the elemental

composition is computed by assuming a cubic variation from

the ablation products composition in the inner layer to the

air composition at the edge of the inviscid outer layer. If

the ablation rate is moderate, a conventional boundary-layer

solution is used in which the stream function at the wall is

defined by the ablation rate. The inviscid outer layer and the boundary layer for either

strong or moderate injection are coupled to provide continuous

enthalpy profiles across the shock layer. First, the inviscid

layer is displaced from the wall by a distance equal to the

displacement thickness of a boundary layer with mass addition.

The boundary-layer edge enthalpy value is matched to a value

in the inviscid layer near the edge; of the boundary layer.

This information provides a smooth enthalpy profile which is

used by the radiation model to recompute the radiative flux.

This flux is used as input for the next iteration on the mass loss rate calculations. The solution continues until a consistent set of flux and mass loss rate values is obtained.

There are several radiation models which might be used

for calculation of radiative transfer through the ablation

products layer. As previously noted, the radiation model used

in Reference (50) was that of Reference (19) . This model

contains the major features for air radiation as well as the

prominent features from a large number of the chemical

compounds resulting from an equilibrium analysis of the ablation products from common ablators such as phenolic nylon

and phenolic carbon. Thus, when this code is coupled with

the boundary-layer solutions described above, the major

influence of the ablation products on the radiative heating

has probably been accounted for. The answers should certainly

be within the factor of 2 mentioned by Anderson. Typical

results from analyses of this type are presented in Figures 9 and 10. Figure 9 shows the extent of the ablation products

layer and the temperature distribution through the shock layer,

and presents the calculated shock stand-off distance, heating

rates, and the char- surface temperature for conditions typical

766
of peak heating during a manned planetary return Earth entry.

Blockage by ablation products reduces the inviscid heating by

about 22 percent. It is of interest to see in which spectral

regions the absorption is taking place. This is shown by the

difference between the dashed and solid lines in Figure 10. Note

that the absorption is almost entirely in the ultraviolet region.

Only by carrying out a coupled analysis such as that just

described is it possible to accurately assess the heat -shielding

requirements for severe radiation-dominated entries. Coleman,

et al., carried out a thorough study of heat- shielding require- ments for blunt and conical manned planetary return entry

vehicles. Their analysis included all the important mechanisms

mentioned previously. For a baseline theoretical model (which

assumed a phenolic-nylon ablator, no mechanical char failure,

equilibrium pyrolysis gases, laminar flow, equilibrium shocklayer chemistry, and no blockage of radiation by ablation

products), they showed the magnitudes of the various heat

absorption and blockage mechanisms as a function of time during

entry. These data are reproduced (with some modification) in

Figure 11. In constructing Figure 11, the data of Coleman

et al. (49) were modified to reflect a 25-percent radiation

blockage due to ablation products since, in light of recent

developments, this is believed to be realistic. Results are presented for the center line of an Apollo-like vehicle. For

the blunt Apollo-like vehicle, radiation is the dominant heating mechanism, and the significant energy accommodation

mechanisms are radiation blockage by ablation products, tran- spiration cooling, siirface reradiation, °nd the enthalpy

increase in the pyrolysis gases. For conical vehicles, the

dominant heating was found to be convective, and the significant energy accommodation mechanisms were transpiration cooling, reradiation, and pyrolysis enthalpy increase.

Ablative Response in Turbulent Flow

All the results presented thus far are rigorously applicable only to laminar flows. Actually, little is known about

the response of ablators subjected to turbulent flows. Since

arc tunnels are generally low-density facilities and can accommodate only small models, ordinary ablation tests never produce Reynolds numbers high enough to produce turbulent

flow. Some ablation tests have also been carried out in

turbulent pipe-flow and channel-flow facilities and have

produced valuable data, but because of radiation exchange with

facility walls and difficulties in precisely defining the

flow, the application of these results to flight conditions

is not straightforward. Some turbulent ablation data have

also been obtained from flight tests. Data for several

materials having densities from about 0.3 to 1 gm/cm5 were obtained from the NASA Pacemaker series of small flight

767
vehicles and data for a variety of high-density materials have

been obtained from flights of military vehicles. Unfortunately,

Mach numbers and enthalpies over which these data were obtained

are considerably lower than those for severe planetary entries.

The analysis of the response of an ablator to turbulent

flow is usually done by calculating the convective heating

rates by available turbulent boundary-layer techniques and

reducing the convective blocking effectiveness of the ablation

products to 1/3 or l/k of their laminar values. Very little in

the way of coupled radiative- convective analyses with turbulent

flow has been attempted thus far since it has generally been

assumed that radiative heating will be insignificant in the

afterbody regions where the flow could be turbulent. In

general, the approach used with nonmilitary vehicles has been

to choose vehicle geometries and entry trajectories so as to avoid turbulent heating. This same approach will probably be

used with planetary entry vehicles.

Planetary Entries

Ablation research for Earth entiy has been reviewed. Now

we shall use this background and consider the problems associated with planetary entries. From our review of Earth entry

research, we have seen that ablative heat shields provide

reliable, reasonably lightweight thermal protection systems,

but because of uncertainties in several key technology areas

(boundary-layer transition, chemical state of pyrolysis gases,

char- surface recession mechanisms) a conservative design

approach must be used. At the present time, these same uncer- tainties will require conseivative designs for planetary

vehicles. For severe Earth entries, the most significant ablative energy accommodation mechanisms were seen to be blockage

of convective and radiative heat flux by ablation products,

char surface reradiation, and pyrolysis enthalpy increase.

Various planetary entries will now be considered and the most

significant ablation mechanisms identified in each case. At this point it should be pointed out that prelaunch and

transit environments such as sterilization, vacuum exposure, and

cold soak can significantly influence the choice of an ablation

material for a particular mission. The study of these effects

constitutes a broad technology area in itself and is outside

the scope of the present paper. It must be remembered, however,

that the material properties used in the analyses described

herein must be those that the ablator possesses after its interplanetary trip. Just prior to atmospheric entry. In Table 1, properties of various planetary atmos- pheres(53,5^,55,56,57) and unpublished Project Viking atmospheric

768
models are presented"^. The near planets may be grouped into

two categories, those with atmospheres considerably less

dense than the Earth's and atmospheres more dense than that of

Earth. In the low-density group, the planets of most current

interest are Mars and Mercury. In the high density group, they

are Venus and Jupiter. In the present paper, we shall consider Mars as being

representative of the planets with tenuous atmospheres.

Pritchard(58) has considered the possibility of using a vehicle originally designed for Martian entry to explore Mercury, Titan (a moon of Saturn), and other planets

with thin atmospheres. He concluded that, for surface pressures as low as 1 mb, an entry vehicle designed for Mars could

be used for Mercury and Titan missions with appropriate

changes in the terminal descent system and, perhaps, in the

guidance system. Hence, Mars is a good focal point for lowdensity-atmosphere entries. In the case of entry into dense atmospheres we shall

consider two planets, Venus and Jupiter. Venus is selected for

consideration since it is unique in posing entry problems some- what comparable to those of Earth entry. Jupiter is chosen as being characteristic of the giant planets (Jupiter, Saturn,

Uranus, Neptune). This is done for two reasons. First of all,

interest in Jovian entry is currently high and, as a result,

much more information is available for Jupiter than for the

other giant planets. Second, Jupiter is thought to possess

many of the atmospheric characteristics that would be encoun- tered during Saturn, Uranus, and Neptune entries.

Entry Into Tenuous Atmospheres - Mars

The essential problem associated with entries into tenuous atmospheres is that of decelerating to conditions at which

parachutes or other high drag devices can be deployed. This

must be accomplished at altitudes which are high enough to allow atmospheric sampling in most planetary exploration missions. In order to achieve this high-altitude deceleration,

low values of the vehicle ballistic coefficient, u/Cjyk, are required. As pointed out by Roberts(59)^ deceleration within

a given atmosphere requires that the atmospheric drag parameter,

P/(Mg/C]yi)

The values presented in Table 1 are intended only to roughly

categorize the various atmospheres. It is recognized that

many of these values are uncertain and debatable.

769
be less than approximately 10. Hence, a vehicle with an M/CjyV

of 51^ g/cm^ will be decelerated by the Earth's atmosphere

whereas^ deceleration in the Martian atmosphere will require

m/C]3A's on the order of 3-1^ g/cm^.

Atmospheric Characteristics

Our concept of the Martian atmosphere has changed

drastically in the past few years. When Martian entry vehicles

began to be seriously studied in I96I-62, it was recognized

that much uncertainty existed regarding the pressure and

composition of the Mars atmosphere. Early estimates of sur- face pressure ranged from 10 to 30 mb (O.Ol to O.O3 atm), and

compositions from 80 percent nitrogen , 20 percent CO2 to

100 percent CO2 were considered possible. In that time period

most studies assumed significant amounts of both N2 and CO2. Recent spectroscopic data and measurements of the occultation

of radio-frequency transmission from the Mariner flyby spacecraft have, however, considerably improved our knowledge of the

Martian atmosphere. The Project Viking Mars Engineering Model

Working Group has reviewed these data and proposed a set of

model atmospheres with surface pressures ranging from k- to

10 mb (0.004- to 0.01 atm), compositions from 100 percent CO2 to 71 percent 002^ 29 percent Ar, and density scale heights (in

the stratosphere) from ^.5 to 12 km. These various atmospheres

correspond to maximum and minimum surface densities and a mean

model. Density and temperature profiles for these model

atmospheres are presented in Figures 12 and I3. Entry Vehicle G-eometries, Entry Modes, and Velocities

By far the most often used geometry for Martian entry

vehicles is the spherically blunted cone. This configuration

has the advantages of low ballistic coefficient, good payload

packaging characteristics, and good aerodynamic stability.

The Project Viking entry capsule is shown in Figure 1^ as an example of this type of configuration.

Nearly all the studies carried out to date have considered

ballistic entries. While many early investigations featured

vertical or near-vertical entries, the trend in recent years

has been toward entry angles of -20° or less. For low

ballistic-coefficient vehicles (m/Cj)A = 3 -> 5 gm/cm^) entering

at shallow angles, heating rates tend to be very low. The

maximum heating rates on such vehicles as calculated in

References (60), (61), (62), and (63) are summarized in

Table 2. From this table it is seen that maximum heating

rates range from 100 to 200 w/cm^ for direct entry and are on the order of 20 w/cm^ for entry from orbit. Furthermore,

radiative heating is inconsequential for entries less than

about 7 km/sec. Hence, for the Martian entries of greatest

770
current interest (e.g., orbital entries such as Project Viking),

the heat- shielding problem is not severe. Only in the case of

very high-speed advanced missions (such as the 11.5 km/sec Mars

Surface Sample Return mission mentioned in Reference (62)) do

the heating rates become high enough to really exercise presentday ablators. Since high-speed entries of this type have much

in common with Venusian entries, they will not be discussed in

this section.

Since the heating rates are so low, the selection of an

ablative material for a Martian entry vehicle is influenced

more by its ability to withstand such environments as prelaunch

sterilization, in-transit cold soak, solar radiation, micrometeorite bombardment, and vacuum exposure(63^6^) than by its

high ablative efficiency. From an ablative performance standpoint, the outstanding requirement is that the material be of

low density and possess good insulating capability. Filled

elastomeric materials seem to be well suited to these require- ments . A typical material might be composed of silicone resin,

silica microspheres, phenolic microballoons, and carbon and

silica fibers. This type of material has been studied and

tested extensively(46, 65^66,67) and, for low heating-rate applications, is well proven.

The heat shield remains an important part of the overall

vehicle design for a Martian entry vehicle. Most design

studies (67, 68) indicate that the heat shield will comprise

10 to 20 percent of the entry vehicle gross weight. The

materials are developed and available, however, and the overall

problem seems we3JL within the present state of the art.

Entry Into Dense Atmospheres

As shown in Table 1, the near planets may be grouped

according to atmospheric density. The planets of immediate

interest which have atmospheres as dense or more dense than

that of Earth are Venus and the giant planets, Jupiter, Saturn,

Uranus, and Neptune. In most prior research, Venus and Mars

have been considered together because of their similar atmos- pheric composition. If comparable entry velocities are considered for the two planets, this would be a logical grouping. In most cases, however (because of the difference in mass of the two planets), Venusian entry velocities are much

higher than those for Mars . As a result, Venusian entries

usually entail severe radiative and convective heating and

pose a significant challenge to present-day ablative heat

shields. As discussed in the preceding section, this is

usually not the case for Mars. Except for the fact that they

both pose appreciable heat- shielding problems, however, Venus

entries and entries into the atmospheres of the giant planets

have little in common. Hence, they will be discussed as separate categories.
Venus

Atmospheric Characteristics

Data obtained from the flights of Mariner V and the Russian

Venera k vehicles have greatly reduced the uncertainties associated with the density-;, temperatures, and composition of the

Venusian atmosphere. From analyses of these data(55)^ it

appears that the altitude from which Venera 4 sent its last

radio transmission is in doubt. Originally it was reported

that Venera h- had measured pressures of from l6,h to 20.3 atmospheres at the Venusian surface. However, if it is

assumed that the last Venera k transmission came from an altitude of approximately 50 km, then the Venera h data fit nicely

with the Mariner V data, define the temperature and pressure

profiles in the Venus atmosphere quite well, and surface pressures of 167 atmospheres are indicated. This interpretation

of the data is appealing but there are still arguments pro and

con. As a result, model atmospheres have been developed

corresponding to surface pressures of l6,h- and 167 atmospheres

and various degrees of solar activity. Density and temperature

profiles for these atmospheres are presented in Figures 15 and

16. The composition of the Venusian atmosphere is now believed

to be between 90 percent C02^ 10 percent N2, and 100 percent

CO2. From an entry heating standpoint, the atmosphere is

reasonably well known since the important parameters are composition and scale height, and the scale heights in the

sensible (from a heating standpoint) atmosphere are nearly the

same for the various models shown in Figure 15

Entry Velocities, Modes, and Vehicle Geometries

The entry velocities encountered for various missions

range from 11 to 15 km/sec. The lowest velocities correspond

to entry from orbit, and the highest velocities are those for a Mercury swing-by mission. For direct entries, as shown by

Norman and Hart (69), the velocities can range from 11.6 to 13 km/sec depending on the year of the mission. Some early studies

considered vertical entries but most recent investigations have

studied entry angles from -30° to -50°. The Venus missions

studied so far have utilized ballistic entries since it appears

that heating rates can be kept within a manageable range without the use of lift and there is no problem in decelerating to subsonic speeds at sufficiently high altitudes for atmospheric

studies

The geometries considered by almost all investigators are,

as was the case for Martian entry, spherically blunted cones. In general, nose radii and cone angles for Venusian vehicles

are smaller than those used for Mars entry. This trend toward

sharper, more highly sweptback vehicles reflects attempts at

772
minimizing the total heat load by reducing radiative heating.

As shown by Norman and Hart (69)^ however not all the effects

of reducing nose radii are favorable. If transition to turbulent flow occurs (in Ref . (69); the assumption of a momentum thickness transition Reynolds number of 250 results in

turbulent flow at the cone edge just prior to peak heating),

then the thick entropy layer caused by a large nose radius

can significantly reduce the levels of turbulent heating and

shear. Most studies have indicated nose radii on the order of

1 foot and cone angles from 50 to 6o*^ (Fig. 1^).

Heating Levels

Studies of stagnation point and laminar convective heating in various atmospheres(70, 71^72, 73^7^) have shown that COg

and CO2/N2 mixtures will produce nearly the same (actually

about 10 percent higher) heat fluxes as air for comparable

vehicles and flight conditions. It is to be expected that

turbulent heating levels will be comparable for these gas mixtures

On the other hand, radiative heating from air and from

CO2/N2 mixtures differs considerably(75^76,77^78) . One reason

for this is that the temperatures at which these gas mixtures

begin to radiate strongly are quite different. For example, at temperatures of approximately 8000° K, where air radiates

weakly, CO2 radiates strongly. Another reason is that because

of the different radiating species involved the spectral distributions of radiation from these gases are quite dissimilar. This

is shown in Figure 17 which was prepared using the radiation

model of Nicolet(79) • The reference conditions for this figure

were selected from Reference (80) with a V-5 atmosphere (55) and

the calculations were carried out for an isothermal slab.

Three gas mixtures are shown: air, 90 percent CO2 with 10 percent N2, and 100 percent CO2. The calculated equilibrium

post- shock temperature varies between 9090° K for air and

95^0° K for 100 percent CO2. Note that there is almost no difference between the CO2 and CO2/N2 spectral distributions

except in the region from ^ .0 to 4.5 eV. Nitrogen and cyanogen

bands dominate in this region so this difference represents

the effect to be expected by the presence of nitrogen in the

atmosphere. This difference, while not negligible, will not make or break a design and, conversely, designs which allow

for a reasonable nitrogen content will not be unduly penalized.

It is evident that CN, while an important radiator, is not the

dominant radiator which it was originally believed to be( 75^81)

for earlier Venus atmospheres which contained as much as 60 percent N2. A comparison of both of these atmospheres with the air

results on Figure 17 shows significant differences. First of

773
all, it is noted that the air temperature is significantly lower

than either of the others so that the radiation is expected to be lower. However, the difference shown indicates the importance

of C0(^+) which dominates the spectral region from approximately

5 to 11 eV. In air there are no comparable radiators in this

region and, hence, the spectral distribution and radiative flux

encountered during a comparable Earth entry will be significantly different from any Venusian entry.

Figure 17 suggests the importance of using an ablator which

will absorb the C0(4+) radiation. As shown in Figure 10 for

air, the absorption by CO in ablation products is very noticeable in the spectral region from 5 to 11 eV. However, for air

there are no prominent radiation sources in- this frequency band

and the effectiveness of the ablation products is minimized.

These results suggest, however, that any ablator chosen for a Venus entry should be tailored to provide large quantities of

CO in the ablation products so as to minimize the radiative

heat load. In Table 3^ heating rates calculated by various investigators(60, 69,75^80) for various Venus missions are presented.

From this table, it can be seen that the maximum heating

rates for out- of- orbit and direct entries are in the same range

as those for high velocity Earth entry missions. For the

out-of-orbit and direct entries, it is also seen that the pressures and aerodynamic shears are higher than those characteristic of manned entries but much less than those typical of

high M/C]y\. ballistic Earth entry vehicles. The duration of

heating for these Venus entries is relatively short (on the

order of 50 sec)

Candidate Materials

The levels of heating rate, pressure, and shear presented

in Table 5 can be used to select a class of ablators for Venus

entry missions. First of all, the high heating rates shown in

Table 5 prohibit the use of silica-dominated materials such as those used for Martian entry vehicles. At high surface

temperatures, chars which contain large amounts of SiOg will

recede rapidly due to Si02 melting, Si/C reactions, or both. At these temperatures, char surface reradiation will be a primary energy accommodation mechanism and so a high emissivity, stable carbonaceous char layer is required. All of

this dictates the use of a carbon-dominated material. As mentioned previously, the pressures and shears associated

with Venusian entry are somewhat higher than those typical of

manned entries. Because of this, mechanical char failure may

become a significant consideration, at least for selected

regions of the entry vehicle. In Earth entry research the

usual approach has been to increase ablator density in an attempt to obtain increased char strength. To a degree this

774
has been successful, but large increases in ablator density have

been found to result in only small changes in the test conditions required to produce mechanical char removal. As mentioned previously, the most commonly encountered type of

mechanical char failure involves a weakening of the char

skeleton by chemical reaction followed by removal of discrete

char fragments by aerodynamic forces . This is not completely

understood and is not analytically describable at the present

time. It is probable that an ablator of somewhat higher density

than those used on manned vehicles should be used on a Venusian

entry vehicle. A density of about 0.7 gm/cm5 seems a reasonable

choice. The use of high-density materials is not desirable

since their resistance to mechanical char failure is only

slightly greater than that of the lower density materials, and

their inferior insulating efficiency (resulting from their high

density) would require large increases in heat-shield weight.

Hence, a carbon-dominated material of medium density is

indicated. Some of the best-known materials of this type may, however, be unsuitable for planetary missions. Most phenolicbased and many epoxy-based materials have been found to be

incapable of withstanding the prelaunch and transit environ- ments of sterilization, vacuum, and cold- soak which were so influential in the choice of an elastomeric material for

Martian entries. A possible approach may be to use a material composed of an elastomeric resin, carbon microspheres,

and carbon fibers in an effort to utilize the good lowtemperature properties of the resinj and with the high carbonfiller content, produce a stable high- temperature char. It would be interesting to know what type of ablator the Russians

used on Venera k-

Present State of the Technology

Since the levels of heating, pressure, and shear during

Venusian entry are significantly higher than those for Martian

entry, the heat- shield analysis and design is a more critical

part of the overall vehicle design. Except for the most severe entries (Mercury swing-by) , however, the entry conditions are

close enough to those of high-speed Earth entries that the

technology already developed for Earth entry is largely applicable. The situation then is that while there are significant

uncertainties in both heating and ablation analyses, it appears

that, through conservative design, workable heat shields can be

produced within the present state of the art. Present indications are that for out -of- orbit and direct entries the thermal

protection system will constitute from 10 to 25 percent of the

entry vehicle weight. (A recent study by Jaworski and

Nagler(82) indicates that, for the less severe Venusian

entries, the weight of the actual ablation material may be less

than 5 percent of the total vehicle weight. These numbers seem

small and bear further study.)

775
In the analyses of aerodynamic heating, there are significant uncertainties associated with the prediction of boimdaiylayer transition, turbulent heating, and radiative heating

levels. In predicting ablator response, the major uncertainties appear to be related to the kinetics of the pyrolysis gas reactions, ablator response to turbulent flows, the prediction

of char recession rates, particularly recession due to mechanical char removal.

The prediction of boundary-layer transition has, for many

years, been recognized as one of the most important problems

in aerospace research. It is also one of the most difficult. In view of the great amount of research already devoted to the

problem, it seems unrealistic to expect any breakthroughs in

this area. Research should, of course, be continued, but in

designing near- future planetary entry vehicles, we may have to accept transition Reynolds numbers with order-of-magnitude

accuracies. On the other hand, it appears that the present

uncertainties in radiative heating levels could be reduced

by carrying out completely coupled analyses, such as those

described earlier for high-speed Earth entry, in which all the

significant phenomena such as self-absorption, radiation

cooling, realistic (lines, band systems, and continuum) radiation models, and radiation blockage by pyrolysis gases are accounted for. In practically eveiy study carried out so far

for Venusian entry, one or more of these phenomena have been

neglected. With regard to the uncertainties in ablator

response, the greatest need is for more and better experimental

data to better define the important mechanisms and to lead to more physically realistic analyses.

Jupiter

Atmospheric Characteristics

Our knowledge of the atmosphere of Jupiter is much less

complete than for Mars or Venus. Because of the Jovian cloud

layers, the planet's surface (if one exists in the usual

sense) cannot be observed. The available data concern the

atmosphere above the cloud layer. The makeup of the Jovian

lower atmosphere is largely a matter of speculation. In 19^7,

Michaux, et al. (57) , reviewed the status of our knowledge of

Jupiter. They presented a possible makeup for the lower

atmosphere and Jovian interior (based on models by Gallet and

Peebles) which is presented in Figure l8 of the present paper.

From this figure it is seen that the atmosphere is pictured

as becoming denser and denser until a surface which may be

liquid or solid is reached. From the standpoint of reentry

vehicles, - however, it is the atmosphere above the clouds that

is of most- interest since vehicles with moderate m/C-^A's will

decelerate to terminal conditions before reaching the clouds.

776
Table k presents a sampling of proposed Jovian model

atmospheres (56,83) . As shown in Table 4, there is general

agreement that the main atmospheric constituents (above the

cloud layers) are H2 and He, but the proposed compositions

range from predominantly H2 to predominantly He. Proposed

cloud-top pressures and temperatures range from 2 to 2k atmospheres and 15O to l68° K> the atmospheric scale

heights range from 12 to 21 km.

Entry Velocities, Modes, and Vehicle Geometries

For Jupiter, escape velocity is approximately 60 km/sec

and surface satellite speed is approximately ^0 km/sec.

Accordingly, most studies of Jovian entry have considered

this range of entry velocities. The equatorial velocity of

the planet is, however, about 13 km/sec and, hence, entry in

the direction of planetary rotation and in the vicinity of

the equatorial plane could reduce the relative direct-entry

velocities to about 50 km/sec.

Most studies to date have considered ballistic entries and,

again, the most widely considered configuration is the sphere

cone. Because of the very high speeds and heating rates involved in Jupiter entry, there are some indications that

guided (lifting) entries may be required to keep the heating

within manageable limits (85) . Tauber(83) has studied the

relative merits of blunt and conical vehicles for Jovian entry.

His conclusion is that the conical vehicles have a clear

advantage with regard to required heat- shield weights. In

fact, he suggests that if the atmosphere were primarily H2, a blunt (Apollo-like) vehicle might have to be composed almost

entirely of heat shield. Present indications are that the cone angles for Jovian vehicles will be smaller than those for

Martian and Venusian vehicles (Fig. 1^). The most advantageous

cone angle and, for that matter, the overall character of the

entry heating appear to depend strongly on the atmospheric

composition. This dependence will be discussed subsequently.

Heating Levels

Heating rates, for most Jovian entries, are predicted to be

extremely high by Earth-entry standards. Tauber(85) studied the

effects of vehicle geometry and atmospheric composition on radiative and convective heating rates for a spherically tipped, conical

vehicle. Some of his results are presented in Figure I9. The

results presented in Figure I9 were computed for a vehicle

having a 30° cone half-angle, a base radius of 1 m, and a nose

radius of 1 cm entering on a shallow ballistic trajectory at 50 km/sec. During entry, boundary- layer Reynolds numbers were

limited to 5 x 10^ in order to insure laminar flow. Both

radiative and convective heating rates are presented for the

stagnation point and the conical afterbody, as a function of

777
atmospheric H2 content. As shovn by Figure I9, a 100-percent

H2 atmosphere would produce heating rates that are sizable

but comparable to those experienced during severe Earth

entries. A He atmosphere, on the other hand, would produce

radiative and convective heating rates far beyond those with

which we have had experience. At the present time, it is difficult to pick a "most likely" Jovian atmosphere. The most recent papers on the subject do, however, seem to be indicating

higher and higher H2 contents. In 19^3^ Spinrad and Trafton(8^)

proposed a 60-percent Ho atmosphere and, in 19^7^ Beckman(85)

proposed an 82-percent H2 atmosphere. If these high H2 contents prove to be correct, the Jovian entry heating problem will be

large but not unmanageable

The theory of laminar convective heat transfer for H2/He

mixtures appears to be on fairly firm ground. Stagnationpoint heating in H2 has been treated by Scala(86) and Marvin

and Deiwert(7l)^ and correlated by Zoby(70) . Stagnation-point

heating in He was measured and correlated by Pope (87)

Zoby(70) has proposed an approximate method (involving sum- mation of terms weighted according to the mass fractions of

the individual gases) for calculating stagnation-point heating

in gas mixtures, and recent unpublished work by Sutton and

Graves at the Langley Research Center has shown that the sum- mation approximation gives accurate estimates of heating for

Jovian atmospheres. With regard to laminar heating away from

the stagnation point, Marvin and Deiwert(7l) found that heatingrate distributions were only affected to a minor degree by

gas composition. The prediction of transition and turbulent

heating in H2/He mixtures involves large uncertainties, just

as it does in the case of air.

The radiation from H2/He shock layers is significantly

different from that produced by air. This is illustrated by

Figure 20. In Figure 20, spectral flux is presented as a function of photon energy for a 6l-percent H2^ 36-percent He shock layer, and for an air shock layer. The results shown

for the Jovian atmosphere were taken from the work of

Stickford and Menard (88) . They were obtained by carrying

out an adiabatic calculation, which included self-absorption,

for a 12 000° K, 1-cm- thick layer. The results shown for

air are those previously presented in Figure 10 (a) and

include self-absorption, radiation loss, and, as indicated

by the short dashed curves, absorption by ablation products.

The H2/He results include Lyman and Balmer line radiation

while the air results are for continuum only. However, as Figure 10(b) indicates, the line radiation from the air will

not change the comparison shown in Figure 20. The temperature of the air shock layer (l2 900° K) is comparable to that of the H2/He layer, but the air layer is nearly an

order of magnitude thicker. From these curves it is seen

that, for comparable shock-layer temperatures, the intensity

778
of radiation will be much greater for the H2/He mixture than

for air. Also^ it can be seen that the relative amount of

radiation in spectral regions not subject to absorption by

ablation products is greater for H2/He than for air. This

suggests that absorption by ablation products may not be as important for Jovian entry as it is for Earth or Venus entry.

Wilson(l5) showed only about a 10-percent reduction in radiative heating due to ablation products.^

Candidate Materials

The extremely high heating rates and relatively high

pressures that are predicted for Jovian entry (see Fig. I9)

immediately suggest the use of high-density ^ carbonaceous

materials. Graphites, carbon-carbon composites, and highdensity graphite fiber reinforced plastics have all been

suggested for use on Jovian entry vehicles. In the nose

regions of the vehicle, such materials will probably be used.

The requirement of reasonably small shape change, alone, may

dictate high-density materials in this region. On the vehicle

afterbody, moderate density ablators appear feasible.

Materials similar to those discussed for Venusian entry may

work well for these applications.

Actually, there are several basic phenomena which must be

investigated before any reasonable choice of materials can be

made. The first of these concerns the response of materials

to very high radiative heating rates. At present, no realistic

tests have been carried out at heating rates over a few

kW/cm^. It has been suggested that materials may be torn

apart by the high thermal stresses and internal gas pressures

that heating rates on the order of tens of thousands of W/cm^

will produce. Such a possibility cannot be discounted. What

is needed is test data at these conditions. Second, there are questions concerning the absorptivity of ablative char layers

in the different spectral ranges. Wilson and Spitzer(^3)

measured the absorptance of several chars and graphites at temperatures up to 3^00° K and over the visible and near- infrared wavelength ranges. They found that, for some chars,

the assumption of gray-body behavior was invalid. In spite

of such findings, most ablation analyses utilize constant values for char surface emissivity and absorptivity. For

Jovian entry, large amounts of ultraviolet radiation from the

shock layer are predicted (see Fig. 20). In the case of air

shock layers, much of the ultraviolet radiation is absorbed

by ablation products. It now appears that such will not be

the case for Jovian entry. It has been suggested that the

_Recent calculations by the same author, using new abosrption coefficient data for the polyatomic carbon species have

indicated reductions in radiative heating of up to 80 percent.

More work in this area is needed.

779
high energy, short wavelength ultraviolet radiation might

penetrate deeply into the ablator rather than being absorbed at the surface as most ablation analyses currently assume. Wilson

ans Spitzer's data show increasing absorptance as ultraviolet

wavelengths are approached and seem to suggest high ultraviolet

absorptance values but absorptance measurements in the ultraviolet are not presently available. If the radiation should

penetrate into the char, present analyses could give grossly

incorrect predictions of ablator response during Jovian entry.

Finally, there is the question of mechanical char failure in

H2/He atmospheres. As mentioned previously, the most commonly

observed type of char failure is intimately associated with

oxidation reactions at the char surface. When tested in inert

streams, even low-density chars resist mechanical failure up to high levels of pressure and shear. If the Jovian atmosphere is

essentially 100 percent He, mechanical char failure may not be

a big problem even at relatively high pressures and shears. If

large amounts of H2 are present, however, it is possible that

H/C reactions at the char surface could cause a type of char

failure similar to that caused by O/C reactions in air.

Finally, the extremely high heating rates associated with

He atmospheres may force us to reorient our thinking about the

relative importance of. the various ablative energy accommodation mechanisms. For severe Earth entries and for Venusian

entries, we found that char surface reradiation is likely to be

the most important single mechanism. As pointed out previously,

however, the amount of energy that can be reradiated is limited

(by the sublimation of carbonaceous materials) to around

1200 W/cm^, Hence, if the radiative heating rates reach levels

of 12 000 W/cm2, char surface reradiation will be relatively

unimportant while sublimation will become relatively more

important. It has been suggested that, for such extreme

conditions, materials with high surface reflectivity (to

reflect the incident radiation) rather than high emissivity

might be desirable, if such materials can be found.

State of Technology

Present indications are that the percentage of vehicle

gross weight required for the heat shield will be significantly

greater for Jovian entry than for Earth, Venus, or Mars

entries. Jovian heat-shield technology is in an early stage

of development and many fundamental phenomena require study

before the Jovian entry problem can be accurately assessed.

Simulation Ground and Flight Testing

Ground Testing

While many types of ground test facilities were used in

the early days of ablation research, the great majority of

780
present-day ablation tests are carried out in arc-heated wind

tunnels. A typical arc tunnel, equipped to provide combined

convective and radiative heating, is shown schematically in

Figure 21. The test gas is passed through an electric arc where it is heated to high temperatures. It subsequently

expands through a convergent-divergent nozzle and exits as a supersonic test stream in a free jet test section. The model

to be tested is mounted on a swing arm which allows it to be

inserted in the test stream after stable arc- jet operating

conditions have been achieved and to be retracted from the

stream after the desired test time has elapsed. The flow

leaves the test section through a diffuser and flows through a heat exchanger to either a vacuum sphere or a steam ejector

which provides the pressure ratio required to fill the supersonic nozzle. Arc heaters have been developed which operate

well at high arc chamber temperatures (enthalpies) and low

arc chamber pressures. Other heaters have been developed

which operate at high chamber pressure but at modest temperatures. It has not been possible, however, to develop heaters

which operate at both high temperatures and pressures, nor

has it been possible to build a single heater that operates

well over the entire range of possible conditions. Hence, most test complexes have several arc heaters, with each heater

equipped with several nozzles of varying expansion ratio to provide a reasonably wide range of test conditions.

As shown in Figure 21, some test facilities are equipped

with radiation sources (usually high-pressure arc lamps or lasers) so that the test models can be exposed to combined

convective and radiative heating. The range of test conditions currently available is shown in Figure 22(89,90). This

range of test conditions was determined for tests conducted

in air and there is some question as to whether this extensive a range of test conditions is available in gases such as

C02f and He. For many years there were persistent reports

of electrode failures resulting from CO2 operation, but most facilities now seem to be capable of routine operation on

CO2 and the test conditions obtained are generally comparable

to those obtained with air. All arc tunnels seem to run very

well on He but, because of the safety problems involved, few

tests have been run to date in H2 or H2/He mixtures. Some

selected values of available radiative flux capabilities are also shown on Figure 22 to give an indication of the combined

heating capabilities now available. Also presented on Figure 22 are representative Mars, Venus, and Jupiter entry

conditions(89) . As can be seen from Figure 22, our present

ground- test facilities can produce conditions typical of

Martian entry but not those typical of Venusian entry.

Needless to say, Jovian entry conditions are considerably

beyond our present test capabilities.

781
Even within the range of available test conditions, however, ground facilities do not actually simulate reentry

flight conditions. This lack of simulation can be illustrated

by considering the stagnation-point heating relation

Ground- test models are usually much smaller than the flight

vehicles they represent. Hence, if the heating rate is the

same for both, then either the pressure or the velocity must

differ. This illustrates one of the basic difficulties of

ground ablation testing. One or two aspects of a given flight

condition can be duplicated, but when they are, other aspects,

often important ones, must be in error. Because of this lack

of complete simulation, ground-test results cannot be used

directly to predict in-flight ablative behavior. Instead,

the ground tests are used to define basic ablation mechanisms

and to verify a theoretical ablation model (usually a digital

computer code) . The theoretical ablation model is then used

to predict in-flight behavior.

Another problem associated with the small size of ablation test models is that multidimensional ablation effects

can become important in ground tests even though they are negligible for the flight conditions that the ground tests are supposed to represent. One example of such an effect

is the reduction in transpiration cooling effect caused by

the pyrolysis gases flowing laterally through the char to the

low pressure regions around the shoulders of the model rather

than flowing perpendicular to the char surface as is assumed

in one- dimensional ablation analyses. This effect, which can result in increased ablation rates, is described in

Reference (25) . Another effect of this type arises when

lateral temperature gradients (due to small model size)

become large enough so that significant amounts of heat are conducted laterally away from a given point on the char surface

rather than perpendicularly to the char surface into the

decomposing ablator below.

In Figure 22 it was seen that present radiative heating

facilities cannot produce fluxes as high as those expected

during advanced (especially, Jovian) planetary entries. These

facilities also have another shortcoming which is possibly more important than the insufficient flux levels. That is, the

spectral distribution of radiation from the lamps or lasers

differs greatly from that of the gases in a radiating shock

layer. This is illustrated in Figure 23. From this figure

it is seen that the facilities are particularly deficient in

the short wavelength ultraviolet region which constitutes a

782
\

significant part of the shock-layer radiation. As mentioned

before, it has been often suggested that ablation materials

may respond differently to ultraviolet radiation than to visible or infrared radiation. It is possible that the ultraviolet may penetrate deep into the ablator rather than being

absorbed near the char surface (as is assumed in practically

all existing ablation computer programs) . Tests of models

under ultraviolet radiation are needed to resolve this question. Another important phenomenon that cannot presently be

studied in ground tests is the absorption of shock-layer

radiation by ablation products. As was shown earlier, this

energy absorption mechanism is quite significant for Earth

entry and may be significant for ma.ny planetary entries

(especially entries into predomdnantly CO2 atmospheres).

Since the incoming radiation is absorbed preferentially rather

than uniformly across the spectrum (actually most of this

absorption is in the ultraviolet), this phenomenon cannot be

studied unless ground facilities with spectral distributions

similar to those of an actual shock layer can be developed.

Flight Testing

Flight tests have played an invaluable role in the development of Earth- entry ablation technology. Flight tests are expensive and require large project teams and long lead times

but, because of the previously mentioned difficulties in

scaling ground- test ablation data to flight conditions, they

are mandatory for cases (such as manned entry) where the

ablator response must be known with great accuracy.

The great advantage of flight testing is the ability to expose the heat shield simultaneously to the correct levels of

all the important entry parameters (i.e., heating rate, pressure, enthalpy, shear, etc.).

Since this situation cannot be achieved in ground testing,

it usually happens that a given set of ground-test data can be

satisfactorily explained by several theoretical ablation

models, each of which predicts differing in-flight behavior

for the ablator. One of the most significant contributions of

a flight test is that it shows which (if any) of the proposed

theoretical models is capable of predicting both ground and

flight results.

The preceding comments apply when the flight test is

carried out in the atmosphere of the planet for which the

entry vehicle is designed. Suppose we are dealing with a vehicle designed for a Mars or Venus entry. Can meaningful

flight tests of this vehicle be carried out in the Earth's

atmosphere? This problem has been studied by several investigators(80, 91^92) . As far as phenomena which are largely

783
independent of atmospheric composition (i.e., decelerations,

aerodynamic loads, etc.), the answer is yes. If one assumes

a straight line ballistic trajectory(93) and an exponential

atmosphere, it can be shown that, for the case where the

entry velocities and ballistic coefficients are the same, we can select a path through the Earth's atmosphere along

which the vehicle will encounter a density history equal to that which it would encounter along a different path

in a different atmosphere. Beginning with this

argument, Kennet(92) goes on to show that, along this equivalent trajectory, any quantity of the form

"^00 00 can be duplicated so long as m and n do not depend on the

atmospheric composition.

Laminar convective heating can be adequately correlated in terms of ^ but, of course, m and n will, in general,

depend on the atmospheric composition. For air and C02> N2

mixtures, however, the correlations are nearly identical

{- 10 percent difference) and, hence, it appears that for

Martian and Venusian entries, the laminar connective heating

could be adequately simulated. There also appears to be a possibility of simulating transition and turbulent heating in

CO2-N2 atmospheres, but this question has not been completely

resolved at present.

Radiative heating, on the other hand, does not lend itself

to simulation by these techniques. Since many of the most significant radiative heating phenomena ( self-absorption,

absorption by pyrolysis gases, possibly ablator response) are strongly dependent on the spectral distribution of the radiative flux, which is, in turn, strongly dependent on the

chemical composition of the shock layer, this is to be expected.

In Reference (80), Spiegel, Wolf, and Zeh carried out a detailed study of the degree to which the ablative response of

a Venus-entry-vehicle heat shield could be studied and simulated in an Earth entry flight test. They found that convective heating could be simulated quite well, but that the

radiative heating differed significantly, both with regard to flux level and spectral distribution, from that calculated for

an actual Venus entry. Spiegel, Wolf, and Zeh went on to compute the in-depth response of a high-density phenolic- nylon

heat shield for the Venusian and equivalent Earth entries and

found that the ablator response was quite similar in spite of

the differences in radiative heating. This would seem to imply

that such an Earth entry might provide a proof test for a Venusian- entry heat shield. Such a conclusion, however, is

predicated on the assumption that the Venusian- entry (and Earthentry) radiative heating and the response of the ablator to it

784
has been accurately described by the theoretical models used in

the analysis. This is precisely the assumption that we wanted

to check with a flight test in the first place.

Hence_, it appears that the simulation of planetary- entry

heat- shield response by an Earth-entiy flight test is only

feasible for missions involving only convective heating. An

out-of-orbit Martian entry is an excellent example of such a mission. Because of the low heating rates involved and the

fact that the most likely candidate materials (filled elastomers) are not greatly sensitive to shock layer composition at

the resulting low surface temperatures^ it appears that nearly

complete simulation could be achieved. For higher velocity

Mars-entry missions and for missions to Venus and the giant

planets, however, Earth entry flight tests do not appear capable

of simulating the significant radiative phenomena.

Concluding Summation

A review of ablative heat- shield technology for planetary

entries has been carried out to assess the present state of

the art and to identify areas in which further research is

needed

To identify important phenomena and to' provide a basis

for comparison, research on Earth- entry heat shields was summarized as a first step in the review. Emphasis was placed

on entries characterized by large radiative and convective

heating rates, since less severe entries can be considered as limiting cases (the case of negligible radiative heating, for

instance) of this more complex problem.

It was shown that realistic assessments of heating and

ablator response for this type of entry required a coupled

analysis in which convective and radiative processes in the

shock layer influence one another, and both the radiative and

convective heating are significantly modified by the introduction of gaseous ablation products into the flow field. From

such analyses it was shown that, for moderate-density charring ablators, the significant energy accommodation mechanisms

are: pyrolysis gas enthalpy increase, char surface reradiation,

and the reduction of convective and radiative heat flux by

the injection of ablation products into the flow field.

Significant uncertainties were shown to still exist with

respect to the chemical state of the pyrolysis gases (which

determines their enthalpy content) and the calculation of

radiative heating rates with absorption by ablation products.

It was pointed out that most present-day ablation analyses

assume that radiation of all wavelengths is absorbed at the

char surface, even though little is actually known about the

785
absorptance of ablative chars at other than visible wavelengths.

Shock-layer spectra were presented that showed large amounts

of radiant energy at vacuum ultraviolet wavelengths . It was suggested that radiation at these wavelengths might penetrate

deeply into the char and cause ablative behavior to be significantly different from that presently being predicted.

The rate at which the ablator is consumed in accommodating

the incident heating is largely determined by the char recession rate which results from chemical reaction with boundarylayer gases, sublimation, and mechanical char failure.

Significant uncertainties were shown to exist in the prediction

of all three of these phenomena.

The prediction of convective heating rates was found to be

on fairly firm ground so long as the flow remains laminar. The

prediction of boundary-layer transition and turbulent heating

rates still involve significant uncertainties, however, with

transition Reynolds numbers being known only to within an order

of magnitude. It was also pointed out that, because of experimental difficulties, the response of ablators to turbulent flow

is a relatively unexplored area of research. For entries into tenuous atmospheres such as those of Mars

and Mercury, it was found that heat- shield design was well

within the present state of the art. Because of the low entry

velocities typical of out-of-orbit entries, radiative heating

was found to be insignificant and convective heating rates were found to be low. It was pointed out that, for these

entries, the most important ablation-material characteristics were low density (for high insulating efficiency) and the

ability to survive prelaunch and transit environments such as sterilization, cold soak, and space vacuum. In considering entries into the atmosphere of Venus, it was found that, because of the comparable entry velocities,

the significant phenomena were essentially the same as those

for planetary return Earth entry. Hence, most of the uncer- tainties discussed for severe Earth entry with combined

radiative and convective heating apply as well to Venusian

entries

Since the Venusian atmosphere is primarily (rather

than air), the spectral distribution of shock-layer radiation

is quite different from that for air and it was pointed out

that most of the radiation from a Venusian shock layer was in

a spectral range where it could be strongly absorbed by ablation products. Since very few coupled analyses (including

absorption by ablation products) have been carried out for

Venusian entries, it was suggested that this might be a fruitful field of research.

786
It was pointed out that the relatively high char surface

pressures and aerodynamic shears which result from ballistic

Venusian entries suggest significant rates of mechanical char

failure. Considerable work is required in the area of char

surface chemical reactions in CO2 test streams, with particular

emphasis on char failure by coupled chemical-aerodynamic

mechanisms. Reference to Earth entry research showed that the

mechanical char failure problem cannot be overcome by simply

increasing the density of the ablator.

Because of the high Venusian- entry heating rates,

elastomeric materials such as those proposed for Mars entries were considered unsuitable for Venusian entries. Because of

their inability to withstand inter-planetary transit environments, many of the materials most often considered for severe Earth entries may also be unsuitable for Venus missions.

Hence, it was suggested that a new class of materials that

could withstand both the transit environments and the severe entries might have to be developed.

Jovian entry was considered as being typical of entries

into the atmospheres of the giant planets (Jupiter, Saturn,

Uranus, etc.). Because of the large mass of the planet

(hence, high escape velocities), Jovian entry speeds and the

resulting heating rates (especially radiative heating rates) were found, to be potentially many times higher than those

for Earth entry. The extent of the increase can not be

evaluated because of uncertainties in the ablation products

and their absorption coefficients, but it is almost certain

that the radiative heating rates will be very large. It was pointed out that the behavior of materials under such

high heating rates may be quite different than at lower heating rates. The possibility of catastrophic failure cannot be

discounted. It was shown that, because the amount of energy

re radiated from the char surface is limited by material

sublimation temperatures, char surface reradiation could be a less important energy accommodation mechanism for Jovian

entries than it is for Earth or Venusian entries. The possible

significance of mechanical char failure was shown to depend

to a significant extent on whether the Jovian atmosphere is

mostly H2 (which will react with carbonaceous materials) or He (which is inert). The state of the art for Jovian entry heat

shields was found to be in an early stage of development.

The test capabilities of present-day ground-based facilities were reviewed and. their inability to completely simulate

entry conditions was pointed out. It was pointed out, however,

that valuable partial simulation is available and, in particular,

small test models can be subjected to convective heating rates

up to the levels experienced in typical Venusian entries. It was also shown that radiative heating rates comparable to those

of Venusian entry should be available soon. It was pointed out,

787
however, that one important aspect of the radiative heating, the

spectral distribution, is not reproduced in any of the present

radiative facilities. Because of this, it is not presently

possible to validate existing theories of radiative heating with

absorption by ablation products. It was shown that Jovian

entry conditions are beyond the capabilities of present ground

facilities

The ability of Earth-entry flight tests to simulate entries

into other atmospheres was reviewed. It was found that those

phenomena, which are independent of atmospheric chemical

composition, could be adequately simulated but that phenomena

which depend strongly on shock-layer composition (such as radiative heating and char surface phenomena) could not be

simulated. Hence it appeared that nearly complete simulation

of an out-of-orbit Mars entry was possible (since char surface

reactions are not too important at the low temperatures

involved), but that the value of Earth entry tests in studying

Venusian and Jovian entry phenomena is limited.

References

Adams, Mac C: Recent Advances in Ablation, ARS Journal,

Vol. 29, No. 9, p. 625, September 1959-

Fay, J. A.> and Riddell, F. R.: Theory of Stagnation Point

Heat Transfer in Dissociated Air, Journal of the Aeronautical Sciences, Vol. 25, No. 2, February I958

Cohen, N. B.: Boundary-Layer Similar Solutions and Correlation Equations for Laminar Heat-Transfer Distribution in

Equilibrium Air at Velocities up to ^1 100 Feet Per Second,

NASA TR R-118, 1961.

Hoshizaki, H.: Heat Transfer in Planetary Atmospheres at Super- Satellite Speeds, ARS Journal, Vol. 32, No. 10, p. I'^kk, October I962.

DeRienzo, P.> and Pallone, A. J.: Convective StagnationPoint Heating for Reentry Speeds up to 70 000 fps Including Effects of Large Blowing Rates, AIAA Journal, Vol. 5^ No. 2, February 19^7

Boison, J. C.) and Curtiss, H. A.: Preliminary Results of

Spherical- Segment Blunt Body Pressure Surveys in the 20-

Inch Supersonic Wind Tunnel at JPL, RAD Tech.

Memo 2-TM-57-77 (Aerod. Sec. Memo No. 152), AVCO Res. and

Advanced Dev. Div., October 9, 1957-

(1)

(2)

(3)

w

(5)

(6)

788
Zoby, E. v.; and Sullivan, E. M.: Effects of Corner

Radius on Stagnation- Point Velocity Gradients on Blunt Axisynmetric Bodies, MSA TM X-IO67, (U),

March I965.

Lees, Lester: Laminar Heat Transfer Over Blunt Nosed

Bodies at Hypersonic Flight Speeds, Jet Propulsion,

pp. 259, April 1956.

Zoby, Ernest V.: Approximate Relations for Laminar

Heat-Transfer and Shear-Stress Functions in Equilibrium

Dissociated Air, NASA TN D-kkdk, April I968.

Zoby, Ernest V.> and Sullivan, Edward M.: Correlation of

Free-Flight Turbulent Heat-Transfer Data From Axisymmetric Bodies With Compressible Flat-Plate Relationships,

NASA TN D-5802, January I967.

Sullivan, E. M.> Erickson, W. D.> Smith, G. L.> and

Suttles, J. T.: Some Aspects of Interplanetary EarthEntry Simulation, Presented at the 15th Annual Meeting

of the Institute of Environmental Sciences, Anaheim,

California, April 21-24, I969.

Anderson, John D., Jr.: An Engineering Survey of Radiating Shock Layers, AIAA Journal, Vol. 7, No. 9, pp. I665-1675

September I969.

Hoshizaki, H.; and Wilson, K. H.: Convective and Radiative Heat Transfer During Superorbital Entry," AIAA

Journal, Vol. 5, No. 1, pp. 25-35, January 1967.

Page, W. A.> Compton, D. L.> Borucki, W. J.> Ciffone,

D. L.> and Cooper, D. M.: Radiative Transport in

Inviscid Nonadiabatic Stagnation-Region Shock Layers,

AIAA Paper 68-784, AIAA Third Thermophysics Conference,

Los Angeles, California, I968

Wilson, K. H.: Massive Blowing Effects on Viscous,

Radiating, Stagnation-Point Flow, AIAA Paper 70-205,

AIAA Eighth Aerospace Sciences Meeting, New York, N.Y.,

January 1970.

Olstad, W. B.: Blunt-Body Stagnation Flow With Nongray

Radiation Heat Transfer - A Singular Perturbation Solution, NASA TR R-295, 1968.

Chin, J. H.: Radiation Transport for Stagnation Flows

Including the Effects of Lines and Ablation Layer, AIAA

Paper 68-664, AIAA Fluid and Plasmadynamics Meeting,

Los Angeles, California, June I968

789
Rigdon, W. S.> Dirling, R. B., Jr.; and Thomas, M.: Stagnation Point Heat Transfer During Hypervelocity

Atmospheric Entry, MSA CR-1^62, February 1970.

Wilson, K. H.: RATRAP - A Radiation Transport Code,

LiMSC 6-77-67-12, Lockheed Missiles and Space Company,

Palo Alto, California, March 1^, I967.

Callis, Linwood B.: Time-Asymptotic Solutions of Flowfields With Coupled Nongray Radiation About Long Blunt

Bodies, AIAA Paper 7O-865, AIAA Fifth Thermophysics

Conference, Los Angeles, California, June 1970.

Hurwicz, Henryk: Aerothermo chemistry Studies in Ablation. Presented at Fifth AGARD Colloquium, on High

Temperature Phenomena, Braunschweig, Germany,

April 9-13, 1962.

Moyer, C. B.j Anderson, L. W.> and Dahm, T. J.: A

Coupled Computer Code for the Transient Thermal Response

and Ablation of Noncharring Heat Shields and Nose Tips,

NASA CR-1630, 1970.

Bush, H. G.> and Dow, M. B.: Multidimensional Gas Flow

Through Permeable Char Layers and Its Effects on Ablation, NASA TR R-296, January 1969-

Swann, Robert T.> Pittman, Claud M.; and Smith, James C:

One- Dimensional Numerical Analysis of the Transient

Response of Thermal Protection Systems. NASA TN D-2976,

Septeniber I965.

Swann, Robert T.: Approximate Analysis of the Performance of Char-Forming Ablators. NASA TR R-195, June 1964.

Kendall, R. M. ; Rindal, R. A.; and Bartlett, E. P.: Thermochemical Ablation. AIAA Paper No. 65-6^2. AIAA

Thermophysics Specialist Conference, Monterey, Calif.,

September 13-15^ 1965-

Hillberg, Lauri H.: The Convective Heating and Ablation

Program (CHAP). The Boeing Co., Seattle, Washington.

Code Ident. No. 81205, No. D2-56402-1, April I966.

Kratsch, K. M. ; Hearne, L. F.; and McChesney, H. R.: Theory for the Thermophysical Performance of Charring

Organic Heat Shield Composites. Report No. 803099,

Lockheed Missiles and Space Co., October 18, 1965-

Wilson, R. G.: Thermophysical Properties of Six Charring

Ablators From 1^0° K to 700° K and Two Chars From 8OO to 3000° K. NASA TN D-2991, October I965.

790
Brazel, J. P.> Tanzilli, R. A.> and Begany, A. R.: Determination of the Thermal Performance of Char Under

Heating Conditions Simulating Atmospheric Entiy. AIAA

Paper No. 65-6^0. AIAA Thermophysics Specialist

Conference, Monterey, California, September 13-15^ 1965*

McLain, Allen G.; Sutton, Kennethj and Walberg, Gerald D,;

Experimental and Theoretical Investigation of the Ablative

Performance of Five Phenolic- Nylon- Based Materials.

NASA TN D-^5T^, 1968.

Sykes, George F.) and Nelson, James B.: Thermoanalysis

of Ablation Materials. Presented at the Am. Institute

of Chemical Engineers Meeting, Houston, Texas,

February 19-23, I967.

Lundell, J. H.> Wakefield, R. M.> and Jones, W. J.: Hypervelocity Heat Protection. AIAA Journal, Vol. 3, pp. 2087-2096, 1965.

Pike, R. W.> April, G. C.j and Del Valle, E. G. : Nonequilibrium Flow and the Kinetics of Chemical Reactions in

the Char Zone, Reacting Fluids Laboratory, Louisiana

State University, NASA CR-66455, July I5, 19^7

Lewis, J. C.> Floyd, I. J.> and Cowland, F. C: A

Comparative Study of the Gaseous Oxidation of Vitreous

Carbon and Various Graphites at 1500-3000° F. Paper

presented at the Eighth Biennial Conference on Carbon,

Buffalo, N.Y., June I967.

Strickland-Constable, R. F.: Theory of the Reaction of

Graphite With Oxygen in the Temperature Range

1000-2^00° C. Paper presented at the Second Conference

on Industrial Carbon and Graphite, London, April 19^5'

Dolton, T. A.j Maurer, R. E.> and Goldstein, H. E.: Thermodynamic Performance of Carbon in Hyperthermal

Environments. AIAA Paper No. 68-75^, Presented at AIAA

Third Thermophysics Conference, Los Angeles, California,

June 2^-26, 1968.

Kratsch, K. M.> Martinez, M. R.; Clayton, F. I.> Greene,

R. B.> and Wuerer, J. E.: Graphite Ablation in HighPressure Environments. AIAA Paper No. 68-II53, Presented

at the AIAA Entry Vehicle Systems and Technology Meeting,

Williamsburg, Va., December 3-5, I968.

Scala, S. M.> and Gilbert, L. M. : Thermal Degradation of

a Char- Forming Plastic During Hypersonic Flight. ARS

Journal 32, Vol. 32, pp. 917-92^1-, I962.

791
Mathieu, R. D.: Mechanical Spallation of Charring Ablators in Hyperthermal Environments. AIM Journal,

Vol. 2, No. 9, September 196^^. Maahs, H. G.; and Schryer, D. R.: Particle Removal in

the Ablation of Artificial Graphite. AIAA Journal,

Vol. 7, No. 11, November I969, pp. 2178-2179.

Sutton, Kenneth: Results of an Experimental Ablative

Study of a Carbon-Phenolic Material. Proposed NASA TN

(L-6455)

Wilson, R. G.> and Spitzer, C. R.: Visible and Near-Infrared

Emittance of Ablation Chars and Carbon. AIAA Journal,

Vol. 6, No. k, April I968.

Walberg, Gerald D.: An Analytical Study of Diffusion

Controlled Char Oxidation and its Effect on the Steady

State Ablation of Hydrocarbon Plastics. NASA TR R-242,

1966.

Wakefield, R. M.; Lundell, J. H.j and Dickey, R. R.: Effects of Pyrolysis-Gas Chemical Reactions on Surface

Recession of Charring Ablators, AIAA Paper 68-302,

Ninth AIAA/ASME Structures and Materials Conference,

Palm Springs, California, April 1, I968. Also J. Spacecraft, Vol. 6, February I969, pp. 122-128.

McLain, Allen G.: Investigation of the Char Formation

and Surface Recession Characteristics of a Silicone

Resin-Based Composite Ablation Material. Proposed

NASA TN (L-6559)

Vojvodich, Nick S.: Hypervelocity Heat Protection - A

Review of Laboratory Experiments. J. Macromol.

SCI. CHEM., A5(5), pp. 567-39^, May I969.

Hoshizaki, H.j and Lasher, L. E.: Convective and Radiative Heat Transfer to an Ablating Body. AIAA Paper

No. 67-327, Presented at the AIAA Thermophysics Specialist

Conference, New Orleans, Louisiana, April 17-20, I967.

Coleman, W. D.) Hearne, L. F.> Lefferdo, J. M.> and

Vojodich, N. S.; A Study of the Effects of Environmental

and Ablator Performance Uncertainties on Heat Shielding

Requirements for Blunt and Slender Hyperbolic-Entry

Vehicles. AIAA Paper No. 68-15^. AIAA Sixth Aerospace

Sciences Meeting, New York, N.Y., January 22-24, 1968.

792
Smith, G. L.', Suttles, J. T.> Siillivan, E. M.> and

Graves, R. A., Jr.: Viscous Radiating Flow Field on an Ablating Blunt Body. AIAA Paper No. 70-218, AIAA

Eighth Aerospace Sciences Meeting, New York, N.Y.,

January 19-21, 1970.

Suttles, John T.: A Method of Integral Relations

Solution for Radiating Nonadiabatic Inviscid Flow Over

a Blunt Body. NASA TN Also, M.S. Thesis,

Aerospace Engineering, Virginia Polytechnic Institute,

Blacksburg, Virginia, March I968.

Garrett, L. B.> Suttles, J. T.> and Perkins, J. N.: A Modified Method of Integral Relations Approach to the

Blunt-Body Equilibrium Air Flow Field, Including

Comparisons With Inverse Solutions, NASA TN

September I969.

Vaughan, Otha H., Jr.: Model Atmospheres of Mercury,

J. Spacecraft, Vol. 6, No. 10, pp. 1171-1175;,

October I969.

Michaux, C. M. : Handbook of the Physical Properties of

the Planet Mars. NASA SP-503O, I967.

NASA Space Vehicle Design Criteria (Environment): Models

of Venus Atmosphere (I968), NASA SP-80II, December I968

Sodek, Bernard A., Jr.: Jovian Occultation Experiment.

J. Spacecraft, Vol. 5^ No. k-, pp. k-6l-k-63, April I968.

Michaux, C. M.: Handbook of the Physical Properties of

the Planet Jupiter. NASA SP-5051, 19^7

Prit chard, E. B.j and Harrison, E. F.: Multipurpose

Entry Vehicle Requirements for Unmanned Landings on Bodies Having Tenuous Atmospheres. J. Spacecraft, Vol. 5, No. 4, April 1968.

Roberts, Leonard: Entry into Planetary Atmospheres.

Astronautics and Aeronautics, October 1^6k.

Strauss, Eric L.> and Sparhawk, H. E.: Ablative Heat Shields for Planetary Entry Vehicles, AAS Paper 68-8-5,

Space Projections From the Rocky Mountain Region,

Brown Palace Hotel, Denver, Colorado, July I5-I6, I968.

Hiltz, A. A.> Florence, D. E.; and Lowe, D. L.: Selection, Development, and Characterization of a Thermal

Protection System for a Mars Entry Vehicle. AIAA Paper

No. 68-50^. AIAA/ASME Ninth Structures, Structural

Dynamics and Materials Conference, Palm Springs, Calif.,

April 1-5, 1968.
Fixler, S. Z.> and Layton, D. M.: Thermo structural and

Material Considerations for Mars Lander Heat Shields

In: Aviation and Space: Progress and Prospectsj

Proceedings of the ASME Annual Aviation and Space

Conference, Beverly Hills, California, June I6-I9, I968

Wagler, Robert G.: The Mars Transit and Entry Environment: A New Problem for Heat Shields. JPL TR 52-11^5.

Reprinted from Proceedings of Institute of Environmental

Sciences, 13th Annual Technical Meeting, Washington D.C.,

April 10-12, 1967, Vol . 2, pp. h-k^-k^'^

Nagler, Robert G.: Ground Simulation of a Mars-EntryCapsule Aeroshell Environmental History. JPL Tech.

Report 52-1^66, February I5, 1970.

Meltzer, J.; Rossoff, J.; and Slaughter, J. I.: Structure and Materials Aspects of the Prime Flight Test

Vehicle. AIAA/ASME Seventh Structures and Materials

Conference, April l8-20, I966.

.Strauss, Eric L.: Response of Superlight Ablators to Various Heat Pulses. AIAA Paper No. 68-301, AIAA/ASME

Ninth Structures, Structural Dynamics and Materials

Conference, Palm Springs, California, April 1-5^ I968

AVCO Corp. Report: Comparative Studies of Conceptual

Design and Qualification Procedures for a Mars Probe/

Lander. Final Report, Vol. V. Subsystem and Technical

Analyses, May 11, I966.

Katz, G. D.> and McMullen, J. C: Entry Vehicles for

Unmanned Planetary Exploration. AIAA Entry Technology

Conference, Williamsburg and Hampton, Virginia,

October 12-1^, 1964.

Norman, Herbert G.> and Hart, Paul M.: Mission Influence

on the Aero thermodynamic Environment for a Venus Entry

Vehicle, AAS Paper 68-4-6, Space Projections from the

Rocky Mountain Region, Brown Palace Hotel, Denver,

Colorado, July 15-l6, I968.

Zoby, Ernest V.: Empirical Stagnation-Point Heat- Transfer Relation in Several Gas Mixtures at High

Enthalpy Levels. NASA TN D-4799, October I968

Marvin, Joseph G.> and Deiwert, George S.: Convective

Heat Transfer in Planetary Gases. NASA TR R-224,

July 1965.

794
Nerem, Robert M.> Morgan, C. Joe; and Graber, Bruce C:

Hypervelocity Stagnation Point Heat Transfer in a Carbon

Dioxide Atmosphere. AIAA Journal, Vol. 1, No. ^, pp. 2173-2175, September I965.

Collins, D. J.> and Horton, T. E.: Experimental Convective Heat-Transfer Measurements. AIAA Journal, Vol. 2, No. 11, pp. 2046-20^7, July 196^.

Gruszczynski, J. S.> Warren, W. R., Jr.; and Diaconis,

N. S.: Laboratory Simulation of Hypervelocity Heat Transfer Problem During Planetary Entry. Space Sciences

Laboratory, General Electric Co. Report r6^SD73« Paper

presented at the 15th International Astronautical

Congress, September 7-12, 1-90^, Warsaw, Poland.

James, Carlton S.: Experimental Study of Radiative

Transport From Hot Gases Simulating in Composition the

Atmospheres of Mars and Venus. AIAA Journal, Vol. 2, No. 5, March 196^.

Menard, Wesley A.> and Thomas, George M.: Experimental

and Theoretical Study of Molecular, Continuum, and Line

Radiation From Planetary Atmospheres. AIAA Journal,

Vol. 6, No. k, April I968

Gruszczynski, J. S.> and Thomas, K. M.: Equilibrium and

Nonequilibrium Radiation in Simulated Planetary Atmospheres. AIAA Paper No. 66-I85, AIAA Plasmadynamics

Conference, Monterey, California, March 2-^, 1966.

Deacon, Howard Jr.> and Rumpel, William: Radiation

Heating Characteristics of Venus Entry. AAS Paper 68-8-6

Space Projections From the Rocky Mountain Region, Brown

Palace Hotel, Denver, Colorado, July 15-l6, I968

Nicolet, William E.: Advanced Methods for Calculating

Radiation Transport in Ablation-Product Contaminated

Boundary Layers. Aerotherm Corp. Final Report No. 69-6I,

December 27;, I969.

Spiegel, Joseph M.> Wolf, Fred> and Zeh, Dale W.: Simula

tion of Venus Atmospheric Entry by Earth Reentiy. AIAA

Paper No. 68-1148, AIAA Entry Vehicle Systems and Technology Meeting, Williamsburg, Virginia, December 3-5,

1968.

McKenzie, Robert L.j and Arnold, James 0.: Experimental

and Theoretical Investigations of the Chemical Kinetics

and Nonequilibrium CN Radiation Behind Shock Waves in

CO2-N2 Mixtures. AIAA Paper No. 67-322, AIAA Thermophysics Specialist Conference, New Orleans, Louisiana,

April 17-20, 1967.
Jaworski, W.> and Nagler, R. G.: A Parametric Analysis

of Venus Entry Heat-Shield Requirements, JPL Technical

Report 32-1^68, April 15, 1970.

Tauber, Michael E.: Atmospheric Entry into Jupiter.

J. Spacecraft, Vol. 6, No. 10, October I969.

Spinrad, H.> and Trafton, L. M.: High Dispersion Spectra

of the Outer Planets. I. Jupiter in the Visual and Red.

Icarus, Vol. 2, pp. 19-28, I963.

Beckman, J. E.: The Pressure at the Cloud Top and the

Abundance of Hydrogen in the Atmosphere of Jupiter.

Astrophysical Journal, Vol. 1^9, No. 2, I967.

Scala, S. M.J and Gilbert, L. M.: Theory of Hypersonic

Laminar Stagnation Region Heat Transfer in Dissociating

Gases. r65SD^O, Gen. Elec. Space Sci. Lab., April I963.

Pope, Ronald B.: Stagnation- Point Heat Transfer in Arc- Heated Helium and Argon. AIAA Journal, Vol. 7, No. 6, pp. 1159-1161, June 1969.

Stickford, G. H., Jr.; and Menard, W. A.: Bow Shock

Composition and Radiation Intensity Calculations for a Ballistic Entry Into the Jovian Atmosphere. AIAA Paper

No. 68-787, AIAA Third Thermophysics Conference,

Los Angeles, California, June 2^-26, I968.

Nagler, Robert G.: A Systematic Review of Heat-Shield

Technology for Extraterrestrial Atmospheric Entry.

JPL Tech. Report 32-1^36, March 15, 1970.

Heister, N. K.> and Clark, C. F.: Feasibility of Standard

Evaluation Procedures for Ablating Materials, NASA

CR-379, 1966.

Beuf, F. G.; Katz, G. D.> and Kern, R. J.: Earth Flight

Test of Mars Entry Vehicles. J. Spacecraft, Vol. 3, No. 4, April 1966.

Kennet, Haimj and Taylor, Roy A.: Earth Reentry Simulation of Planetary Entry Environment. J. Spacecraft,

Vol. 3, No. ^, April I966.

Allen, H. Julian> and Eggers, A. J., Jr.: A Study of the

Motion and Aerodynamic Heating of Missiles Entering the

Earth's Atmosphere at High Supersonic Speeds, NACA

TN ^0^7, 1957.

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104 LOCAL DISPLACEMENT

THICKNESS

TRANSITION

REYNOLDS

NUMBER 103

102

FLIGHT TEST

AIR WIND TUNNEL

BALISTIC RANGE

4 6 8 LOCAL MACH NUMBER

10 12

Figure 1.- Survey of transition data for nona"blating sharp

cones (from Ref. ll)

8x 103r7

6

5 RADIATIVE

HEAT 4 TRANSFER. W/cm2

^'transparent

\ continuum

Qr =(4KpaTs4)6/2

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CODE: —COUPLED —UNCOUPLED O UNCOUPLED

WITH ATOMIC LINES (WILSON AND GRIEF) COUPLED WITH ATOMIC LINES (PAGEETAL)

HOSHIZAKI AND

WILSON (REF. 13)

0 0.5 1.0 1.5 2.0 NOSE RADIUS, m

Figure 2.- Results for stagnation-point radiative heat transfer

reflecting various stages of the state-of-the-art;

V = 15.24 km/sec altitude = 6l km (from Ref. 12).

801
ENTHALPY, h MJ/kgcAS

SUBLIMATION LIMIT

FROZEN

LUNDELL - ARC IMAGE

TEST (.001 atm)(REF. 33) 4x10-

temperature,t,ok

Figure k.- Pyrolysis gas enthalpy as a fanction of chemical

state (from Ref . ^7)

802
CARBON-PHENOLIC l.U<q^ < 13.6 MW/m'

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.24 r-

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ooca o % •

% •

o o§ • •

.01 ?o l o? 1 .10 1.0 10.0 STAGNATION PRESSURE, P^^, atm. Figure 5.- Mechanical char failure for a carbon phenolic

ablator (from Ref. ^2).

100.0

Pt2, atm

GREY CHAR(SiC)

GLASSY CHAR(Si02a ))

BLACK CHAR (Si02(S))

4 6 8 inl 2 HEATING RATE, Pc, MW/m^

Figure 6.- Char surfaces produced by silicon dominated ablator

( from Ref . h6)

803
FORCED INJECTION

Ah

MJ/kg

5-80 VOJVODICH AND POPE

70-116 POPE

5-7 COMPTONetal.

5-46 NASA -SRI

ROUND ROBIN

Figure 7'

1 2 3 4 5 6 NORMALIZED ABLATION RATE, B - riigAh/cic Reduction of convective heating by pyrolysis gas injection (from Ref. ^7)

TRAJECTORY:

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SOLUTION

CONVECTIVE

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TRANSIENT SOLUTION

STEADY STATE SOLUTION

r

COMPUTE

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STATE 3

COMPUTE

YES

STRONG

INJECTION SOLUTION

TRANSIENT

HEAT SHIELD SOLUTION

MODERATE

INJECTION SOLUTION

RATRAP

COMPUTE qR(y)

ACCEPTABLEy

COUPLED / SOLUTION / Figure 8.- Logic for coupled analysis (from Ref. 50).

804
Tx 10'^, °K

Ca

6 14.18 cm

4170 W/cm^

3246 W/cm^

4 W/cm^

143180k

=^ NO ABLATION LAYER *^WITH ABLATION LAYER

0 .2 .4 .6 .8 1.0 BODY y/6 SHOCK

I'igure 9'- Temperature and ablation products profile. Stagnation streamline of 30^.8 cm radius sphere, V^^ = 15.24 km/sec,

P V

61 km altitude, carbon phenolic injection, —«— = 0.2,

T = 3600° K (from Ref. 50 ). ^oo^oo

(W/cm^-ev)

WALL

103^

INV. LAYER EDGE

10 4 8 12 16 20 PHOTON ENERGY, eV

I I I \ I « 4000 2000 1000 620 WAVE LENGTH, A

PR (W/cm^)

INV. LAYER EDGE WALL

NEGLIGIBLE

4 8 12 16 20 PHOTON ENERGY, eV

1002000 1000 620 WAVE LENGTH, A

(a) CONTINUUM (b) LINES Figure 10.- Spectral distribution of radiation heat flux due to

line and continuum processes (from Ref. 50).

805
0 50 100 150 200 250 300 TIME FROM 122 Km ALTITUDE, sec Figuxe 11.- Ablative energy accoramodation mechanism for

phenolic -nylon ablator on Apollo- like planetary- return

Earth entry vehicle (from Ref. 49).
ALTITUDE

z. km

0 50 100 150 200 250 300 350 400 450 500 TEMPERATURE, °K Figure 13-- Temperature profiles for five Martian atmospheric

models

3.5 m hi. 22 mH hi nH

M/CdA ~ 3 gm/cm

MARS

M/C[)A ~ 5 gm/cm VENUS

M/C^A ~ 8gm/cm2

JUPITER

(PROJECT VIKING)

Figure 1^^.- Configurations for typical planetary entry

vehicles

LEGENDMODELS BASED ON RAD I US OF 6048 km — MODELS BASED ON RADIUS OF 6078 km

ALTITUDE, km

10-18 10-16 10-14 10-12 10-10 10-8 io-6 io-4 io-2 loO 10^ DENSITY, gm/cm^

Figure 15.- Density versus altitude for models of Venus

atmosphere (from Ref. 55).

807
SPECTRAL

FLUX, W/cm^eV

REFERENCE CONDITIONS

199.2 OK

- 1.256 mb

5 « .444 cm

100% CO2. T - 9540 0k 90%CO2, 10% N2

T - 9410 °K

PHOTON ENERGY, eV Figure YJ .- Comparison of spectral flux from air and proposed

Venusian atmospheres.

808
GALLET'S (1962) ' JOVIAN ATMOSPHERE

H2, He AND CH4

LIQUID AND/OR SOLID HYDROGEN

TOP OF CLOUD LAYER

(150 Ok. 1 atm)AMMONIA CRYSTALS AMMONIA DROPLETS AMMONIA VAPOR

1 I- ' ICE CRYSTALS J y^VVATFR DROPLETS WATER VAPOR

PEEBLES (1964) JOVIAN INTERIOR

CLEAR CLOUDS

CLEAR

RADIUS = 1

I CLOUD CLOUDS [atmosphere

(IN EXPANDED

SCALE) RADIUS = 0.94 SURFAC- OF PLANET

(2000 Or 2 X 105 atm

400 0k in GALLET'S MODEL)

Figure

RADIUS =0.8, TRANSITION TO METALLIC HYDROGEN

(T - ? P = 106 atm) METALLIC HYDROGEN

RADIUS = ? TRANS ITION TO DENSE CORE OF ROCKY SILICATES AND METALLIC ELEMENTS (PERHAPS 10 TIMES THE MASS OF EARTH)

3.- Jovian cloud layers proposed oy Gallet overlying

Jovian interior oy Peebles (from Ref. 5T)'

809
= 50 km/sec

R 5x 10" r^, = 1 cm

30^ Th = Im

60

HEATING RATE,

q, kw/cm^

STAG. PRESSURE

Ps atni

HEATING RATE,

q, km/cm^

0 20 40 60 80 100 PERCENT HYDROGEN (BY VOLUME)

Figure 19 •- Maximum heating rates and stagnation pressures for

conical "body during shallow ballistic entry^ (a) stagnation

point (ablating nose), (b) flank (at 0.7 of length point).

(From Ref. 83.)

810
SPECTRAL FLUX, W/cm2eV

PHOTON ENERGY, eV Figure 20.- Comparison of typical shock layer spectra for

Jovian and Earth entries.

RADIATION SOURCE AND COLLECTOR

TO VACUUM

SPHERES

TEST-SECTION PRESSURE CONTROL

VALVE

TEST MODEL

SUPERSONIC TEST STREAM

ARC HEATER ^ TEST SECTION ^DIFFUSER Z^AFTER COOLER Figure 21.- Schematic of ground test facility.

811
COLD-WALL

STAG. -POINT CONVECTIVE

HEATING

RATE. W/cm2

105

PEAK HEATING

JUPITER ENTRY-s^^qR - 60.000 W/cm2

GROUND TEST FACILITIES:

1000 W/cm2 RADIATIVE

HEATING AVAILABLE NOW.

10,000 W/cm2 PROJECTED FOR NEAR FUTURE (REF. 89)

O

PEAK HEATING VENUS ENTRY;

= 4500 W/cm2

VENUS ENTRY TRAJECTORY

10-2 10"! loO lOl STAGNATION PRESSURE, atm. Figure 22.- Ground facility test capabilities

lO^r

SPECTRAL

FLUX, lO^h

W/cm2eV

JUPITER (He, Hz) VENUS (C02, N2) AIR (NO ABLATION PROD. AIR (WITH ABL. PROD.)

PHOTON ENERGY, eV Figure 25.- Comparison of typical shock- layer spectra for

Varius planetary atmospheres with spectra for ground test radiation sources.

812
Paper No, 51

ABLATION PHENOMENA IN SUPERSONIC LAMINAR AND

TURBULENT FLOWS

1 2 E. M. Winkler, M. T. Madden, R. L. Humphrey,

and J. A, Koenig"^

REFERENCE: Winkler, E. M. , Madden, M. T.,

Humphrey, R. L. , and Koenig, J. A., "Ablation

Phenomena in Supersonic Laminar and Turbulent Flows,"

ASTM/IES/AIAA Space Simulation Conference, 14-16

September 1970

ABSTRACT: Arc tunnel experiments using pipe specimens of TFE-7 examine the ablation of Teflon under

a supersonic laminar and turbulent boundary layer.

Test conditions include Mach numbers of 2.3 and 3.0,

supply pressures of 2x10^ N/m^ to 3x10^ N/m^ and

supply temperatures of 2200°K to 5000°K. Experimental results are compared with computer predictions.

In two laminar boundary-layer tests, parallel striations appear; in all turbulent boundary-layer tests,

criss-cross striations appear.

KEY WORDS: ablation. Teflon, ablation-induced

transition, cross-hatched striations

NOMENCLATURE

A = cross-sectional area of node

Aq = original cross-sectional area of node

B = factor in density law

Chief, High-Temperature Aerodynamics Group,

Aerophysics Division, Naval Ordnance Laboratory (NOL)

^Aerospace Engineers, NOL

^Aerospace Engineering Technician, NOL

813
E/R = activation energy divided by gas constant

f. = stream function

h = local enthalpy per unit mass

= total enthalpy per unit mass

= mass fraction of the k^^ element

k = thermal conductivity

M = Mach number

m = meter

nig = mass removal rate of gas per unit area which

enters the boundary layer without phase change

N = Newton

P = pressure

t = time

T = temperature _ . u = velocity

X = axial distance into duct

y = coordinate normal to ablating surface with

origin fixed relative to back wall

z = depth, measured from receding surface

p = density

= original density

p^ = final density

r = volume fraction of resin in plastic

T] = transformed normal coordinate in boundary

layer

T = wall shear stress

814
SUPERSCRIPT m ' = reaction order

SUBSCRIPTS

A,B,C constituents

g gas

k element

o supply or original

00 = free stream

INTRODUCTION

The use of ablative materials and the prediction

of their performance for protection of re-entering

vehicles has not always been successful. Some of the

difficulties encountered are traceable to the inability of ground test facilities to accurately

duplicate critical flight regimes, and the lack of

analytical techniques to analyze experimental data

and to make accurate predictions about future

performance. A series of successful experiments have

been performed at the U. S. Naval Ordnance Laboratory

(NOL) in which the ablative performance of Teflon was examined under supersonic laminar and turbulent

boundary layers (1)"^. It is the purpose of this

paper, then, to present the results of these tests

and show their successful comparison with analytical

techniques. Results of skin-friction measurements

are presented for the first time and compared with the

predictions of a computational technique known as the

BLIMP-CMA computer programs (2,3).

Teflon was chosen as the sample material for two reasons. First, and most pragmatically. Teflon is

relatively simple to analyze. It is a homogeneous

material not possessing the complexities of multicomponent ablators, thus reducing the number of ways

in which errors of analysis may arise. Second,

Teflon has practical applications as a desirable

heat shield material (4) . The desirable qualities

include predictable ablative performance, high

specific heat, and low thermal conductivity; detrimental

to its performance are the relatively high density

and low tensile and compressive properties.

The numbers in parentheses refer to the list of

references appended to this paper

815
EXPERIMENTAL PROCEDURE

The experimental set-up is depicted in Figure 1.

Air is heated in the 3 Megawatt four-ring, threephase, ac arc (5) and expanded through an axially

symmetric, contoured nozzle. The emerging supersonic

air is then passed through the TFE-7 Teflon duct, and

exits into a diffuser and exhaust system. The nozzle

as well as the Teflon ducts have been contoured to

insure uniformity of flow by use of a non-ablating

boundary-layer program developed at NOL (6,7) . Pressure data indicate this initial contour maintained

uniform flow throughout the duct for the duration of

the test. The decision to pass air through the model,

rather than over a shape such as a cone or a wedge, was based on the ease of instrumentation of the pipe

specimen. Flight conditions encountered at certain

stations on a re-entering, blunted vehicle can be

duplicated by this arrangement. Air is expanded in

a supersonic nozzle to a Mach number that equals the

local Mach number at the vehicle's station to be

studied; the supply conditions correspond to the

stagnation point conditions on the body in flight.

Even for high free-stream Mach numbers, the local

Mach number may be quite low. Typically for a small

blunted cone re-entering at M = 15, the local Mach

number may be M ~ 3

The test conditions called for supply pressures

of 20 to 30 atmospheres, (i.e., 20 to 30x10^ N/m2)

Mach numbers of 2.3 to 3.0, and testing times of 3 to 13 seconds. The specimens numbered IL through

6L were exposed to a boundary layer predicted to be

laminar, while specimens IT through 5T were exposed

to a predicted turbulent boundary layer. The

complete running conditions are summarized in Table I. All of the ducts were fabricated from TFE-7

polytetraf luoroethylene. The inner diameter was one

inch (25.4 mm), and the wall thickness was 0.75 inch

(19 mm) . A length of six inches (152 mm) was used

for all the ducts; this length had been shown in

prior tests to maintain uniform flow throughout the

length of the duct. This removed the limitations of

internal shock waves. The uniformity of flow was verified by pressure measurements. With the exception

of one duct, all ducts were instrumented with static

pressure taps located 0.25 (6.35 mm), 1.25 (31.75 mm),

2.375 (60.32 mm), 3.5 (88.9 mm) and 4.625 (117.47 mm)

inches from the exit plane of the water-cooled nozzle.

At the X = 4.625 inches location there were two taps

90 degrees (1.570 rad) apart to check for asymmetries

of flow. In addition to these static pressure taps.

816
the Pitot pressure was recorded at the exit of each

duct. For tests requiring a measurement of the Fitot

pressure to verify that the running conditions were

proper, a quick injection mechanism was used to insert

an uncooled probe into the flow centerline. This

mechanism is detailed in Reference 8. The uniformity

of flow was confirmed by a centerline traverse

conducted with a water-cooled probe; the performance

of these probes was not completely satisfactory as their small size hampered effective cooling during

exposure for the duration of the traverse.

For the ablation measurements the ducts were variously instrumented as indicated in Table I. The

ducts were fitted to record surface and in-depth

temperatures. Receding thermocouples were used to measure surface temperature while the in-depth

readings used a thermocouple plug. This plug

arrangement is shown in Figure 2. Wall shear stress

was recorded in four tests. The balance used in

these tests was developed at NOL and is described in

Reference 9. This instrument, which records

continually, uses a floating element head. In

Figure 3, a Teflon element head is shown with the

disassembled balance. Finally, mass loss due to

ablation was determined by comparison of contour

measurements of the ducts before and after each test.

ANALYTICAL PROCEDURES

The analysis of the data was perfoinned with the

assistance of two computer programs. The first of

these, the Boundary Layer Integral Matrix Procedure,

referred to by its acronym BLIMP, (2), allows the

computation of steady-state ablation and laminar

boundary-layer response. The second was the Charring

Material Ablator program, known as CMA (3) . Together

with the CMA, iterative use of the BLIMP allows the

transient response of the ablating laminar boundary

layer and the in-depth response of the material to be

specif led

In particular, the BLIMP, at least the version

used at NOL in the present analysis, performs a nonsimilar solution of the laminar, ablating boundary

layer. This program has great versatility, and the user is free to select from options in six categories:

1. Body shape: axisymmetric or planar

2. Upstream effects: nonsimilar or similar

solution

3. Chemistry: homogeneous, equilibrium, or equilibrium-frozen with rate controlled reactions

817
4. Transport properties: equal or unequal

diffusion or thermal diffusion coefficients

5. Surface boundary conditions: specify

component mass fluxes of char, pyro lysis, and edge

gas; or perform mass-energy balances

6. Edge conditions: specify total enthalpy,

pressure, and pressure distribution, or assigned edge

conditions including entropy layer

As used for the NOL tests, the BLIMP was set to solve

the nonsimilar, multicomponent , equilibrium, laminar

boundary layer through the axisymmetric nozzle and

duct. Unequal diffusion and thermal diffusion

coefficients were allowed. To account for the

difference in surface material between the non- ablating nozzle and Teflon, a discontinuous streamwise surface was selected with the nozzle wall

temperature being held fixed while the duct wall

temperature was solved for by an energy balance

across the ablating teflon surface.

The BLIMP solves the conservation equations and

boundary condition equations by an integral-matrix

procedure using a generalized Newton-Raphson technique to reduce the errors to zero. The conservation equations are expressed in terms of the primary dependent

variables, f, Hrp and K-^ and their derivatives with

respect to n . These equations are integrated across nodes within the boundary layer and across these

nodes the primary variables are represented by a cubic spline fit.

The CMA specifies the transient, in-depth,

thermochemical response of the ablator. Basically

the CMA solves the one-dimensional, unsteady, heat

conduction problem by a method of finite differences.

The basic assumptions are:

1. maximum of three components for the

ablator - two resins and one reinforcement

2. density history follows an Arrhenius type

law

char

pyrolysis gas is in equilibrium with the

4. pyrolysis out-gassing is immediate

5. no coking

6. cross-sectional area can vary

7. one-dimensional conduction

Within the program, the user is free to select one

of three options for boundary conditions:

(a) specified surface chemistry (i.e., species

concentration), (b) specified surface temperature and

recession rate and (c) specified radiation view

factor and flux. For the present experiments option

(b) was used with the Teflon composed of only one constituent, i.e., pg=p^=0.

818
The fundamental equations

1. The energy balance

are: ^ (phA) = -i- (kA -1^)^ + 7^ (m h ) ^T. 'y 9y 9y t 9y g g t

2. density history

-E/RT m

3. composition

p = r (P^+Pg) + (1-r) p^

Allowing the ablating surface to recede, the material

is considered as an array of arbitrary in-depth nodes

and solved by a finite difference technique.

RESULTS OF LAMINAR BOUNDARY-LAYER TESTS

A total of six ducts were tested at M = 3,

p ~ 21 atmospheres (21.2x10^ N/m2) , and T = 9000°R

(5000°K) . These laminar boundary-layer tests lasted

from 11 to 13 seconds. During this time the pressure

distribution within the ducts was observed to be

relatively steady with time and distance, thus

discounting the possible existence of internal shocks.

The data in Figure 4 show that there was a slight

favorable pressure gradient within the duct due to

over-compensation for boundary-layer growth.

The temperature measurements were partially

successful. The surface thermocouples did not recede

with the ablating surface, and as a consequence, they

became exposed to the flow. Upon exposure the

readings became erratic; the onset of this irregular

behavior, however, allowed us to fix the onset of

ablation to be at about four seconds from the start

of the test. The in-depth thermocouple plugs

performed well. There are some limitations imposed,

though. The spacing between the thermocouples,

0.025 inch (0.63 mm) was too wide to accurately

evaluate the wall heat transfer. Furthermore, there

was perhaps a 10 percent inaccuracy in the data as

a result of tolerance errors in the placement of the

wires. The data are in Figure 5. The apparent

discrepancy of trend between experimental data and

theoretical predictions is due to the method of

819
computation which assumes an instantaneously hot

wall rather -^nan a transiently heated wall. During

none of the tests was steady-state completely reached,

as predicted by the CMA "heat storage" term:

which never vanished and was still approximately six

or seven percent of the surface heat transfer.

After the tests the gas-wetted surface of the

ducts developed a glossy or waxy appearance; in some instances portions of the surface were slightly

brown

As shown in Figure 6 a pattern of parallel

striations formed in two of the ducts tested under

the predicted laminar conditions. The ducts were

supposedly tested under the same conditions as those

that did not exhibit the longitudinal grooving. It

is held possible that these grooves are the result

or manifestation of Gfirtler vortices resulting from a concavity in the duct (10) immediately behind the

nozzle junction, or represent the existence of

longitudinal vortices behind a rearward facing step

at the junction of the nozzle and duct (11) . This

type of parallel grooving has also been observed on ablating Teflon cones (12)

The final diameter increase of two of the ducts

is shown in Figures 7 and 8. In each case the

resulting profile is compared with that predicted by

the BLIMP-CMA procedures. The agreement between the

resulting contour and predicted contour is quite goodup to a point. In all of the supposedly laminar

boundary-layer tests, a pronounced increase in ablation

was observed to begin at about 3.5 inches (88.9 mm)

into the duct and reach a maximum at about 4.3

(109.2 mm) to 4.5 inches (114.3 mm). Supporting

data indicate this increase in ablation was due to

transition from laminar to turbulent flow. Predictions by the BLIMP program indicate a momentumthickness Reynolds number, ReQ, of 450 at this point;

this value is high enough to support transition. The

calculated boundary- layer profiles also support

transition. As shown in Figure 9 the profiles at

three locations within the duct exhibit, first, a normal behavior and then an increasing degree of

inflection as the point of ablation increase is

approached and then passed. This inflection in the

velocity profile is indicative of instability that

leads to transition. Transition may also be the

820
cause of the parallel striation pattern previously

noted; this effect is described in References 13

and 14.

Finally, for the laminar tests, the wall shear

response was recorded. In the first test the balance

was placed in the region subject to transition.

Figure 10 shows the balance response. The initial

transient declines to a minimum at about 5.8 seconds,

and then rises toward steady state at about 13

seconds. To further examine this response, two

balances were installed in one duct; one balance was located as before, and the other was placed upstream

in the laminar region. The balance in the transitional region repeated its behavior; the laminar zone balance responded as shown in Figure 11. The comparison with the BLIMP-CMA curve is not quantitatively

as good as hoped, although the qualitative behavior

is as predicted. The final value of measured wall

shear is about 36 percent of the predicted cold wall

shear.

TURBULENT BOUNDARY- LAYER TESTS

The ducts numbered IT through 5T in Table I were tested under conditions designed to produce a turbulent boundary layer. The general conditions were M = 2.3, Pq = 21 to 28 atmospheres (i.e., 21.2 - 28.3xl05 N/m2) and Tq = 4300 to 4600°R (2388 to

2555°K) . The specific purpose of these tests was to

examine the behavior of a cracked ablating surface

(15) . The results described here are those pertinent

only to the surface response of the Teflon.

The ablative performance of the ducts was uniform.

Mass loss increased from zero to a maximum loss about

one inch down the duct and then decreased to about

70 percent of the maximum for the remainder of the

duct. The abla,tion spike observed in the laminar

tests was not observed in any of the turbulent tests.

The mass loss was found to vary linearly with time

as shown in Figure 12, The ablation was very

pronounced and the pressure was observed to drop so that the Mach number increased from 2.3 to 2.6.

Because of this change in running conditions, testing

times for the turbulent tests were kept between 2.7

and 6.5 seconds. Wall shear was recorded in two

tests, at 21 (21.2x10^ N/m2) and 28 atmospheres

(28.3x10^ N/m2) . The balance reading indicated a rapid approach to steady-state, within two seconds,

reaching a value of 6 0 percent of the cold wall

shear.

821
Finally, the now famous cross hatching was observed in all of the turbulent tests. The pattern

started at the duct entrance and formed over the

length of the duct. Typical of the patterns formed

is that shown in Figure 13. The pressure taps were

seen to cause a change in the pattern from a crisscross to a more pronounced "V" form. The general

pattern is that of a depressed groove with a raised,

diamond-shaped, internal island. The pattern angle

for these M = 2.3 tests was 32 degrees (0.55 rad)

this is consistent with data reported by Laganelli

and Nestler (12) . There was no observed liquid layer

run-off during these tests. At present no attempt

at correlation among the various theories (16,17,18,

19,20,21) is being attempted.

SUMMARY

A successful series of ablation tests have been

carried out on Teflon ducts exposed to supersonic

laminar or turbulent boundary layers. Static

pressure, surface and in-depth temperature, mass

loss, and skin friction were recorded and compared

with a sophisticated, theoretical approach. There

was good agreement between the data and predictions.

The wall shear response was measured under laminar,

transitional, and turbulent ablating boundary layers.

In a series of turbulent tests, ablation rate was found to vary linearly with time, wall shear was reduced to 60 percent of the predicted cold wall

value, and striations were observed. For the same test Mach number, the pattern angle was constant

and compares well with previously measured values.

822

V
REFERENCES

(1) Winkler, E. M. , Madden, M. T., Humphrey, R. L.

and Koenig, J. A. , "Supersonic Ablation Studies

with Teflon," NOLTR 69-125, U. S. Naval Ordnance

Laboratory, White Oak, Md. , Oct. 1969

(2) Bartlett, E. P. and Kendall, R. M. , "Nonsimilar

Solution of the Multicomponent Laminar Boundary

Layer," Aerotherm Report No. 66-7, Pt III,

Aerotherm Corp., Palo Alto, Calif., Mar. 1967

(3) Moyer, C. B. and Rindal, R. A., "Finite

Difference Solution for the In-Depth Response

of Charring Materials Considering Surface

Chemical and Energy Balances," Aerotherm Report

No. 66-7, Pt II, Aerotherm Corp., Palo Alto,

Calif., Mar. 1967

(4) Schmidt, D. L. , "Ablation of Fluorocarbon Resins

and Components," Technical Report AFML-TR-68-393

Air Force Materials Laboratory, Oct. 1969

(5) Winkler, E. M. , Humphrey, R. L. and Koenig, J. A.,

"The NOL Four-Ring, Three Phase, AC Arc Heater,

Mark IV," NOLTR 70-12 (to be published), U. S. Naval Ordnance Laboratory, White Oak, Md.

(6) Solomon, J. M. , "Calculation of Laminar Boundary

Layers in Equilibrium Dissociated Air by an Extension of the Cohen and Reshotko Method,"

NOLTR 61-143, U. S. Naval Ordnance Laboratory,

White Oak, Md., Feb. 1962

(7) Glowacki, W. J., "Fortran IV Program for Calculating the Turbulent Boundary Layer Growth

for Contoured Axisymmetric Nozzles Using Air or Nitrogen," NOLTR to be published, U. S. Naval

Ordnance Laboratory, White Oak, Md.

(8) Aerodynamics Department Staff, "The NOL 3 Megawatt Arc Tunnel," NOLTR 6 6-8 0, U. S. Naval

Ordnance Laboratory, White Oak, Md. , Sept. 1966

(9) Bruno, J., Yanta, W. , and Risher , D. , "Balance

for Measuring Skin Friction in the Presence of

Heat Transfer," NOLTR 69-56, U. S. Naval

Ordnance Laboratory, White Oak, Md. , June 19 69

(10) Gortler, M. , "On the Three-Dimensional

Instability of Laminar Boundary Layers on Con- cave Walls," NACA TM 1375, June 1954, translation of original 1940 paper

(11) Ginoux, J. J., "The Existence of Three Dimensional Perturbations in the Reattachment of a Two Dimensional Boundary Layer after Separation,"

AGARD Report 272, Apr. 1960

823
(12) Laganelli, A. L. and Nestler, D. E., "Surface

Ablation Patterns: A Phenomenology Study,"

AIAA Journal, Vol 7, No. 7, July 1969, pp 1319-

1325

(13) Gregory, N. , Stuart, J. T. and Walker, W. S.,

"On the Stability of Three Dimensional Boundary

Layers with Applications to the Flow Due to a Rotating Disk," Boundary Layer Effects in

Aerodynamics , Philosophical Library Inc. , 1957

(14) Knapp, C. F. and Roache, P. J., "A Combined

Visual and Hot-Wire Anemometer Investigation of

Boundary Layer Transition," AIAA Journal,

Vol 6, No. 1, Jan. 1968, pp 29-36

(15) Winkler, E. M. , Humphrey, R. L. , Koenig, J. A.,

and Madden, M. T., "Investigation of Substructure Heating on Cracked Ablative Heat Shields,"

NOLTR 69-122, U. S. Naval Ordnance Laboratory,

White Oak, Md. , July 1969

(16) Larson, H. K. and Mateer, G. G., "Cross-Hatching

A Coupling of Gas Dynamics with the Ablation

Process," AIAA Paper 68-670, Los Angeles, Calif.,

1968

(17) Inger, G. R. , "Discontinuous Supersonic Flow

Past an Ablating Wavy Wall," Report DAC-62375,

McDonnell Douglas Corp. , Santa Monica, Calif.

1968

(18) Nachtsheim, P. R. , "Analysis of the Stability

of a Thin Liquid Film Adjacent to a High-Speed

Gas Stream," NASA TN D-4976, Jan. 1969

(19) Tobak, M. , "Hypothesis for the Origin of CrossHatching,." AIAA Paper 69-11, New York, N. Y.

1969

(20) Mirels, H. , "Origin of Striations on Ablative

Materials," AIAA Journal, Vol 7, No. 9, Sept.

1969, pp 1813-1815

(21) Probstein, R. F. and Gold, H, , "Cross-Hatching

a Material Response Phenomena," AIAA Journal,

Vol 8, No. 2, Feb. 1970, pp 364-366
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825
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826
SCALE: 1 inch (25.40 mm)

I 1

FIG. 1 NOZZLE AND ABLATION DUCT ARRANGEMENT

FIG. 2 THERMOCOUPLE PLUG FOR IN-DEPTH TEMPERATURE MEASUREMENT

827
828
200

(2.666)

100 (1.333)

DUCT #2

- 21.4ATM (21.67 X 10^ N/m^

= 4879°R (2710°K) STEADY STATE

2 (50.8) 4

(101.6)

X, DISTANCE DOWN DUCT

(inches) (mm)

6 (152.4)

DUCT #3

200

(2.666)

100 (1.333)

22.0 ATM (22.86 X 10^ N/m')

4375°R (2430°K) STEADY STATE

-O —O2 (50.8)

X, DISTANCE DOWN DUCT

4

(101.6) (inches) (mm)

6 (152.4)

FIG. 4 PRESSURE DISTRIBUTION IN TEFLON DUCT

829
I I J L \ I \ I 0 2 4 6 8 10 12 14 TESTING TIME (seconds)

FIG. 5 PREDICTED AND MEASURED IN-DEPTH TEMPERATURES

830

0. 12

j|

LUCO 0.08

iS (2.03) Oz

OS

LUs

<

0.04

y-o (1.01)

H 4098 BTU/LB (9.524 x 10" J/kg)

0

2 (50.8) 4

(101.6) AXIAL DISTANCE FROM DUCT ENTRANCE

(inches! (mm)

6 (152.4)

FIG. 7 PREDICTED AND MEASURED TOTAL RECESSION

0. 12 (3.04)

^ 0.08 S (2.03) z

q:

LU5 0.04

t3 (1-01)

= 21 ATM (21.27X lO^N/m^)

h; = 3643 BTU/LB (8. 466 x 10" J /kg)

0

^

t = 13 SEC

J L_

2 (50.8) AXIAL DISTANCE FROM DUCT ENTRANCE

4 6 (101.6) (152.4) (inches) (mm)

FIG. 8 PREDICTED AND MEASURED TOTAL RECESSION

832
0 2 4 6 8 10 12 TIME (seconds)

FIG. 10 WALL SHEAR RESPONSE

833
3 ^

(143.6) ^ E

< (95.7)

1 (47.8)

DUCT #6

X = 2.75 in. (69.85 mm)

BALANCE READING

PREDICTED RESPONSE

J \ \ \ \ \ \ \ I \ \ I L

0 2 4 6 TIME (seconds!

10 12

FIG. 11 WALL SHEAR RESPONSE

TIME (seconds)

FIG. 12 VARIATION OF TOTAL ABLATION OF FOUR DUCTS AS

A FUNCTION OF TESTING TIME

834


Paper No, 52

PREDICTION OF MECHANICAL EROSION OF CHARRING ABLATORS

Lauri H. Hillberg^

REFERENCE: Hillberg, Lauri H., "Prediction of Mechanical Erosion

of Charring Ablators, " ASTM/IES/AIAA Space Simulation Conference,

14-16 September 1970. ABSTRACT: A simple method has been developed for predicting the mechanical erosion of charring ablators exposed to an aerodynamic environment. The method requires the knowledge of two variables: the aerodynamic shear stress at the ablating surface and the surface temperature. Test data are used to obtain two empirical constants necessary

for correlating the erosion recession rate in terms of these two variables. The empirical constants have been determined for phenolic cork, phenolic carbon, epoxy-novolac, and high density silicone rubber. KEY WORDS: ablation, char, erosion, aerodynamic shear, phenolic

carbon, cork, epoxy-novolac, silicone rubber.

Nomenclature

2 2 2 A = erosion constant, Ibm/ft sec/lbf/in (kg/m sec/N/m

B = erosion constant, °R (OK)

M = Mach number

2 2 m = mass erosion rate, Ibm/ft sec (kg/m sec) P = pressure, Ibf/in^ (N/m^)

s = surface recession rate, in/sec (m/sec)

AS = thickness lost, inches (m)

t = time, sec

T = temperature, R( K)

Senior Engineer; Aerospace Group, The Boeing Company,

P.O. Box 3999, Seattle, Washington 98124

837
Oi = angle of attack, deg (rad)

p = density, Ibm/ft^ (kg/m^)

2 = aerodynamic shear stress, Ibf/in (N/m ^ = blocking function

Subscripts v

e = erosion

f = final o = at onset

s = at surface

t = total w = wall

Introduction

One of the strongest factors affecting the performance of a charring ablator is the recession of the char surface. This recession is detrimental to the performance of the ablator since additional material must be provided to compensate for the material loss. In order to deter- mine the amount of this loss and also to provide a means for evaluating

the overall ablator response, it is necessary to evaluate all mechanisms

which contribute to the surface recession. One such mechanism is mechanical erosion, a term which represents the material loss attributed

to factors which include aerodynamic shear, thermal stress, and internal gas pressure. Analytical prediction of mechanical erosion has been extremely

difficult. The most successful method has been to use sophisticated computer programs which incorporate an integrated thermal -structural analysis (1,2,3)^. The purpose of the structural analysis is to provide a means

for examining the stress state in the material and determining structural failure of the char layer. This failure, occurring repeatedly during

ablation, is then considered to constitute the mechanical erosion. The

disadvantage of this type of approach lies in the need to define numerous structural properties of both the virgin plastic and char over a wide

range of environmental conditions. These properties often require extensive testing to determine (e.g.. Reference 4) and may be erroneous

or may have been determined at conditions incompatible with the

The number in parenthesis refers to the list of references appended

to this paper.

838
intended environment. Consequently, the accuracy and usefulness of a complex structural analysis for predicting mechanical erosion may be

lost because of the lack of well defined structural properties.

It is apparent that there is a need for a mechanical erosion prediction method which has greater simplicity than the complex structural analysis and which can be presented in terms of easily defined variables. Two such methods have already been proposed (5,6). In both of these methods, the mechanical erosion rate is determined by specifying a maximum char thickness permissible under certain conditions. When

these conditions are reached the char fails at the point determined by

that maximum thickness. Repetition of this failure process constitutes the mechanical erosion. The maximum char thickness used in these two

methods is presented in terms of aerodynamic shear (5) or pressure (6) and

a few empirical constants. A third simplified method for predicting mechanical erosion is presented in this paper. In this method the only external force influencing

the mechanical erosion rate is assumed to be the aerodynamic shear. Thus, this method is not applicable to stagnation regions or regions having insignificant aerodynamic shear levels. Use of Reference 6 or the

simplified stress analysis in Reference 5 may be utilized in these areas.

Theory

Several assumptions are necessary for developing this method.

First, as was mentioned above, only aerodynamic shear loads are assumed

to act on the char. Pressure effects are neglected. However, the necessary empirical constants are determined under conditions having both aerodynamic shear stress and normal stresses induced by the pressures exerted by the gaseous ablation products as they flow through the char

layer, !n this respect, pressure effects are included.

The second assumption is that mechanical erosion is a continuous

process and not a series of discrete failures of the char layer. If discrete

failures do occur, then it is assumed that the frequency of the failures and the thicknesses of the material lost are such that the recession can

be handled as a continuous process. The third assumption is that the aerodynamic shear acting on the ablating surface is equal to the shear in the absence of ablation times

the blocking function, ^ , which accounts for the effects of mass

injection.

The final assumption is that all thermo -chemical processes such as oxidation, sublimation, and decomposition can be readily analyzed.

This assumption is in keeping with the capabilities of ablation computer

programs now available. The mechanical erosion rate can then be

839
determined from tests by measuring the total recession rate and subtracting the thermo-chemical contribution.

Utilizing the above assumptions, an analytical prediction method

based on the model developed by Gaudette (7) is proposed. Gaudette's

model, which analyzed the mechanical erosion of silica-filled epoxynovolac, assumed that the erosion rate was a function of the surface temperature, this temperature reflecting the resistance of the viscous melt layer formed on the surface of the char layer. His equation for the mechanical erosion rate has the form

e where K is a constant set at 0.5 for epoxy -novo lac. The original erosion data obtained in Reference 7 along with the curve representing the selected functional relationship are shown in Figure 1

The method proposed here assumes that Gaudette's model may be

extended to apply to all charring ablators and not just those which form

a viscous melt layer. This assumption implies that the surface temperature

also reflects the structural failure resistance of the char layer,, Reference

8 briefly discusses this assumption.

It is possible upon examining the erosion data of Reference 7 (and

the data discussed subsequently) to simplify Gaudette's modeL The

erosion data shown in Figure 1 has been replotted on linear scales in Figure 2, A straight line drawn through these data intercepts the origin

and yields an equation for the erosion rate of the form

e

or, after converting to the base e , m - A^l^re'^/'^W

e Dividing through by the shear stress at the ablating wail, , yields

the erosion model proposed in this paper.

-B/Tvv

This model shows the mechanical erosion rate divided by the aero- dynamic shear stress at the ablating wall is dependent only on the surface temperature. It is thus possible to easily predict the mechanical erosion

rate for a charring ablator through the knowledge of three parameters and

two empirical constants which can be readily determined from tests.

840
ResultsThe two empirical constants required for the proposed mechanical

erosion method can be directly determined from wind tunnel or free

flight tests. Determining them involves plotting the variable m^/^r

versus l/T^^ on semi -log graph paper. The slop is constant B and the intercept at 1/T^ equal to zero is constant A. These empirical constants have been determined for four distinctly different ablators: epoxynovolac, high density silicone rubber, phenolic cork, and phenolic

carbon

E poxy -Novolac The equation for the straight line drawn through the data shown in Figure 2 is used directly to obtain the necessary empirical constants for Avcoat 5026-39H/CG epoxy -novo lac. The equation for the mechanical

erosion, after converting the exponential base from 0.5 to e , is A. = 0 52 e^93°/^W

High Density Silicone Rubber

The empirical constants have been determined for Dow Corning

DC 93-072 silicone rubber { p = 70 lbs/ft^). These constants were determined from tests conducted in the Boeing Supersonic Combusion Wind

Tunnel (BSCWT). The nominal wind tunnel conditions were: M = 2.2

P = 37 to 112 psi T^ = 1130to 2020OF

The model was a sharp flat plate with boundary layer trips located 0,2" from the leading edge. A 2"x 4" x 0.2" ablator specimen was bonded to a plate which was then bolted into a recession located 0.7

inches behind the boundary layer trips. The ablator surface initially is flush with the leading edge. The flat plate was set at either 0^, 10^,

or 15^ angle of attack during the tests. The mechanical erosion rates required to obtain the erosion constants were reduced using the following equation

841
The surface densiiy of fhe ablator required in the above equation was determined using a transient ablation analysis computer program (9) which utilizes an Arrhenius decomposition rate law. The thickness lost was determined directly from before and after measurements of lOX

photographs. It had previously been found from TGA tests conducted on

the char in air that chemical reactions did not contribute to surface

recession. Finally, the onset time, or the time when surface erosion began in the tests, and the time terminating the test were determined from

slow motion movies.

The three unknowns remaining in the correlation equation, the blocking function, the aerodynamic shear stress, and the surface temperature, were determined analytically using known material properties

and measured wind tunnel conditions. Several iterations using various erosion constants were required to obtain the correct values for these three parameters.

The erosion data from these series of tests are summarized in Table

1. The resulting data are plotted in Figure 3. The straight ^ine drawn

through these data is represented by

Phenolic Cork

Erosion constants were obtained for Armstrong AC-2755 phenolic

cork using a combination of wind tunnel tests and flight test data. Wind

tunnel test data obtained from two tests conducted in the Boeing Hypersonic Wind Tunnel (BHWT) were reduced in the same manner as the DC 93-072 silicone rubber with the exception that surface recession attributed to chemical oxidation was included. Constants used for the reaction limited oxidation were obtained from TGA tests conducted on

the cork char in air; constants for the diffusion limited oxidation were

obtained assuming

C + ~ 02-^C0

The wind tunnel test model was a 15^ half-angle cone having a

sharp steel nose tip. The wind tunnel conditions were M = 6.05

?f = 850 and 1000 psi T = 1000°F

842
The second of the two tests used a quartz lamp to supplement the aerodynamic heating rates. The wind tunnel test data are summarized

in Table 2 and plotted in Figure 4. Missile flight test data were also used to supplement the wind

tunnel data. Flight ablation meters yielded thickness ablated as a function of time. The local slope of this curve yielded the ablation rate, also as a function of time. Using the known flight trajectories and a

transient ablation analysis computer program (9), the surface density, surface temperature, blocking function, and aerodynamic shear stress were determined. Utilizing the measured ablation rates and the supplemental analytical data, the necessary erosion data were determined.

These data are also plotted in Figure 4. The erosion equation resulting

from these data is

Phenolic Carbon

Erosion constants were obtained for Avco X6300 phenolic carbon

having a 90° fabric layup using the data presented in Reference 10. The

mechanical erosion rates were determined by taking the recession rates indicated in Reference 10, converting them to mass loss rates by assuming

a char density of 70 Ibm/ft^, and subtracting the recession attributed to oxidation. The oxidation recession was assumed to be that for diffusion limited combustion for the reaction

C . l02-^C0

The aerodynamic shear stresses and the surface temperatures shown

in Reference 10 were used, except that a Reynolds analogy factor of

1 .2 was applied to the shear stress. The final variable, the blocking

function, was determined analytically using a transient ablation analysis

computer program (9). The results are tabulated in Table 3 and are

plotted in Figure 5. From this figure it was found that the erosion rate can be represented by

= 7.5 X ,0^-77000Aw

843
Discussion

The results obtained for the proposed erosion rate equation indicate

that mechanical erosion can be correlated, at least for the materials and

conditions examined. It is interesting to note that such a simple equation

does correlate as well as it does for the different ablator types and the

different ranges of test environments.

It is realized that the results may be limited. Except for the

missile flight tests for the phenolic cork, the tests were essentially steady

state tests. Consequently, the temperature profiles and char depths may

not match flight conditions. The data shown for phenolic cork in Figure

4 do, however, indicate a correlation between wind tunnel and flight

test data. However, the cork char is very fragile and was quite thin in both the wind tunnel and flight tests. Consequently, it is very difficult to make any significant conclusions from these data. To test the ability of the erosion model to aid in correlating temperature data, several instrumented DC 93-072 silicone rubber ablation

models were run in the same test series that were used to obtain the empirical constants. The results of one such test are shown in Figure 6. The conditions of the test were M =2.2

= 70 psi T^ = 1850OF a = 15^

The temperatures were predicted for a location 2.0" aft of the leading edge. The correlation between the predicted and measured

values can be seen to be excellent. Correlation without the mechanical

erosion model was extremely poor, largely due to the fact that no recession was predicted.

Conclusions

A simple, straight forward method has been developed for the prediction of mechanical erosion of charring ablators in a shear environment.

By virtue of the ability to reduce test data in terms of the proposed

equation, the method appears valid, at least for the materials and test conditions which were examined.

Empirical constants for use in the proposed erosion equation have

been determined for e poxy -novo lac, high density silicone rubber,

phenolic cork, and phenolic carbon.

844
References

1. Schneider, P. J., Dolton, T.A., and Reed, G.W.; "Mechanical

Erosion of Charring Ablators in Ground-Test and Re -Entry Environments, " AlAAJ^, Vol. 6, No. 1, Jan. 1968, pp 64-72.

2. Hillberg, L. H.; "Influence of Material Properties on Re-Entry

Vehicle Hect Shield Design, " AIAA/ASME 8th Structures, Structural Dynamics and Materials Conference, March 1967.

3. Hoglund, R. F. , Babcock, S.E., and Grovens, R.N., Jr.; "Ablation Test Specimens Environment at High Temperature, " AIAA/

ASME 10th Structures, Structural Dynamics and Materials Conference, April 1969.

4. Ching, A., and Welsh, W.E., Jr.; "Strength and Stress-Strain Properties of Rapidly Heated Laminated Ablative Materials, " AIAA J ., Vol. 6, No. 12, Dec. 1968, pp 2312-2315.

5. Mathieu, R. D, ; "Mechanical Spallation of Charring Ablators In Hyperthermal Environments, " AIAA J ., Vol. 2, No. 9, Sept. 1964, pp 1621-1627.

6. Bishop, W. M. and DiCristina, V.; "A Prediction Technique for Ablative Material Performance Under High -Shear Re-Entry Conditions, " AIAAJ_. , Vol. 6, No. 1, Jan. 1968, pp 59-63.

7. Gaudette, RoS., del Casal, E.P., and Crowder, P. A.; "Charring

Ablation Performance in Turbulent Flow, Volume I," Document

D2-1 14031-1, The Boeing Company, June 1967.

8. Rindal, R.A.; "An Evaluation of Design Analysis Techniques for High Performance Ballistic Vehicle Graphite Nose Tips, Appendix

D: Thermo-chemical and Thermomechanical Ablation," Air Force Materials Laboratory, AFML TR-69-73, Volume IV, January 1970.

9. Hillberg, L. H.; "The Improved Convective Heating and Ablation

Program (CHAP)," Document D2- 26104-2, The Boeing Company,

June 1969.

10. Mitchel, B. J.; "Ablative Material Tests Under Transient Heating

Simulating Ballistic Re-Entry," AIAA Paper 69-150, January 1969.

845
TABLE 1—Erosion Data for DC 93-072 Silicone Rubber

AS

4- * f o

Run (psi) (m) (sec) (*) (°R)

1 .055 .015 26.8 4. Ox 10"^ 1560

2 .102 .035 23.2 5.9x 10"^ 1620

3 .117 .037 14.2 9. Ox 10~^ 1680

4 .114 .076 41.7 6.4x 10"^ 1660

5 .108 .024 24.4 3.6x lO'^ 1520

6 .105 .036 54.8 2.4x 10"^ 1480

7 .056 .016 22.8 4.9x 10"^ 1610

8 .080 .022 27.4 3.9x 10"^ 1530

9 .121 .043 10.5 14.0x 10"^ 1760

10 .083 .044 19.8 11. Ox 10"^ 1750

n .122 .135 21.7 19. Ox lO"^ 1810

12 .061 .025 19.1 8.5x 10"^ 1690

13 .055 .033 27.7 8.7x 10"^ 1660

14 .050 .006 37.2 1.2x 10"^ 1410

15 .098 .013 42.2 1.2x 10"^ 1390

* (lbm/ft^sec)/(lbf/in^)

846
TABLE 2 —Erosion Data For Armstrong AC -2755

Phenolic Cork

Run (psi)

t (sec) (*)

1 .069 30.0 5.1x10"^ 1230

2+ .077 30.0 1.8x10"^ 1310

* (lbm/ft^sec)/(lbf/in^ + aerodynamic heating supplemented by quartz heating lamps

847
TABLE 3—Erosion Data For Avco X6300 Phenolic

Carbon (90° Fabric Layup)

Specimen

Code

s (in/sec) (Ibm/Tt sec) ®

2 (Ibm/ft sec) X6300-4 .026 .152 .134

X6300-5 .051 .297 .265

X6300-3 .024 .140 .104

Specimen T + rrie/^T

Code (psi) (**) (^R) X6300-4 .62 .20 1.10 5780

X6300-5 .63 o35 1.20 5680

X6300-<3 .73 .29 0.49 5150

* based on a surface density of 70 Ibm/ft + values shown in Reference 10 corrected for a Reynolds

analogy factor of 1 ,

** (lbm/ft^sec)/(lbf/in^

848
Fig. 1 —Mechanical Erosion Data for Avcoat 5026-39H/CG.

Fig, 2 —Mechanical Erosion Data for Avcoat 5026-39H/CG.

849
\

0.5 oTi 5.I

Fig. 3 —Mechanical Erosion Data for Dow Corning DC 93-072.

AHMSTRONG AC-2755 PKNOUC COUK

Fig. 4 —Mechanical Erosion Data for Armstrong AC-2755 Phenolic Cork

850
1.2

AVCO X6300 PhENOUC CARBON

(90° FABRIC lAYUP) • lEDUCED FROM DATA IN REFERENCE 10

Fig, 5 —Mechanical Erosion Data for Avco X6300 Phenolic Carbon

Fig. 6 —Temperature Test Data for Dow Corning DC 93 -072.

851

Paper No. 53

HEATING IN CRACKS ON ABLATIVE HEAT SHIELDS

1 9 o E. M. Winkler, R. L. Humphrey, M. T. Madden,^

and J. A. Koenig^

REFERENCE: Winkler, E. M. , Humphrey, R. L.

Madden, M. T., and Koenig, J. A., "Heating in Cracks

on Ablative Heat Shields," ASTM/IES/AIAA Space

Simulation Conference, 14-16 September 1970

ABSTRACT: An experimental program has been carried

out in the U. S. Naval Ordnance Laboratory (NOL)

3 Megawatt Arc Tunnel to study the effect of cracks in an ablative heat shield on the substructure heating.

The tests used a supersonic contoured nozzle with a Teflon duct attached to it. The ducts had transverse

and longitudinal cracks machined into the surface.

They were instrumented for pressure, temperature,

heat transfer and skin-friction measurements. The

cracks were found to have pronounced effects on the

ablative bheavior. The heating is moderate under a laminar boundary layer, but can be catastrophic when

the boundary layer is turbulent, depending upon the

size and direction of the crack. The results for the

transverse cracks were compared with an available

analytical prediction. The heat-transfer measurements

tend to support the concept of vortex cells existing

within the ci'acks. Cracks that are deep as compared

with their width result in very little substructure

heating.

KEY WORDS: ablation, substructure heating,

catastrophic ablation, surface cracks

Chief, High-Temperature Aerodynamics Group,

Aerophysics Division, Naval Ordnance Laboratory (NOL)

^Aerospace Engineers, NOL

^Aerospace Engineering Technician, NOL

853
NOMENCLATURE

b = crack width

H = enthalpy

h = crack depth

M = Mach number m = material loss per unit time

q = heat-transfer rate

Re = Reynolds number

St = Stanton number

u = velocity

X = distance from duct entrance

y = depth of crack

6* = boundary- layer displacement thickness

p = density

SUBSCRIPTS

c = bottom of crack

s = surface of duct w = wall

9 = momentum boundary-layer thickness

T = shear stress

00 = free-stream values

INTRODUCTION

Ablative materials are an accepted means for

protecting re-entry vehicles from the intense

heating encountered during certain phases of the

flight. There is concern, however, that when cracks

develop in the ablative heat shield, the substructure

may be subjected to excessive heating. The ablative

854
behavior of the heat shield may be altered in the

vicinity of cracks which in turn can affect the aero- dynamics of the vehicle. The analytical description

of the processes is inadequate to allow a safe design

of heat shields. A moderate number of studies have

been performed on non-ablating, shallow cavities,

rectangular notches of various aspect ratios, and on the upstream and downstream sides of protrusions as,

for example, reported in References 1 through 4. Most

of the studies were done for low-speed flow. Early

experimental studies were done by Wieghardt and

Tillmann, References 1 and 2. They illustrated the

vortex flow pattern in the cavities and obtained

empirical relations for the effect of the cavities, or notches on the total drag of the vehicle. Analytical

studies (3,4)"^ established a dependency of the number

of vortices within given notches on the flow Reynolds

number. Heat- transfer data are reported in References

5 and 6, for example. These studies serve as a guide,

but since they do not cover a sufficiently wide range

of conditions, they did not lead to a satisfactory

theoretical treatment that could be applied to ablating

surfaces in high-speed flow. The present research

program was formulated to look into this matter.

APPROACH

NOL ' s approach, primarily experimental, endeavors

to produce actual flight conditions in the 3 Megawatt

Arc Tunnel and to measure parameters and material

responses that can be used to further the understanding of the processes involved. Re-entry bodies

except for manned space vehicles are typically

conical with a small to moderate amount of nose

bluntness. The heat shield may be Teflon, silica or carbon phenolic, or a number of other materials.

Representative flow parameters over the conical

part of a re-entry heat shield are: Mach number = 2.0

to 10, wall enthalpy (H^) /stream enthalpy (Hq) = 0.2

to 0.4, friction Reynolds numbers (Re^) = 70 to 2000.

Cracks in the heat shield may have width over boundarylayer thickness (b/6*) ratios of 0.05 to 10 while crack

heights to width ratios (h/b) could range from 0.1 to 10. Such conditions can be duplicated in an arc heated tunnel by using the similarity of flow down a nozzle from stagnation conditions to that around the

forward region of a re-entry body. If one makes the

reservoir conditions of the arc tunnel equal to the

The numbers in parentheses refer to the list of

references appended to this paper.

855
stagnation point values on the re-entry body, the

correct flight conditions on the initial portion of

the conical surface can be duplicated. This allows

the options of studying the flow external or internal

to a model with equal validity of the results. The

latter approach was selected because it is experimentally quite convenient. Even small models can be

easily instrumented for obtaining pressure, temperatur

heat- transfer and skin-friction data. In this case,

the model can be simply an axially symmetric duct as described in the next section. Instrumentation

supports and leads can be arranged externally, thus

eliminating the difficulty of providing aerodynamic

and thermal shields for them.

For the study of substructure heating due to

cracks, parameters other than Mach number, pressure

and enthalpy ratio need to be considered. From flow

studies in cavities, or cracks, with non-ablating

walls (1,2,3,4,5,6) one knows that the friction

Reynolds number, Re^ , characteristic sizes such as b/6* and h/b, as well as the geometry may influence

the flow patterns in the cavity and the heat transfer

to its walls. Geometry factors include orientation

of the cavity with respect to the flow direction,

spacing of the cavities, if there are more than one,

and the cavity edge geometry. With ablation present,

the blowing rate parameter, Cg = m/p^u^St also has

to be considered. The critical, actual size, b, of a crack is somewhat speculative until studies on this

subject matter are concluded. But it appears plausibl

that cracks of 0.25 mm (0.01 inch) width or larger

on the conical part of an ICBM heat shield may lead

to problems for the substructure depending on the

time of exposure to intense heating. The blowing

rate parameter, essentially an indication of the

amount of shielding provided by the ablator, will

vary with the material. The aspect ratio h/b will

depend on the thermal shock resistance of the heatshield material and the cause of the cracks. It is,

however, conceivable that it may range from 0.1 to

10.

EXPERIMENTAL ARRANGEMENT

The experimental set-up is shown in Figure 1.

The arc heater, part of NOL's 3 Megawatt Arc Tunnel

(7) , is a three-phase, four-ring, ac arc which

operates on dry air supplied from a high-pressure

storage field. The heater is followed by an axially

symmetric contoured nozzle designed by the method

of Reference 8. Two nozzles have been used during

856
the present investigations, one was designed for a Mach number 3, the other for a Mach number of 2.3.

From the nozzle, the flow enters an axially symmetric

duct made of the ablative material. The Teflon ducts

(TFE-7, E.I. Dupont Co.) with nominal inside diameter

of 25.4 mm (1 inch) and length of 152 mm (6 inches),

are tightly connected to the nozzle exit which has

the same diameter. The nozzle as well as the duct

contours have been corrected for the boundary-layer

growth. The Mach number 3 nozzle and ducts were designed for laminar flow, while the Mach number 2.3

nozzle and ducts had turbulent boundary- layer

corrections applied to them. Exiting from the ducts,

the flow then discharges into a test cell, diffuser

and exhaust system.

The Teflon ducts are instrumented with copperconstantan thermocouples, and static and Pitot

pressure gauges. The transverse and longitudinal

cracks machined into the ducts are equipped with heat

gauges on the end walls and bottom to measure local

heat transfer rates (Fig. 2) . In some ducts the

local skin friction was measured with a specially

developed balance (9) . Crack sizes, orientation, and instrumentation of

the ducts are summarized in Tables I and II. This

covers a representative but limited range of conditions of interest for the substructure heating

problem.

ANALYTICAL PROCEDURES

Several analytical programs have been used in

the analysis of the experimental data. Test results

obtained under conditions where the duct boundary

layer was predicted to be laminar were compared with

the BLIMP-CMA programs (Boundary Layer Integral

Matrix Procedure-Charring Material Ablator) (10)

In the present test series only one run was laminar.

The BLIMP-CMA procedures are quite versatile and

permit the calculation of Teflon ablation rates,

boundary-layer parameters, and in-depth material

response in non-similar, multi-component, chemicallyreacting, equilibrium, laminar boundary layers along

axisymmetric or planar surfaces. Quasi-steady or transient solutions can be obtained.

Unfortunately, a turbulent boundary-layer option

was not available for the BLIMP-CMA procedures. As

a consequence, the data obtained with turbulent

boundary layer along the duct wall, could not be

compared with predictions. However, a numerical

857
procedure was devised to analyze the experimental

data to the extent considered essential for the substructure heating studies. It combines a computation

applicable to turbulent boundary layers (8) on non- ablating walls, experimental information, and wallspecies and temperature data obtained from the use of

BLIMP. These are quasi-steady-state computations.

It was felt that this is acceptable since the

experimental shear-stress data indicate that for the

turbulent boundary-layer test runs, steady-state

conditions are reached in less than one second.

RESULTS : Laminar Boundary-Layer Test

There was only one test carried out under

conditions where the boundary layer was predicted to

be laminar through the duct. This was done at a Mach number of 3, a supply air pressure of 2.13 MN/m^

(21 atm) and stagnation temperature of about 5000°K

(9000°R) . The test duration was 13 seconds. The

duct had two cracks, one in the flow direction, the

other transverse to it. The size, aspect ratio, and

location of the cracks are listed in Table I. Both

cracks were instrumented with a heat gauge.

There was a small favorable pressure gradient

along the duct due to an overcompensation of the

boundary-layer growth along the duct. The surface

texture of all ducts became glossy and waxy (11)

this texture is presumably attributable to one of the

Teflon degradation species (12) . Ablation increased

from zero at the duct entrance - because of the

contact with the cooled nozzle - to a maximum at x =

30 to 50 mm (1,2 to 2 inches) from the duct entrance,

then decreased as expected for a laminar boundary

layer. At x = 90 mm (3,5 inches), the ablation

increases, reaches a maximum and then decreases again.

This anomaly is observed in all "laminar runs," and was analyzed as evidence of boundary-layer transition.

This duct also showed the longitudinal, almost parallel

striations which are further evidence of laminar

boundary-layer instability and transition, as discussed

in References 13 and 14. At the downstream side of

the longitudinal crack a pronounced wake-like additional ablation pattern is observed as seen from

Figure 3. The transverse crack shows only a slight

rounding on the downstream side.

The heat gauge in the longitudinal crack began

to respond as soon as hot air from the arc heater

reached it, about 10 ms. The temperature rise was

858
linear with time but at a moderate rate. There was

a 13.6°C (24.5°F) temperature rise in the 13 seconds

of the test. For the transverse crack there was no response discernible for the first second of the test.

Thereafter the temperature rose very slowly at a varying rate. The temperature variations correspond

to a heating rate of 2.7x10^ W/m^ (2.4xl0~l BTU/ft2sec)

for the longitudinal crack and to a maximum of

7.9xlo2 w/m2 (0.7x10"^ BTU/ft^sec) for the transverse

crack. In terms of q/qg the values are .0042 and

.0016, respectively.

Turbulent Boundary-Layer Tests

Six tests were performed at a Mach number of 2.3.

The boundary layer was turbulent throughout the duct.

The test conditions were 2.13 and 2.84 MN/m2 (21 and

28 atm) and temperatures of 2390 and 2560 degrees

Kelvin (4300 and 4600°R) . Ablation was very pronounced under these conditions, requiring that the

testing times be kept to six seconds or less. Five

of the six ducts had several cracks, longitudinal and

transverse to the flow, as listed in Table I. The

instrumentation varied, providing for pressure readings,

in-depth temperature recordings, and skin-friction

measurements along the ducts, and heat-transfer

measurements at the bottom and side walls of the

cracks. (See Table I.) The listed values of h/b and

b/6* refer to the cold wall conditions. The sixth

duct was used for a reference test for duct number

4T. This became necessary since duct number 4T was

run at conditions for which no reference data existed

from previous tests. Pressure, wall temperature and

skin-friction data were needed for the evaluation of

the results. The amount of instrumentation required

was in excess of what could be accommodated in one

duct.

The overall ablation behavior of the ducts in

this test series was very uniform. The mass loss

increased from zero at the entrance to a maximum at x - 25 mm (1 inch) and then leveled off to an almost

constant value of about 70 percent of the maximum.

Figure 4. The ablation rates varied from 0.776 to

0.561 kg/m^-sec (.159 to .115 Ib/ft^sec) along duct

number 5T. Expressed in terms of the blowing rate

parameter, m/p^u^St, the turbulent ablation was 0.99.

The experiments showed that the total amount of

ablation for a given value of wall heat transfer

varied linearly with time, allowing the onset of

ablation to be deduced. It was determined to be

0.3 second. All of the turbulent ducts exhibit

859
cross-hatched striation patterns as shown in Figure

5. The patterns start more or less iinmediately at

the duct entrance and form in a more uniform fashion

if a pressure orifice is not located there. A

pressure orifice seems to enhance the wave pattern

and promote its formation. The cracks have a pronounced effect on the striation pattern, they

alter the spacing or even cause them to disappear,

Figure 6. At the downstream side of longitudinal

cracks, the interference between the crack "wake" and

the striations produces a complex pattern, as seen in

Figure 7. Analytical studies are under way attempting

to explain the observed patterns.

The heat gauges at the bottom, and sides of the

cracks responded within 30 ms after arc ignition.

Some of the temperature readings are shown in Figure 9. For the longitudinal cracks the center of the downstream wall experiences practically the cold-wall

surface heating rate, at least initially. The

temperature rose to 950°K (1250°F) within the first

half second, then the rapid ablation caused the

destruction of the heat gauge and the recording

ceased. At the bottom of the crack. Figure 9, the

heating rate varies, first almost linearly with time,

from zero to 0.20 of the averaged surface heating

during the first 1.23 seconds of the testing time and

then reaches 100 percent of the surface heating within

the next 3/4 second. Burn-out of the heat gauge

occurred at this time. A faint pattern of transverse

vortices is discernible at the bottom of the cracks.

The center of the upstream wall showed an erratic

behavior probably indicative of a changing vortex

formation. Gouging, both laterally and longitudinally

occurs at the downstream end of the crack. It is

particularly severe when a longitudinal and a transverse crack interact.

The transverse crack heating rates read from

continuous oscillograph recordings showed a more moderate behavior, reaching at best eight percent

of the surface heating rate.

Unfortunately, the number of runs and the range

of variation of parameters was insufficient to draw

firm conclusions. The following are the observations:

a. For aspect ratios of four, the heat transfer

to the bottom of the transverse crack at t = 1 second

is about four percent of the duct surface heat

transfer. At the end of a four second test it may

be as high as eight percent. At this time, the

nominal aspect ratio has changed to h/b - 2.5.

However, the geometry has also changed, i.e., the

edges are rounded, and there is evidence of gouging

860
due to the vortices that had established in the

cracks

b. The value of h/b does not seem to be a unique correlating parameter nor does b/6* or Re^

However, the absolute width, b, seems to be of some significance, both in the substructure heating rate

and the deformation of the initial crack geometry.

This is evidence that a transverse crack will cease,

with increasing width, to be an "open cavity" and the

flow pattern in it will change to that of a "closed

cavity"

c. With increasing depth of the crack, the

heat transfer decreases very sharply, and is only

1.8 percent of the surface heating under a turbulent

boundary layer for an aspect ratio of 7. This is

consistent with the studies reported in Reference 15.

Detailed analysis and flow studies in transverse

cracks have led to relations for the variation of

the crack heat transfer q^/qg with the depth parameter

y/b, where y is the distance from the surface and

b the width of the crack. The analysis indicates

that the ratio of q^/qg is not a function of Mach

number and is not affected by ablation if the data

is given in terms of the changing crack depth as ablation progresses. The discontinuity in the curve is due to the transition between the solutions

chosen for the regions of y/b. The predicted

variation, for the transverse cracks, is shown in Figure 10 (solid and dashed lines) together with the

present results and other experimental data. It can

be seen that the crack heat transfer depends only

weakly upon the value of b/6 and moderately on the

friction Reynolds numbers.

d. The temperature records show some very

interesting trends. The short duration run, duct

number IT, shows a uniform rise in temperature. For

all other runs a periodic behavior is exhibited, or at least suggested by the temperature traces. After

a uniform rise of the crack-bottom temperature for

a certain time there follows a period where the

temperature remains constant for some time, then

the temperature rises again at about the initial

rate, there may even be a decrease of temperature.

This may repeat itself several times during one recording, as seen from the temperature traces of

Figure 11. The same periods of temperature rise and

constant temperature are not necessarily maintained

in two cracks in the same duct. This observation

strongly suggests that vortex flow exists in the

cracks, with the number of the counter-rotating

vortices being some multiple of the crack width.

861
Then, as ablation progresses, the crack height is

reduced and its width increased to a degree that the

vortex at the bottom of the crack can no longer be

maintained. The lowest vortex breaks up, and the gas

becomes more or less stagnant. The bottom temperature

then remains constant, or even decreases if the

vortex core temperature was lower than that of the

outer region, as seen from Figure 11. When the next

vortex has reached the bottom, or a reformation of

vortex patterns has taken place to fit the changed

height and width of the crack, the temperature

rises again. If one adopts the results of Reference

5, namely that 30 percent of the heat transferred

across the dividing streamline atop of a cavity is

transferred at the bottom of the first vortex, and

so on, the present results suggest that at the end

of the tests two or three vortices existed in the

transverse cracks.

SUMMARY AND CONCLUSIONS

Supersonic substructure heating studies were

carried out in the NOL 3 Megawatt Arc Tunnel. The

purpose was to assist the analysis of the effects of

cracks in ablative heat shields on the performance

of the vehicle. Teflon TFE-7 was selected as the

ablator.

The test conditions were chosen so that the

boundary layer in the Teflon duct was predicted to

be either all laminar or all turbulent. Transverse

and longitudinal cracks in the ducts were instrumented

to obtain heating rates at the bottom and side walls

of the cracks. In addition, wall-static pressures,

in-depth temperatures, and skin-friction data were obtained for the duct itself. Total ablation was evaluated from the wall recession.

The repeatability of test conditions and results was very good.

The non-dimensional heating rate for transverse

cracks was found in good agreement with other recent

experimental data and the prediction of an unpublished

analytical procedure. The temperature records obtained

from gauges on the bottom and side walls of the

cracks indicate a changing vortex flow in the cracks.

The heating rate and ablation rate at the bottom

and downstream walls of one-inch long longitudinal

cracks of aspect ratios of 4 and 8 will lead to catastrophic failure of the substructure. There are indications of transverse vortices on the bottom of

the longitudinal cracks.

862
Parallel striations, in the flow direction were observed ii the laminar duct and cross-hatched

striations in all turbulent ducts. Pressure orifices

and cracks have a pronounced effect on the striation

pattern and the ablative behavior of the material.

A procedure that combines measured pressure and

wall shear data with theoretical chemistry information

and a NOL-developed boundary-layer prediction procedure is judged successful in evaluating the surface

heating rate for a Teflon surface under a turbulent

boundary layer. More complete details may be found

in Reference 19. '

REFERENCES^

(1) Wieghardt, K., "Increase of the Turbulent

Frictional Resistance Caused by Surface

Irregularities," translated from Jahrbuch der

deutsden Liftfahrt forsdung, Vorabdruck lA 022,

1943

(2) Tillman, W. , "Additional Measurements of the

Drag of Surface Irregularities in Turbulent

Boundary Layers," NACA TM 1299 (translation)

Jan. 1951

(3) Pan, Frank and Acrivos, Andreas, "Steady Flows

in Rectangular Cavities," Journal of Fluid

Mechanics, 28, Pt A, 643, 1967

(4) Fox, Jay, "Heat Transfer and Air Flow in a Transverse Rectangular Notch," International

Journal of Heat and Mass Transfer 8 269, 1965

(5) Haugen, R. L. and Dhanak, A. M. , "Heat Transfer

in Turbulent Boundary-Layer Separation over a Surface Cavity," Journal of Heat Transfer,

Trans. ASME, Vol 89, Series 6, No. 4, 1967

(6) Nicoll, K. M. and Bogdonoff, S. M. , "Experimental

Studies of a Specific Cavity Configuration in

Laminar Hypersonic Flow," Archiwum Mechaniki

Stosowanej 1 16, 1964

(7) Aerodynamics Department Staff, "The NOL 3 Megawatt Arc Tunnel," NOLTR 6 6-8 0, U. S. Naval

Ordnance Laboratory, White Oak, Md., Sept. 1966

863
(8) Glowacki, W. J., "Fortran IV Program for

Calculating the Turbulent Boundary Layer Growth

for Contoured Axisyininetric Nozzles Using Air

or Nitrogen," (to be published as a NOLTR)

(9) Bruno, J. R. and Risher, D. B. , "Balance for

Measuring Skin Friction in the Presence of

Ablation," NOLTR 68-163, U. S. Naval Ordnance

Laboratory, White Oak, Md. , Sept. 1968

(10) Kendall, R. M. and Bartlett, E. P., "A Procedure

for Calculating the Non-Similar, Laminar,

Chemically Reacting Boundary Layer over an Ablating Teflon Body with Coupled Quasi-Steady

or Transient Wall Boundary Conditions,"

Aerotherm Final Report No. 68-32, Aerotherm

Corp., Palo Alto, Calif., Apr. 1968

(11) Winkler, E. M. , Madden, M. T., Humphrey, R. L.

and Koenig, J. A., "Supersonic Ablation Studies

with Teflon," NOLTR 69-125, U. S. Naval Ordnance

Laboratory, White Oak, Md. , Oct. 1969

(12) Wentrick, T. , Jr., "High Temperature Behavior

of Teflon," AF BMD-TN-59-15, Research Report

No. 55, July 1959

(13) Gregory, N. , Stuart, J. T. and Walker, W. S.,

"On the Stability of Three-Dimensional Boundary

Layers with Application to the Flow Due to a Rotating Disk," Boundary Layer Effects in Aerodynamics , Philosophical Library, Inc. , 1957

(14) Knapp, C. F. and Roache, P. J., "A Combined

Visual Transition," AIAA Journal, Vol 6, 29,

1968

(15) Donaldson, Coleman, D. P. and Dunn, J. R.

"Final Report on Flow and Heat Transfer Studies

in Transverse Cracks," ARAP Report No. 133

(DASA 2301) , June 1969

(16) Settlage, P. H. and Siegle, J. C, "Keeping

Cool with 'Teflon'," an evaluation of "Teflon"

TFE resins as heat-barriers in high-speed

flight. The Journal of Teflon, Polychemicals

Dept. , Dupont Co. , 1960

(17) Scala, S. M. , "A Study of Hypersonic Ablation,"

Proceedings of the 10th International Astronautical Congress, Springer-Verlag, Vienna, 1959

864
(18) Siegle, J. C. and Muus, U. T., "Pyrolysis of

Polytetraf luoroethylene , " address to American

Chemical Society Meeting, Sept. 1956

(19) Winkler, E. M. , Humphrey, R. L., Koenig, J. A.,

and Madden, M. T. , "Investigation of Substructure

Heating on Cracked Ablative Heat Shields ,

NOLTR 69-122, U. S. Naval Ordnance Laboratory,

White Oak, Md. , July 1969
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869
SCALE 5 cm Scale 1 inch

FIG. 1 NOZZLE AND ABLATION DUCT ARRANGEMENT

THERMOCOUPLE COPPER CONSTANTAN

0.010

CRACK

FIG. 2 HEAT GAUGE INSTALLATION IN CRACK

870
871
LOCATION OF SKIN

FRICTION BALANCE

LOCATION OF 9.52 MM

0 50 100 150 (MILLIMETERS)

FIG. 4 TOTAL SURFACE RECESSION (OO MEASURED 180" APART)

872
873
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TESTING TIME (SECONDS)

FIG. 8 TEMPERATURE VARIATION FOR LONGITUDINAL CRACK

Testing Time ( seconds

FIG. 9 HEAT TRANSFER TO SIDE WALLS AND BOTTOM

OF A LONGITUDINAL CRACK

876
0.1

0.01

0.1

G.E.DATA AEDC

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(V\ALTA DATA

Re^=1600 NOL DATA

A NOL DATA

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• ' b/5«l,ReT* CO

b/5« 1, Ref-0

1.0 y/b

10 100

FIG. 10 TRANSVERSE CRACK HEATING SUMMARY

Courtesy of the Aeronautical Research Associ

OM -2.3

= 2.89 MN/M^ p = 2.23 MN/M ^

T° = 2405 K T = 2440 K

0 0 h/b = 4.1 h/b = 4

1 2 3 4 5 TESTING TIME (SECONDS)

FIG. 11 TEMPERATURE VARIATION IN TRANSVERSE CRACKS

877

Paper No. 5

APPARENT OPERATING LIMITS OF ARC HEATERS WITH RESPECT TO TOTAL ENTHALPY AND STAGNATION PRESSURE^

R. Richter"'"

REFERENCE: Richter, R., "Apparent Operating Limits of Arc Heaters With Respect to Total Enthalpy and Stagnation Pressure ,

ASTM/IES/AIAA Space Simulation Conference, 14-16 September 1970.

ABSTRACT: For many years arc heaters have been operated over a wide range of operating conditions. An apparent upper limit

with respect to stagnation pressure and total enthalpy was

found which could not be exceeded. It is attempted to show

that this limit is not imposed by the enthalpy-pressure combination but is caused by the limit in total power input that can

be achieved with the non-segmented constricted arc heater. The

failure or fast deterioration of arc heater components is found

to be primarily due to high arc currents which are employed to heat the gas to maximum temperature. It is concluded that the

enthalpy-stagnation limit of nonsegmented constricted arc heaters operating under similar constraints can be exceeded. KEY WORDS: arc currents;, arc heaters, constricted arc, enthalpy J operational limits, stagnation pressiire

NOMENCLATURE: C = Constant (Relation (1)); D = Diameter,

cm; H = Enthalpy, Btu/lb or joules/kg;^ I = Current, amps;

P = Power, watt; V = Voltage, volt; W = Mass flow rate,

g/sec; T = Temperature, °R; h = Heat transfer coefficient, Btu/sec-°R-f t ; k = Heat conductivity, Btu-f t/sec-°Rf t^ ; p = Pressure, atm; q = Heat load, Btu/ft^-sec or kW/cm^; t = Wall thickness, inch; a = Stress, psi

SUBSCRIPTS: A = arc; a = anode; all = allowable; c = constrictor; c = cathode; c = copper; M = melting;

T = total; W = water; o = stagnation.

Senior Scientist, Electro-Optical Systems, Pasadena,

California

2Work reported in this paper was performed under several NASA

and USAF contracts.

879
Arc heaters have long been employed for chemical processing

and reentry simulation because of their capability to raise

gases to high temperatures. Presently they are primarily used

for testing of heat shield materials for manned and unmanned

spacecrafts and missiles under conditions encountered during

reentry into the atmosphere of the earth or other planets. For the simulation of the environment encountered during

the reentry phase with respect to heat transfer and fluid dynamics, an arc heater has to heat air at a combination of high

enthalpy and stagnation pressure. The approximate operating

point of an arc heater for simulating the reentry conditions

into the earth atmosphere of a manned spacecraft is 40,000 Btu/

lb at a stagnation pressure of 1 atm, while that for simulating

the reentry conditions of a missile is about 4,000 Btu/lb at a stagnation pressure of 200 atm. The operating conditions for

the manned spacecraft have now been achieved, while attempts

are still under way to reach the operating conditions needed

for the simulation of the reentry conditions of missiles. The

obtainment of the desired conditions is apparently subject to limitations. Figure 1 shows the apparent operating limits of arc heaters with respect to stagnation pressure and total en- thalpy as they have been experienced by many investigators. It

is therefore of great interest to find the physical limits for

reaching, with arc heaters, the desired total enthalpy at a given stagnation pressure as well as to determine the operating

parameters that set these limits.

The limit to the achievable enthalpy at a given stagnation

pressure appeared primarily to be determined by the structural

integrity of the arc heater components. In attempts to raise

the total enthalpy at a given operating pressure, failures of

the electrodes and/or of the containment walls would occur. Failures of the electrodes would be attributed to exceeding the

average local heat load on an electrode surface which is a function of the total arc current, the operating pressure, and

the attachment point velocity over the surface. The failure of

the containment walls could be attributed to stray arc attach- ments and/or local over-heating.

During the development of several arc heaters and in the course of a special Air Force supported program, a better understanding of the apparent operating limits for arc heaters has

been generated. A substantial amount of experimental data was

obtained which permitted some correlations of the most important

parameters. But as had been found previously by many other

investigators, the number of operating parameters in arc heaters

is fairly large and their inter-relation rather complex. The

main parameters appear to be:

1 . Arc Current

2. Operating Pressure ., 3. Mass Flow Rate

4. Constrictor Design v ..^ _

880
a. Diameter

b. Length

c. Wall Thickness

d. Cooling

5. Electrode Design

a. Magnetically Forced Arc Movement

b. Magnetic Field Strength

c. Gas Induced Arc Movement

d. Vortex Strength

6. Electrode Separation

7. Gas Properties as Function of Pressure and Temperature

a. Radiation

b. Electrical Conductivity

c. Thermal Conductivity

d. Viscosity

e. Absorptivity

The possible number of permutations between these parameters is so large that it would be unrealistic to hope for es- tablishing absolute operating limits of arc heaters by an ex- perimental process of elimination. Some form of analysis has

therefore to be developed but whose basic assumptions are well

supported by a broad experimental background.

The first attempts in that direction have been made by a new analysis of the arc heating process in arc heaters. This

analysis still requires some additional refinements and extensions, though it already permits the evaluation and design of

constricted arc heaters over an operating regime in which absorption of radiation has no major influence. The author has

primarily worked with constricted arc heaters. However, many

of the results that will be discussed are applicable to all arc devices, radiation sources included.

Many attempts have been made to analyse the heating process in arc heaters, primarily of constricted arc heaters

(Refs. 1,2,3, and 4). These analyses did not predict too well

the experimentally obtained results of the author. The author

feels that the main reason for the discrepancies between analytical predictions and test data is the inconsistency between

the model underlying the analyses and the actual processes

taking place in the constrictor. Many investigators, for instance, have based their analysis on the laminar flow through a smooth tube (Refs. 1 and 4). This condition is rarely encountered in a segmented constrictor which has been the type of arc heater investigated by the author. Furthermore, most analyses

assumed the flow as well as the arc to develop in a continuous

process along the constrictor (Ref. 1). The results of these

analyses predict the heat transfer to the constrictor wall to be primarily determined by the local total enthalpy of the gas

The numbers in parentheses refer to the list of references

appended to this paper.

881
which increases along the constrictor. This, however, was not borne out by our test data and therefore a different analytical

approach appeared to be indicated.

The author approached the analysis of the very complicated

arc heating process by synthesizing the various physical pro- cesses which apparently take place in the constrictor of an arc heater. These processes are individually analyzed and then com- bined by demanding an overall energy balance as well as stipulating an arc operation with minimum entropy rise. The com- plete analysis will be presented at a later date in a separate

paper

The main difficulty in applying the synthesized analysis

to test data and/or the design of arc heaters lies in the inavailability of fully substantiated material properties for the

various gases. The author employed primarily the thermal con- ductivity, radiation, electrical conductivity, and viscosity

data of Ref. 5. With these material properties for nitrogen,

fairly good agreement between test data and calculated heat

transfer rates and electric fields were obtained over a very

wide range of operating conditions, i.e., from less than 1 atm

to 97 atm. The analysis was also applied to data reported in

the literature. A good agreement with measured arc diameters

of an arc operating in a cross flow was found (Ref. 6). The

largest variations between test data and calculated values was

about 20 percent. This discrepancy is quite acceptable considering not only the uncertainties associated with the measurement of absolute local heat transfer rates and electric fields

in arc heaters but also the uncertainty in the materials properties used.

Two types of constricted arc heater designs have to be

distinguished. In one type of design the electric arc operates

with what the author will call a "natural" arc length, while in

the other type of design the arc length is determined by the

distance between the two electrodes. These two types of heaters

are illustrated in Figure 2. Many variations for each type of

heater are possible, though the basic design must be considered

unalterable. It is also possible that in both types of arc heaters the polarity of the two electrodes might have an influence but it will become clear from the future discussions

that such influence must be minor and can be included into the

overall picture of arc heater operation.

In the first type of arc heaters which is shown schematically in Figure 2a, the natural arc length appears to be a function of all major operating parameters, i.e., mass flow

rate, operating pressure, arc current, and gas, as well as the arc heater component configuration and the magnetic field

strength, if a magnetic field is employed for the control of

the arc. When correlating a substantial number of experimental

The numbers in parentheses refer to the list of references

appended to this paper.

882
data of non-segmented constricted arc heaters (Refs. 7 and 8), an interesting relation between the major parameters appears to be applicable. This relation could be best expressed by

(1)

The exponents of the parameters were found to be m = 0.835

= 0.129

n = 0.267

X = 1.34

and the constant for air, C = 1617.

Naturally this relation cannot be justified in the strictest sense of an analysis as it is well known that the total

voltage is the sum of several individual voltage drops, all of

which are dependent on the operating parameters in a different

way , i.e.,

In most cases the arc voltage is very much larger than the sum

of the electrode voltages. This is especially true for high

power, high pressure arc heaters.

But, nevertheless, the relation as it stands conveys the

interesting effects of the most important parameters on the

total voltage. .The total voltage-arc current characteristic

shows a negative slope with a power of approximately 0.267.

This slope prevails at current levels at which investigators of

stationary arcs have observed arc operation with a positive

slope. The negative slope of nonsegmented arc heaters has

therefore to be attributed to a shortening of the arc length

with increasing arc current. This deduction is further substantiated by the positive voltage current characteristics of

arc heaters which do not permit the arc to shorten itself (Refs.

9 and 10). If the total voltage is divided by the electric

field which was found to be also a function of the major operating parameters, the effect of the major operating parameters

on the arc length would be expressed.

When multiplying the total voltage by the arc current,

relation (1) leads to a relation between the major operating

parameters and total input power

The numbers in parentheses refer to the list of references

appended to this paper.

= V + V + V,

T a c A

(2)

^T

= (3)

883
c i,^-"

P D

O C From this relation it is seen that the total input power, P,p, is primarily a function of the arc current, the main flow rate,

the operating pressure, and the constrictor diameter. The mass

flow rate and the constrictor diameter have a strong influence

in contrast to the relatively weak influence of the operating

pressure. The medium influence of the arc current on the total

power input was always quite obvious by the negative slope of

the voltage-current characteristic of such heaters. But the

most important aspect of this relation is the influence of the mass flow rate, W, on the total power input. The relation

points out quite clearly that if a current limit exists, then

there exists also an enthalpy limit. Because the mass averaged

total enthalpy is a fraction of the total input power divided

by the mass flow rate, the faction being the efficiency, T],

"t = (5)

D p W

c ^o

The efficiency was found to be a strong function of the arc current. If we set as a first approximation

where the value for z is found to vary between 0.12 and 0.35,

the enthalpy, H^, is given by

c o The exponent 1-n-z is very much less than 1. For that reason we find that for a given arc heater configuration a limit to the achievable total enthalpy at a given pressure exists re- gardless of arc current. The important results of this exercise are the following:

1. The arc heater does not have an operating limit

that is determined by the total enthalpy of the gas. Quite the opposite, the enthalpy limit is

determined by the arc heater.

884
2. An enthalpy limit is established by the total electric

power input for a given gas flow rate and arc heater

configuration, as indicated by the strong influence of

the constrictor diameter and the mass flow rate on the

total voltage.

Relation (7) seems to indicate that the total enthalpy at

a given pressure could be increased by decreasing the constrictor diameter, . This is indeed possible. But, when decreasing the constrictor diameter, a new limit is being approached.

When a segmented constricted arc heater was operated over

a wide range of operating conditions, heat transfer rates to the constrictor wall could be measured under various operating

conditions. In Figure 3 the heat loads on a constrictor when

operating with a stagnation pressure of around 1 atm are shown

(Ref . 9) while in Figure 4 heat transfer rates for a second con- strictor operating at stagnation pressures ranging from 38 atm

to 51 atm are shown (Ref. 10). It can be seen in both figures

that the heat transfer rates are almost constant over the entire

length of the constrictor. This result seems to indicate that

the heat loads under the conditions under which the heaters were operated are predominantly a function of the arc current.

Some arc devices were tested under conditions and with constrictor configurations for which this was not at all the case (Ref.

9). Under some conditions it could be clearly shown that the

measured heat transfer rates were the sum of two heat transfer

processes; one as shown in Figures 3 and 4 was almost constant

along the constrictor, the other was directly proportional to the local mass averaged total enthalpy. But those were cases which the author does not consider applicable to the present

discussion. In Figure 5, calculated heat loads for various arc currents and constrictor diameters are shown. Experimental

correlations between local heat transfer rates and constrictor

diameters are not available as it was not possible to build a representative number of segmented constrictors of different

diameters. Since good agreement was found between measured and

calculated heat transfer rates, the use of analytical values

can be justified. These correlations show that the heat load

increases quite rapidly with decreasing constrictor diameter

and that, therefore, a limit exists to how small a constrictor

diameter can be made.

This limit to the permissible heat load cannot be a simple

one. It is determined by the operating pressure as well as the

yield strength at elevated temperature of the material of which

the constrictor is built. For any given heat load the wall

thickness of a container under pressure has two limits, one

limit is set by the maximum allowable temperature of the material; the other limit by the tensile or compressive strength of

the material as a function of the effective temperature. If

The numbers in parentheses refer to the list of references

appended to this paper.

885
the allowable temperature is the melting temperature, T , then

the maximum wall thickness as a function of heat load is given

by

_ ' t = k - - (8) For simplicity, a thin wall tube is being considered. When the same assumption is applied to the stress consideration, the

relation between pressure and wall thickness is given by

Ap D " all

Elimination of the wall thickness t between relations 8 and 9 produces the following limit

r^p °c k^

q = k (T + T ) / ^- — + ^ (10) ^all ^m w 2 a,, h

all

This relation has been plotted for copper into Figure 5. The

combined results show clearly the constraint on the size of the

constrictor diameter.

SUMMARY

The preceeding discussion has attempted to explain the observed

stagnation pressure - total enthalpy limit of nonsegmented con- stricted arc heaters by the rather complex relations between

the major operating parameters and the physical dimensions of

the arc device. When expanding the experimentally determined

relation between total voltage, constrictor diameter, arc current, mass flow rate and stagnation pressure it is realized

that the apparent enthalpy limit that was found experimentally

with such heaters is effectively a design limit. The destruction of heater components is not at all caused by the stagnation pressure-total enthalpy combination. The destruction or excessive deterioration of components has to be attributed to the high arc current which is employed in an unsuccessful at- tempt to overcome the basic limit on the total power input imposed by the heater design. The stagnation pressure - total

enthalpy limit has been exceeded by the use of a segmented con- strictor which makes the arc length independent of the stagnation pressure, the mass flow rate, and the constrictor diameter.

Arc heaters with segmented constrictors are subject to new

limits whose dependence on the major parameters are not included in this discussion.

886
REFERENCES

1. Stine, H. A. and Watson, V. R. , "The Theoretical Enthalpy

Distribution of Air in Steady Flow Along the Axis of a Direct-Current Electric Arc," NASA TN D-1331 (1962).

2. Stine, H. A., Watson, V. R., and Shepard, C. E., "Effect

of Axial Flow on the Behavior of the Wall-Constricted Arc," NASA TM X-54,065, presented to the AGARD Specialists'

Meeting on Arc Heaters and MHD Accelerators for Aerodynamic Purposes, Rhode-Saint-Genese , Belgium, September

21-23, 1964.

3. Weber, H. E., "Growth of an Arc Column in Flow and Pressure Fields," AGARD Specialist Meeting on Arc Heaters and MHD Accelerators for Aerodynamic Purposes, Rhodes-SaintGenese, Belgium, September 1964, (Published as AGARDograph

No. 84, pp 845-881)

4. Skifstad, J. G. and Murthy, S. N. B., "Analysis of Arc- Heating Phenomena in a Tube," International Symposium on

Plasma Phenomena and Measurements, lEEE-PTGNS, Tenth Anniversary Meeting, San Diego, California, October 29 - November 1, 1963 (IEEE Transaction on Nuclear Science,

January 1964, Pg. 92).

5. Yos, J. M. , "Transport Properties of Nitrogen, Hydrogen,

Oxygen and Air to 30,000°K," AVCO Tech. Memo RAD-TM-03-7,

22 March 1963.

6. Roman, W. C, and Myers, T. W., "investigation of Electric Arc Interaction with Aerodynamic and Magnetic Fields," ARL

Report 66-0191, October 1966.

7. Eschenbach, Richard C, Skinner, G. M. , and Arc Laboratory

Staff, "Development of Stable, High Power, High Pressure Arc Air Heaters for a Hypersonic Wind Tunnel," WADD Technical Report 61-100 ASTIA AD No. 266 907 (July 1961).

8. Brown, St. George, and Patton, J. B., "Calibration and

Operation of the Linde Model N4000 Arc Heater with 1/2-

and 3/8-inch-Diameter Constrictors," AEDC-TR-65- 102 , June

1965.

9. Richter, R. , and Buhler, R. D., "High Enthalpy Gas Heater,"

Final Report Contract No. NAS9-3564 (December 1965).

10. Richter, R. , "Ultrahigh Pressure Arc Heater Studies," AEDC

Report No. AEDC-TR-69-180 (September 1969).

887
TTTT I I I I 1 1

1

1

\—I I I I 1 1 1| \—

<

I

TESTED REGIME WITH

SEGMENTED CONSTRICTED

ARC DEVICES

10

UJ

oo 3 < 10 ^ 0.

APPARENT ENTHALPY

STAGNATION PRESSURE

LIMIT OF NON-SEGMENTED

CONSTRICTED ARC DEVICES

111 I I I I I I II J I I Mill

10 100 STAGNATION PRESSURE p^.ofm

J L±

500

Fig. 1--Experimentally established operating regimes of con- stricted arc heaters.

888
AIR

a. NONSEGMENTED CONSTRICTED ARC HEATER CONFIGURATION

b. SEGMENTED CONSTRICTED ARC HEATER CONFIGURATION

Fig. 2--Two basic types of constricted arc heater

configurations

889
1

\—I\—I—I

\—I

\—\—\—I

\—\—I

\—

W= 6g /sec

2 5 8 II 14 17 20 SEGMENT

Fig. 3--Measured heat transfer rates along the constrictor of

a low pressure arc heater 1 atm).

890
SEGMENT

T

1 1 10 n 12 13 14 15 1

\—\—\—I—

2 '0 COi

<

O

RUN PRESSURE Po , atm ARC CURRENT, A A 74 51 410 O 72-2 40 395 • 73-1 42 375 72-2 31 375 0 72-3 72-4

3335 380 ^SEG 3- 72-1 29 375 0 72-5 38 385

J L J \ L_J \ I \ [ 6 7 8 9 k5 11 12 13 14 15 SEGMENT

1 2 3 4 5 6 7 DISTANCE FROM CATHODE J^c.cm

Fig. 4--Heat transfer rates along the constrictor of a high

pressure arc heater.

891
E

6

10

<UJX

I

9

I—CO

Z

o

u

NOMINAL 5-MW CONSTRICTED ARC

HEATER

OPERATING PRESSURE Po= ^00 atm

MASS FLOW RATE W = 200g/sec

RADIATION HEAT LOAD q,. a» ARC CURRENT I*

V

ALLOWABLE

HEAT LOAD

h = 10 Btu/hr-ftVR

= 8000 p$i

Ap = 50 arm

ALLOWABLE

HEAT LOAD

^all °*

4 h = 10 Bfu/hr.*R.ft

(7^11= 3000 p$i

Ap = 100 otm

I

1

,000 A

750 A

500 A

250 A

ALLOWABLE HEAT LOAD q

°* = OO

O^ll = 3000 psi

A p = 100 atm

all

CONSTRICTOR DIAMETER , cm

Fig. 5--The effect of constrictor diameter on the maximum heat

load and the allowable heat load.

892
Paper No. 55

OPERATING CHARACTERISTICS OF THE AIR FORCE FLIGHT

DYNAMICS LABORATORY REENTRY NOSE TIP (RENT)

FACILITY

J. C. Beachler

Research Aerospace Engineer, Air Force Flight

Dynamics Laboratory, Wright-Patterson AFB, Ohio

ABSTRACT: Reentry missile nose tip development

requires ground test facilities capable of simultaneously producing extreme pressures and heat transfer rates. The Reentry Nose Tip (RENT) Facility

was developed by the Air Force Flight Dynamics

Laboratory to produce this high pressure hyperthermal environment. The present and planned capability of this facility is described and typical

pressure and heat transfer profiles of the test

jet are shown.

KEY WORDS: reentry, test facility, heat protection,

ballistic missiles

The continuing requirement for higher performance ballistic reentry vehicles has placed greatly

increased demands on thermal protection techniques.

Improved ablation materials and several new heat

protection methods have been brought to bear on the

problem

These programs have required extensive ground

and flight test efforts to screen and develop better

materials and methods for heat protection. The

893
Reentry Nose Tip (RENT) Facility was recently developed by the Air Force Flight Dynamics Laboratory to extend the nation's ground test capability for

thermal protection system evaluation.

Facility Description and Performance

The RENT Facility is a high pressure high

power arc heated continuous supersonic wind tunnel

which produces a high pressure high shear test

environment with stagnation temperatures up to 14,000°R. Figure 1. is a schematic of the facility.

The primary component of the Facility is its

high pressure high voltage arc heater which heats

up to 8 Ibm/sec of high pressure air to stagnation

enthalpies of 6400 Btu/lbm at 125 atmospheres pressure using 37 megawatts of electrical power. This

air is expanded from a .900 inch throat to Mach 1.8

at the nozzle exit. A cross section view of the

heater is shown in Figure 2. The test models are injected to the arc jet

centerline at a station .100 inches downstream of

the nozzle exit plane by a hydr aulically actuated

linear drive system capable of producing model

speeds from 4 to 150 inches per second. The model

carriage drives 5 models and/or probe struts through

the test jet in either a sweep or sweep-pause mode.

This system produces up to 7g decelerations as it

stops the model in the test flow.

The RENT exhaust system consists of a collector

cone 18 inches downstream of the nozzle exit followed by a water-spray-cooled 18 inch diameter duct

leading to a 7 foot diameter centrifugal water

separator and a 40 foot long vertical exhaust stack.

The operating range of the facility was determined in a series of tests conducted in mid-1969.

A typical operating characteristic for 2400 amps

current and a nozzle throat diameter of .900 inches

is shown in Figure 3. The DC power to the arc heater is determined by monitoring the operating

voltage and current. The air mass flow rate is

measured with a subsonic venturi located upstream

of the heater. The air pressure in the heater was measured at the outside wall of the swirl chamber

and at the center of the rear electrode as indicated

in Figure 2. Based on References 1-3 the rear electrode pressure (Pqr^ considered to be the most

representative heater stagnation pressure. The flow

rates and the temperature rise of the heater cooling

894
water were used to determine the energy loss to the

heater for calculation of the heater efficiency and

heat balance enthalpy:

= [EI - (w Cp AT)„^Q]/i^,^

Since the heat balance or bulk enthalpy is an average measurement, it is typically not a true test

jet centerline value due to the use of highly

cooled nozzle and front electrode walls and the

centrifuging action of the arc stabilizing vortex

which tends to concentrate the hot, less dense

plasma near the centerline of the heater. However,

since accurate direct measuring total enthalpy and

heat flux probes are beyond present technology for

the conditions of these tests, the data is presented

using the heat balance values of . A summary of

facility performance is shown in terms of arc heater

stagnation pressure and heat balance total enthalpy

in Figure 4.

Typical radial impact pressure profiles taken

with graphitic probes for nominal Mach 1.8 conical,

flared and contoured nozzles are shown in Figure 5. The maximum heater pressure (P^^) of 125 atmospheres

produces a model stagnation pressure of about 100

atmospheres at a station 0.100 inches downstream of

the nozzle exit. Radial and axial pressure surveys

have been made with both conical and flared expansion contours on d* = 1.00 and 0.900 inch nozzle

throats. Reference 4 discusses the performance and

relative merits of each of these nozzles for high

shear tes t ing

A characteristic stagnation point heating rate

profile taken with null point calorimeters in the

contoured nozzle at 100 atmospheres heater pressure

is shown in Figure 6. The ratio between centerline

and intermediate plateaus of heating rate is approximately 2:1. The centerline estimated for laminar heating is 6000 Btu/lbm which is about 2.5 times

the heat balance value.

Pressure and temperature measurements were made

to determine the mass handling and cooling system

performance of the RENT Exhaust System. Facility

operation demonstrated that it had adequate cooling

and mass handling capacity at maximum heater operating conditions. With an arc heater mass flow of

6.5 lb/sec an additional 24 lb/sec of air is en- trained through the exhaust system by the test jet.

895
Acoustic levels of up to 163 dbm with a frequency peak at 1000 Hertz were observed with a microphone at the exhaust collector cone. Since the

facility is enclosed by concrete and an acoustic

barrier, this sound level caused no damage to equipment and all operational personnel were safe without

ear protection at their normal work stations. Power

spectral density analyses indicated the sound was generated at the free jet boundary.

High speed photography, high speed digital data

and pyrometer measurements are normally used for

analysis of ablation tests conducted in the RENT

Facility

Arc Heater Performance

The general operating characteristics of the

N = 4'^ heater with 1.00 inch and 2.00 inch nozzle

throats are described in Reference 2. However, the

high shear tests in the RENT required extending the

pressure capability to the limit of a new less

restrictive air supply system and improving heater

enthalpy at all conditions.

Tests were made with a 96 inch front electrode

to determine the maximum P^j^ attainable without

blowing the arc through the heater nozzle throat.

An extended pressure range is available when a front

field coil is used to hold the downstream arc termination in the heater. The data in Figure 3 represents the maximum achievable Pqj^ with the 0.900 inch

nozzle throat as limited by the existing air supply

capacity.

A front electrode burnout problem previously

thought to be a radiation heating performance limit

(Reference 2) was found to be a backside cooling

problem and has been remedied by redesign of the

cooling passage. Cooling water velocities have been

raised from 80 to 120 ft/sec to prevent film boiling

on the electrode.

A test series was run with a 96 inch front

electrode to determine the maximum attainable

enthalpy at heater pressures corresponding to a model stagnation pressure of 60 atmospheres. Since

the impact pressure at the model test station was about 77% of Pqr* tests were conducted at a heater stagnation pressure of 1100 psia. The heater

enthalpy increased with increasing arc current up to

the 5400 amp le vel where front electrode erosion at

"'"Heater scale factor described in Reference 3.

896
the arc termination became excessive. Improvements

in front electrode cooling permitted extending the

operable arc current from 2800 to 5400 amps at this

pressure. All figures are shown with from heat

b alance

Summary and Future Plans

The pressure capability of the AFFDL N=4 arc heater has been extended to 123 atmospheres at 37

MW through improved front electrode cooling, the use of a front spin coil and improvements in the air

supply system. The current capacity of the heater

has also been extended from 2800 to 5400 amps at 75

atmospheres through improved electrode cooling.

Facility operation at maximum heater conditions

has demonstrated the adequacy of the RENT exhaust

and model injection systems and shown that sound

levels produced by the test jet are acceptable.

Pressure and heat flux surveys of the test jet

with conical and flared nozzles and ablation tests

for AF Materials Laboratory have demonstrated the

quality and steadiness of the test flow. Use of a front spin coil has also precluded the possibility

of arcing to the test models and assisted in producing uniform pressure profiles with fluctuations

of less than ±5% in jet impact pressures. Test

flows appear very clean and axisymmet r ic

The RENT Facility is operational and has been

routinely performing ablation tests for the Air

Force Materials Laboratory and the Air Force Space

and Missile Systems Organization.

Future plans include extending the operating

pressure of the N=4 heater to 140 atmospheres,

design and development of a 200 atmosphere heater

and increasing the heater power supply capacity.

897
REFERENCES

R. T. Smith and W. A. Kachel , "Design and Performance of a Small (8-Megawatt) Arc Heater for

the AFFDL 50 Megawatt Hypersonic Facility,"

AFFDL-TR-68-17 , February 1968.

R. T. Smith and J. L. Folck, "Operating Characteristics of a Multi-Megawatt Arc Heater Used

with the Air Force Flight Dynamics Laboratory

50 Megawatt Facility," AFFDL-TR-69-6 , April

1969.

R. C. Eschenbach and G. M. Skinner, "Development

of Stable High Power, High Pressure Arc Air

Heaters for a Hypersonic Wind Tunnel," WADD-TR61-100, Linde Company, Indianapolis, Indiana,

July 1961.

F.J. A. Huber, "A Tailored Plume Nozzle for

Testing Fast Reentering Missile Nose Tips,"

Paper to be Presented at AIAA 6th Aerodynamic

Testing Conference, March 1971.

898
899
900
FIGURES N=4 OPERATING CHARACTERISTICS, d* = .900, 96 INCH FRONT

ELECTRODE, WITH FRONT COIL

901
902
Po^ = 1500psi

•arc

= 2600 A

d* = 0.900 in.

RADIAL DISTANCE FROM NOZZLE Cfc , inches

Axial Station 0.100 from Exit Plane

FIGURE 5 IMPACT PRESSURE

PROFILES OF RENT NOZZLES

903
RADIAL DISTANCE FROM NOZZLE <t , inches

Axial Station 0.100 from Exit Plane

FIGURE 6 HEAT FLUX PROFILE

OF CONTOURED RENT NOZZLE

904
Paper No, 56

INVESTIGATION OF THE CHARACTERISTICS

OF FLUIDS VENTED INTO A VACUUM

Riddle Steddim,^ Dupree Maples, and Frank M. Donovan ^

ABSTRACT: For manned space stations planned for the future,

it becomes necessary to eject fluids into the space environment. These ejected fluids form ice clouds which remain in

the vicinity of the spacecraft and could interfere with optical instrumentation. In order to study the dynamics of fluid

jets in a space environment, a high vacuum system was modified

to allow the injection of fluids into a vacuum region while

maintaining a pressure below 0.1 N/m . The cone angles, individual particle velocities, and mass flow rates of water

sprays under vacuum conditions were measured. The individual

particle velocities of sprays produced by outgassed water were

lower than the particle velocities of sprays produced by water with a definite amount of dissolved gas. KEY WORDS: vacuum tests, particle velocities, flow rate tests,

high speed movies, water jet

INTRODUCTION: During long duration space missions liquid wastes must be ejected from spacecrafts. The behavior of liquids in a space environment becomes important because the

possibility exists that the ice particles formed from ejected

liquids could interfere with optical instrumentation.

Major characteristics of jets produced by ejecting fluids

into a space environment which should be investigated are the

total included cone angle of the spray; the velocities of individual particles within the spray; the size, shape, and distribution of particles; the state of the particles, i.e.,

whether they are liquid or solid; and the total volume flow

rate. The above characteristics are functions of at least two important general parameters: the nature of the ejected fluid

- its chemical and physical composition and its thermodynamic

^Louisiana State University, Baton Rouge, Louisiana

905
properties; and the physics of the injection process - orifice

size and geometry and injection pressure and temperature.

Some previous investigations of this problem include a study presenting both theoretical and experimental work by

Simmons, et al of the Atlantic Research Corporation (1)". The

experimental program included static tests where bulk samples

of liquid were suddenly exposed to vacuum, orifice and gasket

leak tests, and vent line tests. Some work was done to determine individual particle size and velocity by using still photographs. Other work includes that done at the Boeing Com.pany

by Dawley (2). This investigation had three stated objectives

to confirm that there exists a problem of flow stoppage due to freezing; to determine the effects of flow rates, water tem- perature, and initial nozzle tip temperature on nozzle perform

ance; and to record the ice formations photographically, A

report by B, L. Mann and 0. T. Stoll (3) also deals primarily

with the problem of flow stoppage due to freezing. Jet dynamics were not considered. In the Journal of Spacecraft and

Rockets, Mikatarin and Anderson" (4) report studying liquid

jets expelled into a vacuum, but could not obtain particle

velocity measurements, A report by the Brown Engineering Company's Research Laboratories (5) considers work which has been

done to determine the drop sizes which form from jets and the

expected life times of ice spheres in a space environment as

a function of radius.

This investigation has been directed toward the study of cone angles and individual particle velocities of water jets

under vacuum conditions. Sprays produced by distilled, out- gassed water were compared to sprays produced by water with a known amount of gas (CO2) dissolved in it as a function of injection pressure and orifice size. The degree of vacuum was kept below 0, 1 N/m^. EXPERIMENTAL APPARATUS AND PROCEDURE: This investigation was conducted in a Murphy and Miller Vacuum Chamber at the National Aeronautics and Space Administration's Mississippi Test Facility - an overall view of the apparatus is shown in Figure

1, The test chamber is approximately 1,25 m in diameter by

1,5 m long and is evacuated by a Stokes mechanical pump, a Roots high volume blower, and an oil diffusion pump. The sys- tem is capable of producing a minimum pressure of 3.33 x 10"-^ N/m^, Chamber pressure was measured by a thermocouple and an ionization gage. A hole 0.6 m in diameter covered by a plate

provided a mounting for various f eedthroughs

A cy linderical , liquid nitrogen cooled, cryogenic liner was designed and installed within the test chamber to provide

cryopumping for the water vapor. A fluid injection system was

"The numbers in parentheses refer to the list of references appended to this paper

906
also designed and installed in place of a glass port on the

door of the test chamber. A close-up view of this system is shown in Figure 2. This system consisted of a tank approximately lO'-^m^ capacity which held the fluid to be injected

into the vacuum system, a slight glass arrangement in the side

of the tank, a valved fill-drain line and a pressurization

line, a plunger for starting and stopping the injection of the

fluid, a threaded hole for mounting various nozzles and orifices, a hot water line for heating the nozzle, and a small

port and mirror for close viewing of the nozzle surface. The vacuum side of this system is shown in Figure 3. Tank pressurization was supplied by a cylinder of high

pressure gas which fed a ballast tank (capacity approximately

lO'^m-^) through a pressure regulator. The ballast tank was connected directly to the fluid injector tank by a rubber vac- uum hose. Ballast tank pressure was measured by a pressure

gage which read from 30 inches of mercury vacuum to 15 pounds

per square inch above atmospheric pressure. This fluid pres- surization system is shown schematically by Figure 4. The fluid spray was photographed using both still and

high speed cameras. The still photographs were taken with a 4x5 inch view camera with a 135 mm focal length lens. Exposure times varied between 1/60 and 25 seconds. The high speed

motion pictures were made with a model WF-145 Wollensak Fastex

"S" high speed camera. The particle velocity measurements

were made from movie film taken at approximately 4,000 frames

per second with a 35 ram focal length lens. Successful high

speed motion picture photography of sprays in a vacuum depends

greatly on obtaining proper lighting. Experience gained thus

far in this project indicates a modified backlighted spray

photographs best. A diagram of the lighting arrangement used

in the footage f-rom which velocity measurements were made is shown in Figure 5. The fluid tested was distilled water which was boiled for

one hour to expell dissolved gasses and then cooled in a closed flask. The first series of tests used helium as the pressurizing gas because it is nearly insoluable in water. In the

second series, carbon dioxide was used as the pressurizing gas

and is quite soluable in water. The mole fraction of gas dissolved in the water was obtainable from Henry's Law, The water was injected into the vacuum at temperatures between 24C

and 27C. RESULTS: During the first few runs, a slot approximately

3 . 2 mm wide was put directly in front of the spray in order to obtain a thin section through the middle of the spray and

hence photograph only a section of the spray parallel to the

film plane. Icing problems were anticipated for previous

investigators (references 1 and 2) had reported the formation

and growth of ice structures whenever a water jet struck a surface. Therefore, the slot was provided with a heater.

907
Nevertheless, ice would still form on the slot, occasionally

bridging the slot and blocking flow although the slot material

was heated to a temperature in excess of 38C. Although the

slot material was kept in excess of 38C, generally one-half

hour was required for the ice formation to melt sufficiently

for gravity to break it away from the slot. One other problem

arose with the slot: the sprays tended to be unstable and

sweep from side to side and miss the slot. T?iis was particularly true of sprays produced by the smaller orifices sizes

and at lower pressures. These sprays would issue from the

orifice as a thin, pencil-like jet for a length of several jet

diameters and then break into a cone spray of small particles. Just before this pencil-like jet broke into a spray it would

begin to weave, causing the entire cone to sweep. For these reasons the slot was abandoned and the entire spray photographed . A series of tests were conducted to obtain particle velocities as a function of injection pressure for outgassed water

and for water with dissolved carbon dioxide in it. The re- sults of this series of tests is presented in Figure 6. Each

velocity plotted is the average of the velocities of approximately twenty particles tracked at each pressure. These velocities were obtained by photographing the spray with a high

speed 16 mm movie camera. By proper choice of camera location

and lens focal length, it was possible to photograph the spray

such that individual particles could be tracked for about fifty frames. This corresponded to an actual distance of about

150 mm. A time base was provided by timing marks put on the

edge of the film at one thousand per second. Due to reasons discussed above it was not possible to obtain only a section

of the spray, but rather the entire spray had to be photographed; thus, some dispersion of particle velocities would

be expected because many particles would not be moving exactly

parallel to the film plane. This dispersion did occur but

not to the degree which might have been expected. This is probably due to small cone angles. Table 1 compares results

from this work to the results obtained by the Atlantic Re- search Corporation group (1). The latter did measure the

velocity of particles which were moving only parallel to the

film plane. However, these velocities were obtained using

still photographic equipment and a single flash strobe and

the time which a particle track registered on the film was difficult to measure and they reported that their tabulated

velocity measurements were probably in error by a factor of

two. Note also the large range of velocities reported. This

group did not attempt to outgas the water. Since the range of velocities observed in this work is not unduly large and the velocities fall within the range of the few observations of previous investigators, it is felt

that the average velocities presented in Figure 6 are representative of the particle velocities of the entire spray under

908
the various conditions. There was some error because some particles were not moving parallel to the film plane. Corresponding to earlier repo ts (1), many particles appeared to be

rotating. When viewing the high speed movies, many particles

appear dim but their brightness will increase sharply and then

decrease as they move across the screen. This is probably due

to specular reflections from uneven particle surfaces.

Another series of tests was conducted to measure the

volume flow rate of outgassed water through seven nozzle sizes

(orifice diameters of 0.34, 0.51, 0.66, 0.79, 0.84, 1.0,

and 1.17 mm) at five injection pressures. These results are presented in Figure 7. As would be expected and as seen from

Figure 7, the volume flow rate and injection tank pressure are related by the following equation:

V = cp2 where: V = volume flow rate

p = injection pressure

c = constant (different for each nozzle)

Included cone angles as a function of injection pressures as shown in Figure 8 were obtained from still photographs of sprays produced by outgassed water. It appears from this

figure that cone angles and injection pressure could be cor- related by an equation of the form:

a = ap'^ where: a = cone angle

p = injection pressure m,a = constants. A problem arises when attempting to obtain cone angles of sprays photographically. The sprays are dense near the center

and become progressively less dense away from the axis of the

jet and no clear boundary at the jet's edge exists. The intensity of the photographic image of the spray depends on the

size and concentration of the drops comprising the spray, the

arrangement and intensity of the lighting of the spray, as well as the film speed, exposure time and aperture used. It

is quite likely therefore, that small and widely separated

drops at the periphery of the spray did not register on the film and the total included angles of the sprays were some- what larger than those measured. The data shown in Figure 8 for the large nozzle (0.99 mm) does not have nearly as much

scatter about its representative line as does that of the

small nozzle (0.51 mm). It is felt that this is due to the spray from the small nozzle being finer and more diffuse than

the spray from the larger nozzles, making it more difficult to correctly define the boundaries of the small jet.

CONCLUSIONS: Experience gained in this program indicates that

909
icing will occur whenever a spray impinges on a surface, even

a relatively hot surface. The work accomplished to date dem- onstrates that sprays in a vacuum can successfully be analyzed

photographically if the spray is properly lighted. Additional

work is proceeding to determine particle velocities, sizes,

and distributions within water sprays for various nozzle sizes,

injection pressures, and dissolved gas content. Another series

of experiments will attempt to define the freezing time of the

individual particles comprising the sprays. REFERENCES: •

1. Simmons, Gift, and Markels, "Investigation of the Effects

of Vacuum on Liquid Hydrogen and Other Cryogens Used on Launch Vehicles," Final Summary Report, Atlantic Research

Corporation, Contract No. NASA-11044, December 18, 1964,

NASA Assesstion X65-12159.

2. Dawley, Richard A., "Expulsion of Water Based Liquids into

a Hard Vacuum," 02-84191-1, Boeing Company, 1966 (N66-

30440). .

3. Mann, B. L. , and U. T. Stoll, "Experimental Investigation

of the Freezing Problem Presented in Discharging Water

Based Fluids to Space," SAE Paper 912D, SAE/NASA Engineering & Manufacturing Meeting, Los Angeles, California,

October 5-9, 1964.

4. Mikatarian, R. R., and R. G. Anderson, Journal of Space - craft and Rockets

Vol. 3, No. 2, February, 1966, pp. 267-268.

5. Sodek, B. A., "The Resultant State of Water Expelled into

the Near-Ear th Orbital Environment," RL-SSL-207, Brown Engineering, July, 1968.

6. Brennan, J. A., "A Preliminary Study of the Orifice Flow

Characteristics of Liquid Nitrogen and Liquid Hydrogen

Discharging into a Vacuum," Advances in Cryogenic Engineer - ing , Vol. 9, Plenam Press, New York, 1964.

7. Schmidt, J. M. , "An Experimental Study of the Behavior of Liquid Streams Injected into a Low Pressure Chamber," Jet Propulsion Lab,, California Institute of Technology,

Pasadena, California, April, 1949,

910
Table 1 Comparison of the Distribution of

Particle Velocities Between Current

Work and Previous Work

Velocities from Velocities Presented

this Investigation in Reference 1

Injection Pressure: Injection Pressure:

85.5 KN/m^ 101.3 KN/m^

Fluid: outgassed water Fluid: water

Particle Velocity -

s Particle Velocity —

s 16.05 11.05

16.55 42. 10 17.30 21.00

15.00 22.70

16.05 13.35

14.80 22. 10 15.45 10.18

14.55 23.20

16.55 34.60

15.05 12.37

15.30 10.61

14.70 10.40

16.40 12.03

16.75 11.17

14.70 11.70

16.90

15.90

16.30

911
Fig. l--Overall View of the Test Apparatus

Fig. 2--Injection Tank

Sys tern

912
913

Paper No. 57

MICROMETEOROID IMPACT SIMULATION SYSTEM

1 2 Robert A. Golub and John R. Davidson

REFERENCE: Robert A. Golub and John R. Davidson, "Mlcrometeoroid Impact Simulation System," ASTM/IES/AIAA Space Simulation

Conference, lk-l6 September 1970.

ABSTMGT : This paper describes a Micrometeoroid Impact Simulation System (MIMS) at NASA Langley Research Center and outlines current research efforts. The MEMS electrostatically accelerates electrically charged, micrometer- size particles to velocities in excess of 30 ktn/s. Particles can be accelerated one at

a time or at various rates to above lO/s. The major components

of the MIMS are a rnicroparticle charger- injector , a horizontal

4-million-volt Van de Graaff accelerator, an assemblage of

particle detectors, a particle deflection system, a data system,

and a series of target chambers. The data system is a computercontrolled, real-time system which records information to determine the velocity, mass, and diameter of each particle which

enters the target area. The system also controls and selects

particles, and rejects particles which do not have the desired

velocities. The system is used to study damage to sensitive

surfaces, develop micrometeoroid detectors, and obtain data

about meteoroid entry physics.

KEY WORDS: micrometeoroid, hypervelocity, particle detectors,

impact damage

The existence of meteoroids in space presents a hazard to vehicles and satellites outside of the earth's atmosphere. The

particle sizes vary greatly, with the smallest being about l|am

in diameter. Their velocities vary from about 11 to 72 Ism/s. ^Aerospace Technologist, NASA Langley Research Center,

Hampton, Va

^Aerospace Technologist, NASA Langley Research Center,

Hampton, Va. L-7223 915
The small particles, called micrometeoroids , can damage the sur- faces of space telescopes, external windows on spacecraft,

lenses, protective coatings, and exposed instrumentation. The

damage is the result of "both a high particle flux and long exposure time which causes a gradual deterioration of the surface

efficiency. To establish a damage criteria and degradation

rates, the cumulative effect of micrometeoroid impacts must be

measured.

A Micrometeoroid Impact Simulator (MIMS) was built to measure these phenomena. The major components of the simulator

system are a microparticle charger-injector, a horizontal

^-million-volt Van de Graaff accelerator (VDG), an assemblage

of particle detectors, a particle deflection system, a unique

data acquisition system, and a series of target chambers. The

simulator has accelerated micrometer-size, spherical particles

to velocities in excess of km/s . It can provide particles

at rates in excess of lO/s.

The purpose of this paper is, first, to describe the MIMS

and, second, to outline the current experimental studies.

MIMS THEORY OF OPERATION

The basic principle under which the system functions is

the conservation of energy. It states that a particle with

charge, q, put in a potential field V, will change kinetic

energy by an amount equal to the product of V and q. Mathematically, - mv = Vq

where m is the mass of the particle and v is the velocity

of the particle. Solving the equation for the velocity:

V = V2 Vq/m

The resultant velocity, then, is a function of the potential

field V, the charge on the particle q, and the particle mass

m. From theory and geometry q varies as "V^/d and since the

machine voltage is usually fixed and the charger- injector maintains a relatively constant charging efficiency, the velocity

will vary with particle size; the smallest particles will have

the highest velocities. Figure 1 shows the variation of the

velocity with particle size. With the simulator system at 4-MV, 0.1 |xm diameter particles are accelerated to velocities

about 35 km/s and 10|j,m diameter particles to velocities of

about k km/s

A diagram of the system is shown in Figure 2. The vacuum

environment through which the particles travel is maintained by

the VDG vacuum system and the vacuum systems of the adjoining

916
target chambers. A particle is accelerated as follows: The

particles are initially at rest in a cavity in the chargerinjector. A particle is given a charge and then injected into

the VDG along the accelerating axis. Once in the field of the

VDG, the particle is focused and accelerated through the full

potential of the machine. After it leaves the VDG;, the particle is traveling at a constant velocity. It then traverses

several particle detectors where its charge -induced voltage^

time of flighty and the potential by which it was accelerated

are measured by the data system. If the particle is within

preset time- of- flight limits, it is counted, has its measured

parameters recorded, and is allowed to pass through the deflection system without being deflected. On the other hand, if the

particle does not have the proper velocity it is considered

unacceptable, is not counted, does not have its parameters

recorded, and the deflection system deflects it into a deflection

aperture plate ^-fhere it is stopped before it can enter the target chamber. The data system records the parameters for those

particles which it has allowed to pass into the target area. The particle beam size is controlled with the VDG focusing

system. A focused particle beam would contain about 90 percent

of the particles within a 1-mm-diameter area. A defocused beam,

which is useful in some experiments, can be as large as about

6 mm in diameter.

The particles normally used are carbonyl-iron. Their density is 7.86 X 105 kg/m3 and their diameters range from about

0.1 to 10|j.m. Particles can be injected either by manual operation or by automatic operation at rates of 0.5^ 1^ 2, 5^ or over 10 per second.

PARTICLE CHARGER-INJECTOR

The charger- injector is shown in Figure 3- Details of

operation are available in reference 1. The methods for charging and injecting particles will be described briefly. The

particles are initially at rest in the particle cavity. The

charging electrode-particle cavity assembly and the plate are maintained at a potential of about 12 to 1^ kV. A negative

pulse is then applied to the plate which causes the topmost

layer of particles to achieve erratic motion. This motion is

similar to that of particles in a gas. Some of these particles

will effuse through the cavity exit into the forward charging

chamber. Of the particles which pass into the fonward charging

chamber which are still in motion, some will strike the electrode and gain a very high charge. When one of these highly

charged particles passes through the collimating apertures, it

is focused and accelerated by the VDG. A charging electrode is shown in the figure inset. It is

basically a tapered shaft with a small ball at the end. The

shaft length is about 2.5 mm and the diameter of the ball is

917
about J)0\im, The electrode is usually tungsten. For particle

rates of about 10 per second, the average operating lifetime of

a good charging electrode is about kO to 80 hours. The application of a 12 to 1^ kV potential on the electrode results in a charge transfer to a particle amounting to about 12 to 1^ percent of the theoretical maximum charge that the particle can

hold. If the electrodes are operated at higher potentials, the

charging efficiency does not increase appreciably but the electrode lifetime is shortened considerably.

ACCELERATOR

The h-m Van de Graaff Accelerator (Model KN-4000) is

operated in a positive mode. The high potential is attained by

transferring charge from a power supply located at the base of

the machine (via a specially woven, continuous belt) to the

high voltage dome. Figure 4 is a photograph of the VDG with

the outer tank and the dome removed to show the high voltage

end of the column where the terminal electronics of the machine

are mounted. The primary components of the column are porcelain insulating supports, a series of rings coupled by resistors, the accelerating tube, and a drive motor and alternator

pully which hold the belt. The alternator- also supplies power

for use inside the high-voltage dome. The particle chargerinjector and its electronics can be seen in the end of the

column. The column is cantilevered from a heavy base plate

which is anchored to the floor. In order to operate the accelerator, the tank must enclose the column and be sealed to the

base plate. The tank is pressurized to about 550 kN/m^

(80 psig) with sulfur hexafluoride gas. The SFg is an excellent insulating gas and allows operation at 4 to 4.5 MV with

almost no machine sparking. The accelerator operates stably

if it has been properly prepared for running. At 4 MV the

terminal voltage variation can be held below ±10 kV with some effort.

DETECTORS

The velocity, charge, and position detectors are identical

electrically and differ only in the configuration of the active

elements. Figure 5 shows a diagram of a typical detector. The

detector is a coaxial capacitive pickup consisting of an inner

active element through which the particles pass, a bootstrap

shield, and a vacuum- enclosure support outer tube. The detec- tor active element behaves like a capacitor and senses a passing charged particle. The governing equation is: q = CV

o

918
where q is the particle charge^ C the detector capacitance,

and Vq the induced voltage. A typical detector waveform

showing the voltage Vq as a function of time for a passing

charge particle is shown in Figure 5. The bootstrap shield is

operated by negative feedback from the detector amplifier and

drives the electrical capacitance of the assembly to a lower

effective value. This bootstrap shield action increases the

sensitivity of the detectors. These detectors can reliably

sense charges as small as about 1 fC.

TARGET AEEA

The arrangement of the target chambers is shown in Figure 6. The target chambers are deliberately arranged in a "piggyback"

style. The arrangement facilitates quick experimental changeovers by minimizing the alinement problems. Each chamber is

independent of the others because each has its own vacuum system

and can be closed off from the others. The first chamber, which

is called the single -impact chamber, contains a movable,

remotely controlled, target positioner. The positioner can move the target across the beam horizontally and vertically to

about 5 cm from the center line. The positioner also contains

a gimbal system and can allow pitch and yaw motions of the target to ±45°. The chamber has two 12.5-cm glass ports so that

the target can be seen from the front at angles of ^5°- The second chamber is used for mirror surface tests and

impact flash experiments. It has two chamber sections on a common axis. The first section contains two sets of parallel

plates which are alined along the particle beam axis; one set is

perpendicular to the other. The plates deflect the particle

beam when a modulated high voltage is applied to each set of

plates. In this way, the surface of a mirror can be scanned

with the beam. The impact flash section has four equally spaced^

5-cm front viewing ports and one 5-cm rear viewing port. All of

the 5-cm ports have retainers for mounting filters and photomultiplier tubes.

The third chamber is a specially designed, two- stage,

differentially pumped system used for luminosity and drag experi- ments. The chamber test section, which can be at pressures as high as 300 N/m2_, can be coupled to the other vacuum systems

which are at about 10"'^ N/m2 (lO''^ torr) . The test section is

about 15 cm in diameter and over 1 m in length. It has fifteen

5-cm diameter ports along its axis to view particle luminosity.

DATA SYSTEM

The data system, in addition to recording data, also con- trols the particle beam. The computer- controlled interface

coupled to the multielement detector system selects particles

919
with desired velocities and controls the number of particles

which strike the target. The data system can accept particles

with velocities from less than 1 km/s to velocities greater

than ^0 km/s. It measures the charge- induced voltage of the

particle^ time of flight of the particle , and the potential of

the VDG for each particle within preset velocity limits. From

these three parameters the velocity, mass^ and diameter of a particle can be calculated.

Equipment

The KIMS control room is shown in Figure 7- The three

instrument racks at the left contain the controls and monitoring equipment for the VDG and the particle charger- injector.

The other components are parts of the data system; they are the

computer^ the interface rack, the magnetic tape unit, the

keyboard/printer-punch/reader, and the card reader.

The computer has a l6-bit word length and multiple hard- - ware registers. It contains l6,000 words of core memory to hold the MIMS control program. The control program is a realtime program and uses the interrupt features of the computer

for selection of functions. The operator controls program

functions with the data switches and the interrupt switch on the computer control console.

The magnetic tape unit, the card reader, and the keyboard/

printer-punch/reader are standard computer peripheral devices.

They are used to store data, to modify the program, and for

operator control of the data system.

The interface rack is a specially designed unit. It contains

1. A dual counter scaler unit for counting particles.

2. A time interval counter for measuring particle time of

flight.

^. An analog-to-digital converter (ADC) for measuring

charge- induced voltage and VDG potential.

k. The control unit for the deflection system.

5. Two logic chassis which handle the real-time signals

for the system.

Figure 8 shows the multielement detection system and the

deflection system. Detectors VI, V2, and V3 are particle

velocity sensors. Detector CH-1 is used to sense the particle

charge- induced voltage. Detectors PI and P2 are position detec- tors and are used only to aline the detector system to the

particle beam. The deflection plates, the deflection aperture

plate, and the deflection system power supply can be seen between detectors V3 and P2.

920
Method of Operation

Basically^ the system works as follows (see Fig. 9)- A. particle passes through detector VI and starts the time interval

counter and a 1-MHz clock counter. If the particle gets to detector V2 (^0 cm downstream from detector VI ) it stops the

1 MHz clock counter; the clock counter value is stored in a storage register^ and the clock counter is reversed so that it counts down to zero. Meanwhile the stored count (from the 1-MHz

clock counter) is compared to two preset time- of- flight limits

which fonii a "window." If the particle is within the limits

then it is tentatively accepted; otherwise the system is reset

and waits for another particle. When the 1-Mz clock counter

reaches zerO; the particle should be within the active element

of detector CH-1. The analog-to-digital converter is given a convert pulse which tells it to convert the charge-induced

voltage of the particle. The stored count (time value) is now

loaded hack into the 1-MHz clock counter and the clock counter

is started to count down a second time. If the particle gets

to detector V3^ which is 1 m from VI, then the time interval

counter is stopped and the particle is counted as an acceptable particle. Now the charge- induced voltage and velocity of

the particle are read into the computer. The particle should

reach the entrance to the deflection plates (4o cm downstream

from detector CH-l) when the 1-MHz clock counter reaches zero for the second time. This clock counter triggers a pulse to remove the 10 kV from the deflection plates for a short period

of time to allow the particle to continue in a straight line

path. The particle is then free to enter into the first of the

target chambers. If, for some reason, the particle had been

unacceptable, that is, had either an unacceptable velocity or had struck one of the detectors and failed to get through the

detection system in the proper timing sequence, the 10 -kV potential would not have been removed from the deflection plates and

the unwanted particle would have been deflected into the down- stream deflection aperture plate. During the period in which

a particle was between detector V3 and the deflection plates,

the ADC was given a signal to convert a value from the VDG for

the accelerating potential and subsequently store it in the

computer. When sufficient particles have been accepted, the

data stored in the computer memory are automatically transferred

to magnetic tape for later analysis. Each time a particle is

counted by the system it is compared to a preset maximum allowed

particle count which, when reached, automatically terminates

data system activity.

Data Output

Data can be output during real-time operation to a printer

and stored on magnetic tape. In Figure 10 is sho^ra a printer

921
listing from a typical experimental run. At the top is the

coraputer program initialization which tells the operator ahout

his system options. In this case Mode A has been selected.

This is a real-time mode where the accumulated data will he

automatically transferred to the magnetic tape for storage.

The alphabetic information starting with "PROJECT NUMBER" and

ending with "USE DS, AI^D INTERRUPT TO GO" is typed out auto- matically in sequence. At the end of each line the operator

supplies the n-umeric information. When the last numeric digit

is typed, the project number specimen number, and file reference number are automatically transferred onto the magnetic

tape for identification. After the run has begun, the operator

may print particle data on the printer at any time. If the

operator interrupts the computer, the parameters will be printed

for the next particle the system accepts after the interrupt

was performed. When the operator terminates a run or (automatically) when the maximum allowed particle count limit is

reached, the data system will stop operation, type out "END OF

RUT^," and await further instructions from the operator.

Mode B is like Mode A except that the tape identification

is not generated (PROJECT NUMBER through FILE REFERENCE NUMBER)

and the data are not stored on magnetic tape.

Mode C prints out any data which have been stored on the

magnetic tape for checking or analysis.

Mode D is a test mode which gives a quick and simple system

test of the velocity limits, the ADC conversions, and the timeof- flight measurement.

The data collected on magnetic tape are analyzed on a digital computer at Langley Research Center. Programs exist which

read the data from the tapes and calculate parameters such as particle charge, velocity, mass, radius, momentum, and kinetic

energy. Typical outputs of data are shown in Figure 11. Figure ll(a) shows a data .display for the relative frequency of occurrence (for particle velocity). Figure ll(b) shows the

results of a more extensive analysis and shows the log of

kinetic energy versus relative frequency of occurrence. Listed in Table 1 are some data which illustrate the sys- tem capabilities. Time-of- flight limits were selected in delib- erate groups . The system was required to allow only one particle to be accepted during each run. The data illustrate that

the data system can control the particle velocity by using the

preset velocity limits. Particles in the 10-km/s range are available at rates of several particles per minute. For particles with velocities in excess of 30 km/s, the rate is about

one particle in about 15 minutes when the system is set to shoot

particles of lO/s. The calculated radii and masses for each

particle are listed. The velocities and radii listed show that

the smaller particles are the fastest, as predicted by theory.

922
EXAMPLES OF MICROMETEOROID EXPERIMEfTTS

The experimental programs presently underway are concerned

with damage to sensitive surfaces, micrometeoroid detectors,

and meteoroid entry physics.

Mirror surfaces, such as those to "be used in the reflecting telescopes proposed for the manned space station, will he

struck by micrometeoroids . A study is underway at the MIMS

facility to show how the specular and diffuse reflectance change

when such surfaces are bomharded with microparticles . Coated

aluminum mirror surfaces have been bombarded with up to

109 particles/m2. The narrow angle specular reflectance changes

are being measured.

The simulator is being used to test and evaluate thin- film,

capacitor- type micrometeoroid detectors. These detectors will

be flown on the Meteoroid Technology Satellite. The satellite

has three objectives: to measure meteoroid penetration rates

in b^jmper protected structures (the rates will be compared with

data from previous satellites which had unprotected sensors);

to measure the velocities of meteoroids which penetrate

stainless-steel sheets 13i-im thick; and to measure the impact

flux of small mass meteoroids in near-earth space. The thinfilm capacitor detectors (about 1 or 2[im) are part of the

velocity measuring sensors, and also detect the small-mass

meteoroids. The characteristics of the detector discharge and

the meteoroid parameters which trigger the discharge are not fully understood. Particle characteristics are being corre- lated with discharge signal characteristics to establish specifications for flight hardware. Also, because the velocity

measuring device uses a thin-film capacitor bumper, the effect

of the bumper on possible particle breakup is being studied.

Thin-film capacitor detectors of 0.2 and O.^um thiclmess are also being studied.

When micrometeoroids hit a surface a flash of light is

given off; this could be detrimental to space experiments which

use photosensitive materials. The intensity-history of the

flash and the spectral characteristics are being measured. If

these are known, it may be possible to discriminate against

them. Various kinds of targets, both hard and soft, are being

tested. Preliminary measurements indicate that the flash from

soft targets is somewhat different from the flash from hard

targets. Other experiments suggest that the impact flash

characteristics can be correlated with projectile characteristics, consequently the flash may be used both to detect micrometeoroid impacts and to identify the velocity and mass of the

impacting particle; correlations of this type are being studied.

The MIMS solid-state detector studies are an effort to develop a reliable, long-life micrometeoroid impact detector

for particles mth masses as low as approximately 10"^^ kg.

Presently, no detector exists for these particles. Solid-state

detectors are used by nuclear physicists to analyze particulate

923
radiation. Three types of these solid-state detectors "being

studied for micrometeoroid measurements are:

1. Lithium- drifted silicon detectors

2. Silicon heavy-ion detectors

3. Silicon surface-barrier detectors

To date, two of the lithium- drifted silicon detectors have been

tested with encouraging results. Every microparticle which

struck the first detector generated a signal. The second

detector was bombarded by over 700^000 particles intermittently

during a several-day period and was still giving out signals

when the run was terminated. Two surface-barrier-type detectors

were tested but both of these failed after only a few impacts.

For the luminosity and drag experiments a program is

being developed at the MIMS facility on the laboratory simulation of meteor phenomena and associated atomic collision processes. A differentially pumped, vacuum chamber is being used

to provide a rarefied atmosphere. The particle is shot into

the gas and, depending upon conditions, may cause a streak of

light or be decelerated. Both the light. trail and the particledeceleration will be measured. It is hoped that the results

will supply data about the luminous efficiencies of meteors

and help calibrate meteor photographs and radar observations.

CONCLUDING REMARKS

The Micrometeoroid Impact Simulator is one of the few

apparatus which can accelerate particles of known mass to meteoric speeds (above 11 km/s). With the data acquisition

system now in operation, the velocities of the impacting particles can be controlled and all particle parameters recorded for

future automatic data analysis. The automatic features of the

system permit the effect of thousands of particles on sensitive

surfaces to be measured and analyzed. The simulator is being

used to study the effect of micrometeoroids' impacts upon sensitive surfaces, and to develop new micrometeoroid sensors for

flight experiments

924
VD

iH

1 CQ O W H

VD rH

o\-=i- N-^o o o\Lr\ococo lta

LI \ MJ r> > U \ U \ MJ

i—1

1—

^> U

(Da n. •HQ

-4-VD CVl-^VDVD OVD-^ OJVD I^-C—D-H OJ lf\r-\ a\K\\:~tr\

i>--=j- ^-tn_:t \f\-^ Lr\ OJ OJ o

Velocity (Ws) H ir\ OWD H 0-VD cQ

LTAOJ K-\ir\l^O-4- OJ UAH LfA O OJ OJ H H ITACOVD-^CO-^ HHHHHHHHHOJKA

Particle Time

of

Flight

(^s) I>— HCQ LiAVD KA0\O-^ OJ H Lr\ r-\ -=i- CO if\ a\ ITNH OVDCOVO-=f HCO LP\CO O\C0C0aDC0VD LTAVDVD hf^OJ

Time

of

Flight

Limit

([j,s)

Slow

Fast

OOOOOLALTNLrAUAOO oooooi>-i>-i:— i>--=i- H H H H H

Particle No. H OJ KA-^ ltnvd o-co a\ o h

H H

925
100

Vel. km/s

10

—' !

-i \—

i —J —I +

— Dim i !— K u n 1 — O Run 1

—X—

-—1 •

1 —

4- •\ —

i-:;t"^

J^Thec

_4_ - retical ci Charging

! I

jrve for 4

efficiency

4--\. . MV potentia

- 12. 5%

i

-

1

1 U -—k

\ }

1 i —

—+ —'~ _;;_^J__ 1 -L1^

Sii-rTfftittt;ttttr iiTr^iSiulirh;

I ."I 1 ^

i

j

[in;i;i;;TntiTit-

+——-

1——h- - ITTTtittiitiil

i

!

1 'r~

1 , 1

1 ~ ! L..-. :.

"

1

1 4.

i

4; J. - — -I- .1 1-0 10 Diameter, [im 'igure 1. -Variation of microparticle velocity with diameter for

4-million-volt potential.

926
CHARGER

INJECTOR

VELOCITY AND

CHARGE

DETECTORS

\

DEFLECTION

PU\TES

TARGET CHAMBERS

\ MIRROR

SURFACE

VAN de GRAAFF ACCELERATOR

SINGLE IMPACT ^fTaSH^ LUMINOSITY

INTERFACE COMPUTER

TAPE

UNIT KEYBOARD/

PRINTER

DATA ACQUISITION SYSTEM PERFORMANCE CHARACTERISTICS CARBONYL IRON PARTICLES; DENSITY 7. 86 >c 10^ kg/m^ PARTICLE DIAMETER RANGE; 1 to 10 um

PARTICLE RATES TO GREATER THAN KVs

PARTICLE VELOCITY RANGE; 1 > v > 35 km/s MEASURES AND RECORDS PARTICLE PARAMETERS Figure 2. —Micrometeoroid Impact Simulator System.

927
Charging

electrode

Van de Graaff accelerator tube

Charging electrode

Figure 5, -Mlcroparticle charger-injector.

928
To amplifier
Figure 7« — MIMS control room.

930
Particle

travel

Storage

register

1^, 1 MHz clock up-count

1 IVlHz clock down-count

40 cm •40 cm I I

V-2

Time interval counter

1 MHz clock down-count

20 cm- 20 cm— El

^

CH-I V-3 Deflection

plates

Particle count Figure ^. —Data system schematic

Target area

• Deflection aperture

plate

931
DATA SWITCH MODE

A REAL-TIME

B BY- PASS TAPE

C PRINT (nJT OF TAPE

0 SYSTEM TEST

SET MODE DS, INTERRUPT

INITIALIZATION

PROJECT NUVBER 333

TEST NUMBER 003 SPECIMEN NUMBER 023

FILE REFERENCE NUMBER 000001

MAX. PARTICLES ALLOWED 0015000

MAX. ENTRIES ALLOWED 0015000

MINIMUM Fll TIME 0005 MAXIWJM FLT TIh€ 0250

READY FOR REAL-TIME.

USE DS, AND INTERRUPT TO GO.

IDENTIFICATION

CHRG FLT TIME VOLTMETER

00025 017075 00389

PARTICLE COUNT 0000023 CHRG FLT TIME VOLTMETER

00019 01 6082 003 8 7 PARTICLE COUNT 0000072 CHRG FLT TIME VOLTMETER

00012 0^5^ikk 00390

PARTICLE COUNT 0000117 CHRG FIT TIME VOLTMETER

0001 4 01500 4 00390

PARTICLE COUNT 0000180 CHRG FLT TIME VOLTMETER OOOn 015215 00388

PARTICLE COUNT 0000241 CHRG aT TIME VCLTMETER

00017 015108 00391 PARTICLE COUNT 0000301 CHRG FLT TII€ VOLTMETER

00043 017982 00386

PARTICLE COUNT 0000375

DATA FROM PRINTER

END OF RUN Fiffure 10. Data from printer list of experimental run,

932
25

20-

.40

.35

.30

Relative frequency

of occurance

.25

.20

.15

.05

0

10

.05

0 5.0 10.0 15.0 -.9 .8 -7 -6 -5 Velocity Log of kinetic energy

(km/s) 2, 2, (Kg-m Is ) (a) Velocity distribution. (b) Kinetic energy distribution.

Figure 11. —Typical results from Data Acquisition System.

933

Paper No. 5^

A SPACE RADIATION ENVIRONMENT SIMULATOR FOR THE

EVALUATION OF SOLAR CELLS

P. A. Newman, J. J. Hirshfield, H. E. Waimemacher , NASA, Goddard

Space Flight Center, Greenbelt, Md., and M. Eck, Howard Slack Associates, Baltimore, Md.

ABSTRACT: A facility for studying the properties of solar cells and

other optically sensitive devices under a wide range of environmental

conditions such as could be found on extra-terrestrial space missions

has been developed at the NASA Goddard Space Flight Center. The

system includes control of temperature, vacuum and solar irradiation

while irradiating a 23 cm by 23 cm array of samples with 4 MeV protons or electrons. Typically the samples can be controlled in temperature between -170°C and 150°C, while a shroud, used to control

radiative coupling, can be independently controlled between -196°C and

150°C. The vacuum system, which is virtually free of hydrocarbon

contamination, will evacuate the chamber to less than 1 x 10"^ torr in less than 4 hours. A 400 l/s Noble Vac-Ion pump backed by a titanium

bulk sublimator and a titanium filament evaporator has been designed

to accommodate the large gas loads possible during irradiation. A

4.2 KW compact Xenon arc solar simulator provides a close match to the solar spectrum from low irradiances to an irradiance of four solar

constants. An automatic data acquisition system modified to generate

digital solar cell I-V characteristics is used to collect data in a form

suitable for computer processing.

KEY WORDS: space simulation, space radiation, solar cells, tests, thermal control, high vacuum.

935
INTRODUCTION

A facility for the in-situ testing of solar cells and other optically

sensitive devices under a simulated space environment has been built and is in use at the Goddard Space Flight Center of NASA. The system

is capable of irradiating a 23 cm x 23 cm array of solar cells with both

particulate radiation (protons or electrons) and simulated solar radiation while being maintained under a vacuum of 10"^ torr or less and

controlled in temperature over a range of -170°C to 150°C. Data acquisition is accomplished automatically using a conventional system

with a number of special modifications.

CHAMBER DESIGN

A plan view of the system is shown in Figure 1. The main chamber is a stainless steel (304) cylinder 45 cm in diameter and 203 cm

long with a Wheeler flange holding the sample holder and all of the

electrical feedthroughs at one end and an entrance port for the particulate radiation at the other. A port with a u.v. grade quartz window

is located at right angles to the particulate beam adjacent to the beam

entrance port. The use of a slightly off-axis turning mirror permits

the simultaneous irradiation of the samples by photons and protons or

electrons. The length of the chamber is needed to accommodate the divergence of the particulate beam, A cryogenic shroud that covers

the region from the sample holder to the turning mirror is used to prevent unwanted radiative coupling. The sample holder can be rotated about 5° on an axis normal to the plane of the particulate beam

and the simulator beam so that measurements can be made with each

of the components normal.

VACUUM SYSTEM

All of the high vacuimi pumping equipment is located below the main chamber (Fig. 2). The various pumps used are listed in Table 1 along with their pumping speed and range of operation. The basic

pump is a 400 l/s Noble Vac-Ion^ pump that is assisted in its various

ranges by a titanium filament evaporator and a titanium bulk sublimator. The roughing system is moimted on a separate cart and is con- nected to the system by a flexible line. It consists of a gas aspirator

and three cryogenic sorption pumps. Under ideal conditions (system

^Varian Assoc. Palo Alto, Calif.

936
not opened to atmosphere for too long) the chamber can be evacuated

from atmosphere to 1 x 10"^ torr in less (Fig. 3) than four hours.

With the exception of one viton "O" ring used to seal the quartz window all openings are closed with copper gaskets to provide a hydrocarbon free system. The system pressure is monitored by a nude ion

gauge located in the main chamber. The introduction of high energy

protons into the system during irradiation of the solar cells could provide for a substantial gas load both from induced outgassing of the chamber walls and substrate and from the protons themselves. Fortunately the titanium bulk sublimator has an enormous pumping speed

in this region so that the pressure can be kept under control for these

peak gas loads.

SAMPLE PLATEN

Solar cells to be irradiated are mounted on a temperature controlled copper platen (Fig. 4) located at the end of the chamber. The

platen size is 23 cm x 23 cm and can accommodate about 50 solar cells

at one time. A series of 9 Faraday cups are mounted at various points on the platen so that their apertures are in the plane of the solar cells. The walls of the Faraday cups are fabricated from tungsten so that the wall thickness could be made greater than the range of 4 MeV electrons. The Faraday cups are used to monitor the uniformity of the radiation over the surface area. Thermocouples are mounted on some

of the cells to record the actual cell temperature during the experi- ment.The wires used to transmit the cell characteristics are run along

the platen surface and are shielded by copper guides. The thermal

coupling of the solar cells can be monitored under various experimental conditions by measuring the open circuit voltage. It has been found

that the cell junction temperature can be kept within 3 or 4°C of the sample platen with 1400 W/m^ (approximately one solar constant) inci- dent on the solar cell. This is about what one would expect based on

the published values of the thermal conductivity of silicon,

TEMPERATURE CONTROL SYSTEM

The thermal system is used to control the temperature of the

cryoshroud and the sample platen. Several critical design constraints

had to be met. These were:

1. Rapid shroud temperature response over a range of -170°C to 150°C.

937
2. A shroud gradient of less than ±5°C over this temperature

range under a distributed load of 1400 W/m^,

3. Accurate shroud temperature control over the desired temperature range.

4. Rapid specimen temperature response.

5. Accurate specimen temperature control {±2°C) while being

subjected to varying incident solar radiation intensities.

In addition both the shroud and the sample platen had to have an ac- cumulated leak rate that was less thanl x 10 ~^ ml/s referred to standard pressure and temperature over the operational temperature range. The thermal conditioner for the shroud consists of a pumped

gaseous nitrogen loop, a series of heaters, and a gaseous nitrogen to liquid nitrogen shell and tube heat exchanger. The entire system is housed in an equipment cabinet 76 cm wide, 137 cm long and 183 cm

high. A block diagram of the system is shown in Figure 5. A three stage centrifugal blower is used to circulate approximately 104 1/ s (220 CFM) of nitrogen gas through the system. Modulating butterfly valves proportion the gas flow through the heat exchanger. This variation in flow combined with a variation in power

input to the heater produces a very precise means of effecting temperature control. There is also a large potential heat source or heat

sink reserve. The response time and temperature overshoot of the system can be minimized by biasing the system source and sink

against each other. Rapid transients require that system mass be minimized. This is especially critical in the heater design since this element runs at the highest temperatures of any system component. The system heater

shells are fabricated from thin wall stainless steel tubing 7.6 cm in diameter. Each heater module contains a 1.6 KW quartz infra red

lamp and a set of thin aluminum fins. The fins are radiatively heated

by the infra red lamp and are convectively cooled by the gaseous nitrogen stream. The weight of the heater extended surface is approximately 450 g. The system heat sink is a liquid nitrogen boiler. Gaseous nitrogen is pumped through the tube side of the shell and a tube heat ex- changer. The shell side of the unit is flooded with liquid nitrogen.

The liquid level is maintained by the use of a temperature probe, controller, and a solenoid valve to permit filling from an external reservoir. Boil-off from the liquid nitrogen is mixed with the exit gas flow

stream to further lower the circulating gas stream temperature. A

pressure relief valve is located in the main gas stream to prevent ex- cessive pressure buildup either from the introduction of boil off gas

or from gas expansion during loop temperature up-transients. The

relief valve maintains the gaseous nitrogen loop pressure at 21 KN/m^

(3psig).

938
The shroud thermal system tests showed that the system met the design goals. The transient performance of the system is shown in Figure 6. To meet the temperature control requirements of the sample

platen stated previously requires the use of a secondary cooling medium to eliminate the temperature rise that would otherwise result from the incident radiation.

It was desired to have a work surface temperature change capability of -170°C to 150°C. Conventional mechanical refrigeration temperature control techniques were examined and discarded as being too massive and slow in response.

The sample platen is shown in Fijnire 4. The surface that the

samples are mounted on is made of 1.3 cm thick copper plate to obtain

uniform temperature distribution. Heat removal from the platen is accomplished by conduction to a liquid nitrogen boiler and heat sink. Heat addition is provided by the use of four (4) 300 W infrared quartz

lamps. The conduction path between the liquid nitrogen boiler and the platen is a square box section fabricated from stainless steel. The

dimensions of this section are carefully selected and controlled. The

infra red lamps are mounted inside copper "U" sections which are welded to the conduction path box section. The resultant thermal net- work provides the rapid response and ability to absorb varying

amounts of incident flux without changing the surface temperature outside the design limits. Platen temperature control is provided by

sensing the surface temperature, comparing it to a setpoint temperature, and varying the power input to the infrared lamps to effect a

balance. A schematic representation of the system conduction paths

and their electrical analogues is shown in Figure 7. The results of the system transient tests are shown in Figure 8.

RADIATION SOURCE AND BEAM OPTICS

The radiation source initially employed was a horizontally

mounted 4 MeV Dynamitron^ that was located in a shielded room adja- cent to the environmental chamber. The Dynamitron, a d.c. charged

particle accelerator, whose equivalent circuit is shown in Figure 9, is basically a set of cascaded vacuimi tube rectifiers capacitively coupled

I to an r.f. power supply operating at 130 KHz. Two large r.f. elec- trodes, located along with the vacuum tubes and the ion source in a

i large pressure vessel filled with 1800 Kg of sulphur hexaflouride as an

! insulating gas (552 KN/m^) (80 psi), induce an r.f. potential in a set of

Radiation Dynamics, Westbury, N.Y.

939
corona rings causing a d.c. flow through the rectifiers and establishing a large potential at the output in the high voltage terminal. The ca- pacitance from the r.f. electrodes to the pressure vessel and the previously mentioned inductive load result in resonant circuit operation.

An electron gun or duoplasmatron t5^e ion source with an einzel

lens, located in the high voltage terminal, provides an electron beam.

The duoplasmatron and einzel lens is also used for the production of

ion beams.

The beam is accelerated thru an evacuated accelerator tube across which the full d.c. terminal voltage appears. The mult-section

glass and electrode acceleration tube assembly with an external resistance bleeder network provides a voltage gradient across the tube

and establishes beam focussing by the tube electrodes. Typical base vacuum in the accelerator tube and its drift tube extensions is between

5 X 10"^ and 5 x 10"^ torr. Operational pressure is on the order of

10~^ torr. After acceleration from the terminal potential to groimd

the beam leaves the grounded pressure vessel and is transported to the target through evacuated stainless steel drift tubes with the aid of electrostatic and electromagnetic lenses and deflection elements as well as beam defining apertures.

The beam, after entering the beam transport system, is passed

through a horizontal 46 cm circular pole face switching magnet. The

magnet has a focussing effect in the horizontal plane but its primary

function is to deflect the beam through a selected angle for transport

to the target area in the adjacent room. In positive ion operation the magnet acts as a momentimi analyzer insuring the purity of ion species

and mono-ergic properties of beams reaching the target.

After deflection of the beam into a drift tube aligned with the center line of the environmental chamber, beam focussing and/or defocussing is accomplished with electromagnetic lenses. The beam

focussing properties in the horizontal plane of the switching and

analyzing magnet are also utilized. Electrostatic fields are used to provide slight beam deflections for alignment of the beam within the drift tube. All of these beam

optic elements are designed to introduce a minimum of beam aberration or energy spread.

A large cross-section electron beam can be produced at the target

plane by placing a scattering foil in the path of a well focussed beam.

This technique has limitations, particularly for the evaluations described here, but has been used. An air core solenoid lens is also

used with the electron beam to produce a beam convergence that will allow entry into the environmental chamber with subsequent cross- over and divergence.

The solenoid lens is not a practical method of handling proton

or ion beams. These beams are focussed and defocussed with

940
electro-magnetic quadropole lenses. It should be noted that the

accelerator without supplemental focussing elements provides a low

emittance beam with 90% of its charged particles located within a

1 cm diameter circle. A quadropole magnet is used to supplement the focussing of the beam by the analyzing magnet. Pairs of quadropoles or quadropole doublets are used to introduce the required beam con- vergence for entry into the chamber. Quadropole lenses produce con- vergence in one plane and divergence in a 90° displaced plane. Doublet

lenses with poles displaced 90^ will thus permit control of circular

beams and the introduction of convergence or divergence as desired.

SOLAR SIMULATOR^

The solar simulator (Fig. 10) uses a high pressure, compact arc, Xenon lamp as the energy source. It can be operated as high as

4.2 ftw for high intensity simulation. The short term intensity is regulated by a feedback amplifier that uses the output of a solar cell

in its feedback loop and controls the firing angle of a silicon controlled rectifier power supply. The long term intensity is maintained

by a calibrated solar cell mounted at the same distance from the simulator as the cells being studied but outside the chamber. The

solar beam is diverted by an external mirror and the lamp intensity

is periodically adjusted to maintain the calibrated short circuit current. The external standard is maintained at a temperature close to ambient by a separate control system. A Lenticular lens system is used to achieve a imiform irradiance at the sample plane that is better than 2% uniform over a 23 cm by 23 cm area.

DATA ACQUISITION

It is desired to measure the current-voltage characteristics of the solar cells in their active region after each dose of radiation.

Various experiments have been designed to study the effects of temperature and solar intensity during irradiation to simulate certain

extra-terrestrial missions. The data system has been designed to record the data in digital form so that computer processing can be

utilized. To achieve this the circuit shown in block form in Figure 11

is used in conjunction with a data acquisition system. The cell characteristic is traversed by driving current from a separate supply and

measuring the voltage across the cell and the current by measuring

Spectrolab Inc., Sylmer, Calif.

941
the voltage drop across a precision resistor. The current is driven in stepwise fashion by an I.C. operational amplifier used as a staircase

generator triggered by the data acquisition system; an I.C. voltage

comparator; an I.C. integrator; an I.C. gated feedback amplifier and a power transistor. The feedback amplifier is used to modify the volt- age step of the staircase generator as a function of the change in the

cell current so that the data points on the I-V curve are imiformly

spaced. The gating of the feedback amplifier is required to suppress

the transient caused by the cell being switched by the cross-bar scan- ner in the data acquisition system.

In the data acquisition system each cell is assigned to a data

channel. A tape controlled programmer selects the channel, the range, the integration period and the measurement mode. The current and voltage are measured simultaneously on separate digital volt- meters and are recorded on punch paper tape sequentially. The number of data points collected for each cell can be varied to suit the particular experiment. While the data acquisition system provides

precision of better than ±.5% the actual precision is limited at present

to ±2% due to short term fluctuations on the intensity of the simulator

believed to be due to the wandering of the arc. Because of long lines necessary to connect the data acquisition system and the samples in the chamber a Kelvin connection is used to eliminate the problem of the voltage drop in the 1800 cm of wire.

A typical data run would sample and record the Faraday cup currents, the thermocouple voltages, the system pressure and the time

during the particulate irradiation and the cell I-V characteristics and

temperature after each irradiation. Thus from the data tape it is possible to computer generate plots of 1^^, V^^, and Pmax vs. fluence for each cell with cell temperature and light intensity as parameters.

CONCLUSION

The Space Radiation Environment Simulator can be used to create

in the laboratory a wide range of conditions that could be encountered

on space missions and that could degrade solar cells, solar cell cover

glasses and solar cell cover glass adhesives. Properties of these devices and materials can be studied as a function of particle fluence,

simulated solar intensity and temperature over a wide range. Raw

data output in digital form can be used with a computer to provide any

significant type of analysis as required.

942
ACKNOWLEDGEMENT

The authors would like to express their gratitude to a number of people who gave valuable inputs at various points in the design of this chamber. We owe special thanks to W. Gdula and M. Moore formerly

of this center, W. Wappaus, GSFC, A. Lauer of Varian and H. Slack of Howard Slack Assoc.

TABLE 1

Pump Max. Pumping Speed Pressure Range

Noble-Vac- Ion. 4 x lo^ 1/s @ 4 x 10"^ Torr below 10"^ Torr

Titanium

Filament

Evaporator. . 1 x 10^ 1/s @ 1 x lO''^ Torr* below lO'^ Torr

Titaniimi Bulk

Sublimator . . 7 x lO"^ 1/s @ 1 x lO'^ Torr* below 5 x 10^^ Torr

*Area/Conductance limited at lower pressures.

943
203 CM

SAMPLE PLATEN

INSTRUMENTATION ]ljn FEEDTHROUGHS ^if"

CALIBRATION CELL

BULK Tl SUBLIMATOR CONTROL

THERMAL SHROUD

ROUGHING CART

THERMAL CONTRO n tn GAsouTn n ,

, ' *GASRET ^XJ I .-!d1. OUir

DRIFT TUBE QUARTZ FOR PARTICULATE WINDOW RADIATION MOVABLE MIRROR

FOR CALIBRATION

SOLAR SIMULATOR

iLOWER LN , BOILER HEATER

Fig. 1— Space radiation environment simulator — plan view.

QUARTZ WINDOW

400 US NOBLE VAC ION

BEAM PORT

ION GAUGE

TO ROUGHING CART

Tl BULK SUBLIMATOR

Fig. 2—Space radiation environment simulator — vacuum pumping

components

944
Fig. 4—Typical pump-down cycle. 945
+20 -

0 5 10 15 20 25 30 35 40 45 50 55 0 10 20 30 40 50 60 70 80 TIME - MIN. ; TIME - MIN. Fig. 6—Shroud transient performance.

V

Ml Fig. 7—Thermal mock-up of

platen.

V

Fig. 8—Platen transient performance,

946

r
Paper No, 59

PROPERTIES OF OPTICAL MATERIALS IN SPACE

Fred W. Paul"""

REFERENCE: Paul, Fred W., "Properties of Optical

Materials in Space," ASTM/IES/AIAA Space Simulation

Conference, 14-16 September 1970.

ABSTRACT: A program has been established to provide

a central clearinghouse for information about the

effects of space environment on the optical and

other physical properties of optical materials.

Information is being collected and organized on the

effects of space environment on refractive index,

dispersion, transmittance , reflectance, thermal

expansion coefficient. Young's modulus, birefringence, and yiel.d and breaking points of all materials

of interest for the transmission or reflection of

optical radiation in space.

In technical areas where information about the

behavior of materials does not exist a program of

laboratory investigations will provide the required

information. In preparation for some in-house

studies of materials at very low temperatures and

also during exposure to a charged particle flux an experiment has been devised and instrumented to

provide measurement of dimensional changes and

refractive index changes at several wavelengths.

The essence of the method is the use of Fabry-Perot

interferometry to measure independently the dimensions of the sample of material and the optical path

length through the material.

^Test and Evaluation Division, Goddard Space Flight

Center, Greenbelt, Md . 20771

949
KEY WORDS: optical materials, physical property

changes, space environment, information center,

interferometric measurements

Introduction

The behavior of optical materials and components and systems in space is of primary interest to

the designers and users of space borne optical

devices. Questions often arise, at all phases of a program, i.e., design, manufacture and use, as to

what effect the environment will have on a particular element. During the years , of space studies a substantial body of knowledge about this subject has

been developed. It has often been difficult to

locate the required information to answer a question

about what behavior can be expected. The program

which is being described in this paper is an effort

to collate and codify this knowledge to make it more readily available to those who need it.

The environment of space is not a single

definable thing. Depending on where in space one intends to operate there may be a variety of atomic

ions, or a concentration of electrons, or trapped

radiation belts, or solar wind and a variety of

electromagnetic radiation, and low or high temperature. There are also environmental components which

are carried with the spacecraft, for example con- densible vapors or a nuclear power plant. It may

or may not be desirable to include the environments

of manufacturing, test, prelaunch and launch in this

discussion. The material properties that are to be

included are transmittance, absorptance, emissivity,

reflectance, scattering, refractive index, dispersion, birefringence, themal expansion coefficient.

Also Young's modulus, hardness, yield point and

breaking point for all materials of interest for the

transmission and reflection of radiation in space.

If it is decided that photovoltaic devices and

various detectors should be included, then the

electric properties of these devices and materials

must also be included. The complexity of this

950
subject, which is illustrated by the above remarks,

is one of the reasons that answers to questions

about the behavior of materials in space are difficult to find. Another factor is that the determination of what material will serve for a specific

mission is often a routine engineering task which is

forgotten as soon as it is accom.plished . However,

there are some notable exceptions to this, e.g., the

vigorous and effective program on spacecraft coatings by the Thermophysics Specialist Group of AIAA

and ASTM has established a solid and well publicized

basis for choice of materials and methods of evaluating new materials.

Information Center

A program has been established at GSFC to

collect and interpret and interrelate information

about optical materials suitable for use in space

borne systems. The optical properties like transmittance and refractive index and the other physical

properties like moduli and expansion coefficients

will be recorded and the effect on these properties

of exposure to the various factors of the use environment will be determined. In instances where

available publications such as periodic literature,

meeting reports, and conference proceedings do not

provide adequate information studies will be established to fill the gaps. This program is expected

to produce extensive and adequately cross referenced

bibliographies. In addition it is planned that

several summaries or position papers will be issued

which will provide comparative information on candidate materials for specific applications. It

is intended that the bulk of the literature research,

extraction and compiling of pertinent information

and the studies to provide new information will be

done under contract, probably with research institutes or academic departments. In this period of

shrinking budgets this program has not competed

successfully for funds with some of the more glamorous items like manned flight and orbiting observatories. Nonetheless a small start has been made as an in-house task and some interesting items about

951
materials behavior in space or in simulated space

environment have been noted.

There is a very extensive and well-known literature about radiation effects on optical materials

(for example. Reference 1) which has evolved from

nuclear studies and from solid state physics. It is

usually time-consuming to extract from this large

volume of literature information pertinent to a particular space application problem. Bibliographies of materials directly related to space applications (Reference 2) are very helpful and more of

them are needed. There are also some very useful

general articles (References 3, 4) which elucidate

the problem areas and the practically possible

solutions. In addition there are a number of

studies of more restricted scope which confine their

attention to a certain class of materials or a single environmental factor, (References 5,6,7,8) or similar restriction. Information in this form is

the easiest to use, provided that the circumstances

involved in the engineering application fall within

the scope of the study. There are also two handbooks (References 9, 10) which provide guidance for

the use of optical materials in space.

It would not be appropriate at this time when

but a small fraction of the available and desirable

information has been collected, to draw conclusions

or to summarize. However there are one or two items

which might be mentioned briefly. One is that

reflecting surfaces appear to endure the space

environment very well. Canfield, Hass and Waylonis

(Reference 11) showed that aluminum mirrors over- coated with magnesium fluoride will stand years of

space exposure without loss of reflectance. Pittman,

et al (Reference 7) and Gunter, et al (Reference 12)

have observed that diffraction gratings, both original rulings and replicas, show only very small

adverse effects of exposure to a simulated space

environment. The chief hazard to reflecting sur- faces seems to be contamination by deposits of con- densible materials, Gunter also found that some mirror substrate materials are better than others in

resisting change of optical figure at high doses of

electrons. Among fused quartz, pyrex, borosilicate

crown glass, synthetic fused silica, and Cervit, the

952
synthetic fused silica retained its shape best and

also showed least discoloration due to the irradiation. Another item of interest is the study of

changes of refractive index of optical glass due to

gamma rays (Reference 13) and electrons (Reference

14) . Malitson, et al observed substantial changes

in refractive index, as much as a few parts in the

fourth decimal place due to irradiations . The

refractive index showed continuing instability for

long periods after the irradiation. Unfortunately

this study was left incomplete due to lack of funds

A third item that might be mentioned is the conclusion reached by Heath and Sacher (Reference 6) that

magnesium fluoride, barium fluoride and synthetic

sapphire will provide the most stable high transmittance for ultraviolet wavelengths. These are but

a few of the interesting bits of information about

the behavior of optical materials in space. They

serve to illustrate that many useful and sometimes

unexpected results have appeared in the literature,

perhaps the above few paragraphs may serve to introduce the literature search and organization of

information which is planned. I would like now to

describe a modest in-house laboratory program which

is being undertaken to provide information in an area which seems not to be well covered elsewhere.

An Experiment

13 14 Malitson, Dodge and Gonshery have reported ' on the effects of penetrating radiation on the

refractive index of some optical glasses. Gunter,

et al^^^ , have observed changes in optical figure of

diffraction gratings and their substrate material

due to doses of 1 MEV electrons. It seems that it

would be interesting and useful to know at what rate

these changes occur during exposure to the radiation

environment and whether the changes in refractive

index and figure exhibit any rapid increases or decreases when the radiation field is removed. In

order to provide the required information a system

has been built which permits continuous measurement

of refractive index and physical dimensions and

optical figure during exposure of materials to

953
penetrating radiation. The essence of the method is use of interferometry to measure index of refraction

and dimensions during exposure . This is done without mechanical contact of the measuring system with

the sample to be measured.

The method being used is illustrated in Figure

1. Plates A and B are arranged as a Fabry- Perot

interferometer. The sample is inserted part-way

between the F-P plates and is held parallel to them.

By patch-work coating of the sample four different

interference patterns are obtained. The first

pattern, formed by the rays indexed (1) , is used to

determine the interferometer spacing, d, by conven- tional Fabry- Perot procedures'^^) . Rays (2) are used

to determine the optical path length between the

interferometer plates when the rays pass through the

sample. This length is d + (n - l)t where n is the

refractive index and t is the thickness of the

sample. Rays (3) and (4) are used to determine

respectively the distance between the upper interferometer plate and the top surface of the sample,

and between the lower interferometer plate and the

bottom of the sample. The difference between the

distances found from rays (3) and (4) and (1) , when

appropriate corrections are made for phase changes

on reflection and for thickness of the coatings,

gives the thickness of the sample, t.

To consider the required measurements in a little more detail the following equations may be

written

P^X = 2d + 26^ (1)

P^X = 2[d + (n - l)t] + 26^ (2)

: P3X = + 6^ + (63 + £3) (3)

V = + 6^ + (64 + 6^) (4)

Where P-^* ^2 ' ^3' ^4 experimentally found

orders of interference for the four sets of rays.

X is wavelength, d is the F-P interferometer

spacing. 6-^ is the phase change on reflection at

954
the F-P plates, assumed to be the same for both

plates. n is refractive index and t is the thickness of the sample. L -^^ and Z 2 the distances

between the Fabry-Perot plates and the nearest sur- face of the sample. 63 and are the phase change

on reflection and the thickness of the patch coating

in the upper side of the sample and 6^ and are the corresponding quantities on the lower side

If equation (1) is subtracted from (2) one has

- P^\ = 2(n - 1) t (5)

The phase change drops out and the left hand side

contains only measurable quantities. If we subtract the sum of (3) and (4) from (1) we have

P^\ - (P^X + P4X) -

(6)

2d - 2{L^ +L^) - (6^ + + 6^ + e^)

Since

d - (j: + L ) = t, we have

P^X - (P^X + P^X) = 2t - (6^ + 4- 5^ + s^) (7)

Once more the left side contains measurable quantities. However, to evaluate the 6's and e's on the

right side a separate experiment is required. For

this fringes of equal chromatic order will be used

as described by Bennett "'"^^

. In this method the

coated area for which 6 and e are to be determined

will be partially overcoated with an opaque layer of

aluminum. The equal chromatic order fringes which

are formed by interference between reflected light

from the double coated surface and a uniformly

coated flat comparison surface can be analyzed to

determine values for both 6 and e of the coated area

for which these quantities are to be measured.

Inserting all the measured values in equation (7)

yields a value for t. Putting this value of t in

equation (5) yields a value of n. This process is

repeated for as many wavelengths as necessary.

The principal application envisioned for this

955
method of measurement is in situ measurement in

intense penetrating radiation fields. For several

reasons, however, the first application will be

measurements at reduced temperatures. For one thing,

a new method always required "shaking down" and the

available radiation facilities are so heavily loaded

with work that to take up facility time for shakedown operations is undesirable . Another factor is

that information about the performance of some ultraviolet transmitting materials at low temperatures has not been published and such information

has been requested for planning purposes on some space missions. Also the measurement of refractive

index and thermal expansion coefficient of some twenty optical glasses by Molby-^^^ and the thermal

expansion of some infrared transmitting materials by

Browder, et al ^ provide a convenient means of com- paring results obtained with the new method with

those obtained by older methods.

For the low temperature work a liquid helium

cryostat is used which is provided with both electrical heaters and pumping tubes so that temperature

can be stabilized at points both below and above the

boiling point of helium at atmospheric pressure.

The sample is attached to the liquid helium chamber

by a massive OFHC copper block to which the sample

is attached with indium solder to provide good

thermal contact . The Fabry- Perot eta Ion, which

consists of fused silica plates separated by an invar spacer ring, is held in position by thin

stainless steel strips with adjusting screws so that

the etalon plates can be maintained parallel to the

surfaces of the sample which is positioned between

the two plates, as shown in Figure 1. The cryostat

has two fused silica windows through which observations of the fringe pattern are made.

The cryostat is evacuated by a turbomolecular

pump with a large liquid nitrogen cooled trap

between pump and cryostat followed by two room temperature right angle bends and finally a right angle

bend at the liquid nitrogen shield wall of the cryostat. Pumpout and backfill operations are carefully

managed to reduce the likelihood of contamination of

the cryostat. Monitor mirrors are frequently

inserted in place of the interferometer and are

956
measured for contamination effects by the method

recommended by Hass and Hunter-^^^ . No contamination

has been detected in three months of operation.

The interferometer is illuminated by an argon

ion laser, which provides five strong lines in the

range 476 nm to 514 nm. On occasion when longer

wavelength information is desired a helium-neon

laser radiating at X633 nm is substituted. The

fringe patterns are photographed either with a view camera or with a Hilger constant deviation spectrometer with camera attachment . To avoid confusion

among the several interference patterns that exist

due to the various rays shown in Figure 1, an aperture is used which limits the area of the interferometer which is recorded by the camera to dimensions

of the order of 1 mm x 3 mm for any particular

photograph. The chosen area is moved sequentially

to obtain pictures of the fringe pattern forined by

rays (1), then rays (2), etc., as shown in Figure 1.

A set of four pictures, which can be taken in an elapsed time of 5 to 10 minutes, gives the necessary

information for determining the refractive index at

several wavelengths and the mechanical thickness of

the sample for a specific temperature. The temperature of the specimen is measured with a thermocouple

embedded in the specimen but outside the portion

used for optical measurements. A series of measure- ments with several thermocouples simultaneously has

shown that the temperature measured at the thermocouple is not significantly different from that

where the optical measurements are made.

The apparatus which has been described has been

completely assembled. Preliminary experiments have

been run on temperature sensing, on detemination of

phase changes on reflection, on contamination and on photography of the Fabry- Perot fringes. The first

complete experiments to determine refractive index

and linear themal expansion coefficient at temperatures between room temperature and liquid temperature will be done this month. Further work will

include additional materials in a temperature range

of interest to space missions followed by a transfer

to a radiation facility.

957
4- [—\

k-INTERFEROMETER

d 1 1

t (3

1. 1 ^1 —^ 11 n.rnt -^SAMPLE

1 H

r^^

I i=

11 11

1 ^ INTERFEROMETER PLATE ©: : ® (D ©

Figure 1. Schematic of Fabry-Perot Interferometer with Sample of Material to be

Measured Inserted

References

(1) "Radiation Effects on Glass - An Annotated

Bibliography," Sept. 1965, Radiation Effects

Information Center, Battelle Memorial Institute.

(2) Avery, L., "A Bibliography on the Effects of

Short Wave and Particle Radiation on Glasses,

Astrogeophysical Memo #164, July 1964, High

Altitude Observatory, Boulder, Colo.

(3) Becker, R. A., Applied Optics, Vol.6, p. 955,

(1967)

(4) Barnes, W. P. Jr., Applied Optics, Vol. 5,

p. 701 (1966)

(5) Neu, J. T., Philbin, E. J., Mahadwan, P.,

Compton, J., "Evaluation of the Effects of

Space Environment Exposures on Index of

958
Refraction and Extinction Coefficients of

Apollo Window Materials," NASA CR 1019, April

1968.

(6) Heath, D. F. and Sacher, P. A., Applied Optics,

Vol. 5, p. 937 (1966) . (7) Pittman, M. A., Ofelt, G. S., Becher, J.,

Kernell, R. L., Applied Optics, Vol. 8, p. 209

(1969) . (8) Lusk, T. E., Electronic Design, 26 Oct. 1960.

(9) Cooley, J. L., Janda, R., Handbook of Space

Radiation Effects on Solar Cells, NASA SP 3003

(1963) . (10) Rittenhouse, J. B., Singletary, J. B., "Space

Materials Handbook," 3rd ed., (especially

chapter 8) NASA SP 3051 (1969) . (11) Canfield, L. R., Hass, G., Waylonis, J. E.,

Applied Optics, Vol. 5, p. 45 (1966).

(12) Gunter, R. C. Jr., Kennedy, E. F., Kaseta,

F. W., Jour. Opt. Soc. Am., Vol. 57, p. 1430

(1967) . (13) Malitson, I. H., Dodge, M. J. and Gonshery,

M. E., Jour. Opt. Soc. Am., Vol. 55, p. 1583

(1965) . (14) Dodge, M. J., Malitson, I. H., Gonshery, M. E.,

Jour. Opt. Soc. Am., Vol. 56, p. 1432 (1966).

(15) Meissner, K. W., Jour. Opt. Soc. Am., Vol. 31,

p. 405 (1941)

(16) Bennett, J. M., Jour. Opt. Soc. Am., Vol. 54,

p. 612 (1964) . (17) Molby, F. A., Jour. Opt. Soc. Am., Vol. 39,

p. 600 (1949)

(18) Browder, J. S. and Ballard, S. W., Applied

Optics, Vol. 8, p. 793 (1969).

(19) Hass, G., and Hunter, W. R., "Laboratory

Experiments to Study Surface Contamination and

Degradation of Optical Coatings and Materials

in Simulated Space Environments," presented at

the Optical Contamination in Space Symposium , Aspen, Colo., Aug. 13-15, 1969 and submitted

for publication in Applied Optics .

959

AUTHOR IaDEX

Baurer, T. [5j S-xtornal Spacecraft Conta'^u nation

.lodelin^r and Countemeasures 7; Beacliler, J.C. [55] ^^oeratinv^ Characteristics of the Air i'orcc rii:::ht Dynanics Laboratory Aeentry

Aose Tip (Rh2sT) racility ^'^3 Bean, B,L. [24] Ultraviolet and Charged

Particle Der.radation ol Aiu;:iinu;i and Silver

Coated I'LP Teflon Second Surface :iirrors 345 Blado, L.I). [9] In Situ Vacuun Cane Clalibration •

- by tiie Reference Transfer ;lethcd 141 Bortner, .i.ii. [5] iCxternal Spacecraft Containination

ilodelinu and Counterneasurcs 7'J Biillis, 'v.'.:!. [31] leasurenent "Tethods for .Micro- circuits 449

Buna, T, [4b J operation of a Larne Tbernal Vacuum

Chanljer at Martian Pressure Levels 725

Cappel, K.L. [Uo] The Dynamic Environment

Simulator, A Multi environmental Man-Rated Centrifuge . . . 587

Carden, Iv.H. [39] Explosive Oxidiations Initiated

by Sinulated .'eteoroid Penetration into Spacecraft

Atnospherc 569

Ciiar.fuerlin, P.C. [15] Method for .'lock Proper tv Determination in I'ltraiiigh Vacuum -09 Chuan, K.L. [6] Particulate Contaminant Measurement pv Quartz Crystal Microbalancc 1-^5 Cooper, A.M. [5] External Spacecraft Conta.T.ination

Modeling: antl Countermeasures 79 Corbin, w.il.dr. [7] Space Measurements of the Contamination of Surfaces by OGO-6 Outgassing and their

Cleanina bv Sputtering and Desorption -'-13 961
Cridlin, M.S. [44] A Method to Obtain an Ultraclean

Environment 661

Cunningham, F.G. [24] Ultraviolet and Charged

Particle Degradation of Aluminum and Silver Coated

FEP Teflon Second Surface Mirrors 345

Davidson, J.R. [57] Micrometeroroid Impact Simulation System 915

Donohoe, M.J. [26] The Degradation of Alzak by

Short Wavelength Ultraviolet Radiation . . 383

Donovan, F.M. [56] Investigation of the Charncteristics of Fluids Vented into a Vacuum 905

Dubinin, E.M. [20] Laboratory Experiment on Solar

Wind Interaction with Geomagnetic Field 297

Eck, M.G. [45] A Programmable Dynamic Thermal Vacuum

System for Solar Array Component Testing 669

[58] A Space Radiation Environment Simulator for

the Evaluation of Solar Cells 935

Enlow, D.L. [4] Contamination Studies in a Space

Simulated Environment 51

Evlanov, E.N. [13] Apparatus for Laboratory Simulation

of Ionospheric Flow Over On-Board Instruments of an AES . . 189 Fairbanks, J.W. [45] A Programmable Dynamic Thermal

Vacuum System for Solar Array Component Testing 669

Fayet, Ph., [23] Some Problems Concerning the

Simulation of Plastic Films Degradation by Ionizing

Radiations 323

Ferguson, M. J. [41] Use of the Ben Franklin Submersible As A Space Station Simulator 599

Flanagan, T.M. [36] Charged Particle Track

Implications for Accelerated Testing 525

Fleming, R.M. [14] Thermal Conductivity Measurements

of a Candidate Viking Heat-Shield Material After

Sterilization, and During Exposure to Vacuum, and

to a Simulated Martian Atmosphere 197

Foley, T.P. [37] A Simulation to Validate Future

Space Missions 535

962
Foreman, K.M. [47] Simulation of SurfaceAtmosphere Interaction Processes for Mars 699

Goldsmith, J.C. [2] Molecular Contamination in

Environmental Testing at Goddard Space Flight Center ... 1 Golub, R.A. [57] Micrometeroroid Impact Simulation

System 915

Greenwood, L.R. [14] Thermal Conductivity Measurements

of a Candidate Viking Heat-Shield Material After Sterilization, and During Exposure to Vacuum, and to a Simulated Martian Atmosphere 197

Harvey, W.L, [33] Establishment of an Optimum

Duration for Spacecraft Component Thermal Vacuum

Tests 473

Hearst, P.J. [21] Effect of Photodegradation on Organic Coatings 299 Henninger^ J.H. [26] The Degradation of Alzak by

Short Wavelength Ultraviolet Radiation 383

Hillberg, L.H. [52] Prediction of Mechanical

Erosion of Charring Ablators 837

Hirschfield, J.J. [58] A Space Radiation Environment Simulator for the Evaluation of Solar Cells 935

Humphrey, R.L. [51] Ablation Phenomena in

Supersonic Laminar and Turbulent Flows 813

[53] Heating in Cracks on Ablative Heat Shields 853

Karafiath, L. L. [16] Friction Between Solids and

Simulated Lunar Soils in Ultrahigh Vacuum and Its Significance for the Design of Lunar Roving

Vehicles 225

King, H.J. [19] Solar Wind Simulator 279

Koenig, J. A. [51] Ablation Phenomena in Supersonic

Laminar and Turbulent Flows 813

[53] Heating in Cracks on Ablative Heat Shields 853

Kostkowski, H.J. [12] Radiometry Standards - Present and Future 185

Leonas, V.B. [13] Apparatus for Laboratory Simulation

of Ionospheric Flow Over On-Board Instruments of an AES 189

963
Lilienkamp, R.H. [9] In Situ Vacuum Gage Calibration by the Reference Transfer Method 141

[10] Methods of Determining Residual Gas Composition from Residual Gas Analyzer Data 151

Maclnnis, J.B. [43] Advanced Diving Techniques

Applicable to Extended Mission Neutral Buoyancy

Space Simulation 651

Madden, M.T. [51] Ablation Pheonmena in Supersonic

Laminar and Turbulent Flows ... 813

[53] Heating in Cracks on Ablative Heat Shields 853

Mains, R.M. [35] Some Effects of Equalization on Acceleration Response and Fatigue Life Expectancy .... 509

Managadze, G. G. [20] Laboratory Experiment on Solar Wind Interaction with Geomagnetic Field 297

Maples, D. [56] Investigation of the Characteristics of Fluids Vented Into a Vacuum 905

May, C. B. [41] Use of the Ben Franklin Submersible

as a Space Station Simulator 599

McCally, M. [40] The Dynamic Environment Simulator, A Multienvironmental Man-Rated Centrifuge 587

McClung, R.W. [29] Monitoring Service Testing by

Nondestructive Testing 405

McClure, CM. Ill [42] Development of an Open

Circuit Cryogenic Life Support System for Use

in Neutral Buoyancy Space Simulation 623

[43] Advanced Diving Techniques Applicable to Extended Mission Neutral Buoyancy Space Simulation .... 651

Mcintosh, R. [26] The Degradation of Alzak by

Short Wavelength Ultraviolet Radiation 383

McKeown, D. [7] Space Measurements of the Contamination

of Surfaces by OGO-6 Outgassing and their Cleaning

by Sputtering and Desorption . 113

Metz, A.W. [37] A Simulation To Validate Future

Space Missions 535

Nakagawa, Hiroshi [46] The Cryopumping System

for the Space Simulation Chamber 683

Nelson, E.R. [2] Molecular Contamination in Environmental Testing at Goddard Space Flight Center 1 964
Nevman, P.A. , Jr. [8] Simulation of Radiation

from R.T.G. Power Sources 12s> [58] A Space Radiation Environment Simulator

for the Evaluation of Solar Cells

O'Connor, J,.>/. A Method to Obtain an Ultraclean Environment

Paillous, A. [17] ileal Time Sinulation of Atmospheric Balloon r,nvironment 245

[23] Some Problems Concerninr: the Simulation of Plastic Films Dcciradation by Ioni;',inn Radiations 323

Paul, F.W. [59] Properties of i.)ptical 'laterials

In Space 949

Parameswaran, V.]\. [34] Equipment to Study Hi;zh Velocity Dislocations in Matter ............ .437

Park, S.G, [24] Ultraviolet and Charged

Particle Dej^radation of Aluminum and Silver

Coated FEP Teflon Second Surface .'irrors 345

Pikus, I.M, [5] External Spacecraft Contan Lnation

Modelins and Countermeasures 79 podgorny, I.'i. [20] Laboratory Experiment on Solar i.'ind Interaction Wltw Geomagnetic Field , 297

Podnieks, E.K. [13] lethod for ilock Proper tv Determination in Ultrahi^ih Vacuum 209 Rabinowicz, E. [32] Life Testin?^, Using

Continuous Acceleration 457

Katliff, J. [43] Operation of a Large Therm.al Vacuum Chamjer at Martian Pressure Levels 725 Reetz, A. Jr. [1] Introduction to Contamination

Session C-Qt /iv.ai.lab],eX

Richter, R.L. [54] Apparent Opera tine: Limits of Arc Heaters with Respect to Total Enthaloy and

Stagnation Pressure <>79

.{iggs, '.-/.E, [49] A Cryopump for Steadv-State

Testing of Rocket Engines under Altittide

Conditions , 749

P.itter, .l.L, [3] Contaminant Measurement In Space Environment Simulation Chambers -3 965
Rogers, D.B, [40] The Dynamic EnvironnGnt Simulator, A Multienvironmental :ian-Rnted Centrifup;e 537

Roizes, A. [22] Relation Between Dose Rate Effects and Laboratory Simulation of Space

Radiation Damage , 313

Roussel, ;I. [17] Real Time Simulation of Atmos- pheric Balloon Environment 245

Roux, J. A. [11] Spectral Absolute Reflectance

Measurements of CO,, Frosts in the 0.5 to 12.0

Iticron Region , ,^ , » . . Ib5

Schutt, J.B. [25] The Interaction of Lov; Energy

Electrons witfi Polymeric Perf luorinated EthylenePropylene (FEP) 359

Sciiuttler, R. [22] Relation Between Dose Rate Effects and Laboratory Simulation of Space

Radiation Damage 313

Seiber, B.A, [11] Spectral Absolute Reflectance

;ieasurem.ents of CO^ Frosts in the 0,5 to 12.0,' ?ticron Region 165

Shane, R.S. [28] Plans and Status of N:IAB ad hoc

Committee on Testing for Prediction of Material

Performance in Structures and Components- 403

Shiralkar, B. [32] Life Testing Using Continuous

Acceleration ........ 457

Smitli, A.M. [11] Spectral Absolute Reflectance

Measurements of CO,, Frosts in the 0.5 to 12.0

Micron Region . . 165

Smith, K.A. [30] Signature Analysis - Non-Intrusive

Techniques for Incipient Failure Identification

Application to Bearings and Gears , 1^07 Steddum, R.E. [56] Investigation of tlie Characteristics of Fluids Vented Into A Vacuum 905

Sullivan, E.M. [50] Ablative Heatsiiields for

Planetary Entries - A Technology Reviev/ 751

Triplett, M.J. [49] A Cryopumn for Steady-State

Testing of Rocket Engines Under Altitude Conditions . . . 749

966
Tsimoda, R. [^6] The Cryopimiping System for the

Space Simulation Chamber <>V>2 UnansUy, 3.V, [13] Apparatus for Laboratory Simulation

of Ionospheric Flow Over On-Board Instruments of an AES ^^^'^

Van Lint, V.A.J. [36] Clinrecd Particle Track

Implications for AcceleratPtl Testin^^, 32b

V.'albcrr; , <.,,D. [50] Ablative ilea tshif.lds tor

i'lanetary }:]ntries - A Tcchnoloc:y Revievj 751 wannenacher, ii.L, [38] A Space Madintion

Lnvironment Simulator for tlie valuation of Solar Cells 933 wappaus, W.A, [27] The Photo-Response at an Alkaline

Paint Interface witli Aluminum .393

iv'certnan, J. [34] Kquipnent to Study liijr.h Velocity

Dislocations in Matter 4C7

Vieiclibrodt , li. [30] Signature Analysis - Non-Intrusive

Techniques for Incipient Failure Identification

Application to Bearings and Gears ^'^'^

h-illiains, C.i.'. [42] Development of an Open Circuit

Cryop,enic Life Support System For Use in Neutral

Buoyancy Space Simulation, . 623

'vvinkler, E.;i. [51] Ablation Phenomena in Supersonic

Laminar and Turbulent i'lov;s . ^"'^-^ [53] Heatinf^ in Cracks on Ablative Heat Shields ..... o53

Wolff, CM. [3] Contaminant M.easurement in

Snace Environment Simulation Chambers

VJood, B.E, [llj Spectral Absolute keflectance

Measurem.ents of CO^ Frosts in the 0.5 to 12.0

Micron Region . . -^^^ Yonda, A.V;. [37] A Simulation to Validate Future

Space Missions * ^35

Zuccaro, D, [18] Analysis of th.^ Simulation of

the Solar Wind 261

Zupan, Frank [38] General Purpose Centralip-.ed

Automatic Testinp, for Imnrovement of Shipboard

System Performance . 549

967

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