19820003334.Pdf

Fiwl Report No. IITR1-M06055-10

EVALUATION OF THE EFFECTS OF SOLAR RADIATION ON GLASSES

National Aeronautics and Space Adninistration George C. Marshall Space Flight Center Marshall Space Flight Center, A1 ahma 3581 2 Attention: Mr. R. L. Nichols, EH-34

Prepart hy:

Y. Harada I IT Research INsti tute 10 West 35th Street Chicago, 111i.dis 60616

August 1981

Ill RLStARCH INSTITUTE FOREWORD

The work described in this final report, IITRI-M06055-10, entitled "Evaluation of the Effects of Solar Radiation on Glasses," was performed under the sponsorship of the Marshall Space Flight Center. The work was conducted under Contract No. NAS8-33388 over the period July 9, 1979 to December 31, 1960, at IIT Research Institute. Mr. Y. Harada was the principal investigator on this program. We are pleased to acknowledge the valuable assistance and counsel of Mr. R. L. Nichols, NASA/Marshall Space Flight Center Contracting Officer's Representative on this program.

Respectful ly submi tted, IIT RESEARCH INSTITUTE

Senior Research Engineer Nonmetal1 ic Materials Section Materials and Manufacturing Technology

APPROVED :

llaurice A. :i,Howes, Director Materials and Manufacturing Technology

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ii TABLE OF CONTENTS

Page 1. INTRODUCTION ...... 1 2 . TECHNICAL BACKGROUND ...... 2 2.1 Spacecraft Optical Materials ...... 2 2.2 Space Environment Definition ...... 3 3 . RESULTS AND DISCUSSION ...... 4 3.1 Glass Materials ...... 4 3.2 Test Simulation Conditions ...... 6 3.3 Test Simulation Facilities ...... 6 3.4 Space Simulation Effects on Glasses ...... 7 3.5 Spacecraft Glass Materials: State-of-the-Art ...... 12 4 . CONCLUSIONS AN3 RECOMMENDATIONS ...... 13 REFERENCES ...... 14 APPENDIX A: Radiation Effects Laboratory. Boeing Areospace Company . 15 APPENDIX B: Space Environment Simulation Laboratory. TRW Defense and Space Systems Group ...... 20 APPENDIX C: Reflectance vs . Wavelength Data ...... 25

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iii LIST OF FIGURES

Figure No. Page 1 Effect of real-time solar UV, vacuum UV, protons, and electrons on tne reflectance of sapphire in situ ..... 8 2 Effect of solar UV, vacuum UV, protons, and electrons on the reflectance of 7971 low expansion glass in situ ... 9 3 Solar absorptance vs. exposure time to simulated space environment for glass materials ...... 11 4 Spacecraft W indows ...... 12

LIST OF TABLES Table No. PEge 1 Window Materials for Space Simulation Studies ...... 5 2 Solar Absorptance Changes in Space Simulation Studies ... 7

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iv 1. INTRODUCTION

The space environment to be encountered by space vehicles is extremely complex, involving particulate as well as ultraviolet radiation. The penetrat- ing radiation environment may result from a variety of sources of which the most important are probably cosmic radiation, trapped radiation, auroral radiation, and solar flare radiation. It ir possible that such high energy protons and electrons will have a more significant effect than ultraviolet on spacecraft materials such as glasses, and that synergistic effects from the variety of radiations will occur. A wide variety of functions must be performed by glass components on a space structure, ranging from windows for simple viewing in the visible wavelength region, to lenses for ir detection devices which must transmit at much higher wavelengths. Thus, radiation damage must be evaluated with a view toward the particular application. For example, mslly infrared-transmi tting materials have been darkened to near opaqueness in the visible region, but the ir transmission remains virtuslly unaffected so that the functimality is ma i n tai ned . 1 In a recent program conducted at IITRI, studies were conducted to evaluate the degradation of glass due to .*lectromagnetic and particulate radiation in a space environment. initial work was concerned with attempts to define the space environment. Secondly, a literature review was made on radiation damage mechanisms in glasses. Four optical materials were exposed to simulated solar and particulate radiatic.1 in a space environment. Sapphire and fused silica experienced little change in transmittance while optical crown glass and ultra low expansion glass darkened appreciably. The objective of the current program was to carry this technology further by achieving a more complete ana1,ysis of the 500 hour simulated space exposure test. Additionally, studies were conducted to aid in sample selection for a 100 hour simulated exposure test.

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I 2. TECHNICAL BACKGROUND

Current materials for space applications must endure increasingly stringent environments as missions are displaced outward from near-earth orbits, and are designed for longer durations of several years. In a synchronous equatorial orbit, for example, electrons and protons as well as u;traviolet radiation are considered to be prime contributors to spacecraft surface meterials degradation.' The ramifications for glass materials are discussed in the following sect i ons .

2.1 SPACECRAFT OPTICAL MATERIALS The materiais used aboard spacecraft for optical applications are in general those which transmit in the solar spectrum, nominally from 200 nm to 3000 nm. For the 200-1000 nm regions, silicon oxide based materials, notablj fused silica of high purity, have foirnd widest application. The selection of suitable optical materials, however, involves not only initial optical properties, but optical stability, mechanical properties and performance, contamina- tion potential, contaminabili ty, electrical properties, rf transparency, and, of course, cost. From a materials point of view, there are three regions of optical interest: UV (<300 nm), visible (300-700 nm), and near ir (>700 nm). For UV applications, silica appears optimum, although some alkali halides might be considered; in any CV applications, inorganic Raterials are almost mandatory. Visible regicn applications quite probably can utilize organic- based transparencies, except there UV transparency is also essential. Infrared applications--both traditionally and, quite logical ly, from materials initial properties and space stability considerations--have required inorganic materials Organic materials inevitably exhibit either ir "fingerprint" spcxtra, poor mechanical properties, or other objectional properties. Simply stated, the primary criteria for selection of spacecraft opticdl materials are initial properties, environmental stability, and rnechanica I/physical/environmental considerations. Onlj .)ne region, the visible, was investigated on %P IITRI

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2 program, and the materials were all inorganic oxides,' These are discussed in a later section. For spacecraft use, in general, standard radiation resistant optical materials we sapphire and Corning 7940 fused silica glass. These are used for the front elements of optical cystems whenever possible. If other materials must be used, a 2-3 nun thick fused silica glass window is used in f-ont of tile optical components. The window eliminates almost 90% of the particulate radiation and also the very short wavelength ultraviolet. If a window cannot be used, then the equipment is designed to compensate for the degradation. For example, solar cell cover glass, Corning 0020, degrades approximately 5% in one year when exposed directly to solar radiation. Hence, for a one-year mission, the cell is designed to have an initial output which is 5% greater than required.

2.2 SPACE ENVIRONMENT DEFINITION The highly complex and dynamic nature of a space environment demands in- situ and current data for good definition. This type of information is best obtained from a source such as the Goddard Space Flight Center which continual- ly monitors and analyzes such data. IITRI in its previous program established a dialogue with Mr. E. G. Stassinopoulis, Senior Acquisition Scientist, Radiation Environment, NASA/Goddard Space Flight Center. Mr. Stassinopoulis pointed out that the space environment near the iarth was extremely variable, both spacial ly and temporally. Photon radiation remains relatively constant at 1 s~',, )ut particulate radiation, which is mainly protons and electrons, strongly depends on the orbit of the spacecraft and the solar activity during the time in orbit. The space environment in which the glasses are to be used was also discussed with Mr. R. Nichols and Mr. 3. Wright, Astrophysics, and with Mr. P. Priest, Space Programs, NASA/Marshall Space F1 ight Center. The conclusion from these discussions was that average conditions should be selected for initial simulated solar radiation experiments. Later experiments could then be designed to explore the effect of variations from average,

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3 3. RESULTS AND DISCUSSION

The current program was concerned with evaluation of a space simulation test of optical materials for spacecraft and with determination of materials anticipated for use in future spacz missions. The results are discussed in the following sections.

3.1 GLASS MATERIALS

In current technology, sapphire or fused silica glass appear to be the best materials available for resisting optical degradation in an ultraviolet/ particulate radiation environment. In establishing glass materials for testing the following ranking of resistance to radiation damage was developed: 1. Excellent - sapphire 2. Good - fused silica 3. Fair - optical crown glass 4. Poor - ultra low expansion glass These materials were chosen for exposure for the following reasons:

1. Sapphire is exceptionally strong and hard (9 on the Mohs scale) and is by all reports the best material for resisting radiation damage. Union Carbide Czochralski grown sapphire crystals were chosen. 2. Fused silica is a close second to sapphire in strength and hardness (7 on the Mohs scale) and in radiation resistance, and has been used on numerous space missions. Corning 7940 synthetic fused silica was chosen because it is a well-characterized, uniform material with very low solar absorptance.

3. Optical crown glass is used in many optical devices as the front ele- ment of the objective lens. It was observed to krken when exposed directly to the space environment in Spacelab experiments. Shott BK-7 was chosen since it is the most widely used glass of this type and readily available.

4. Ultra low expansion glass is a very interesting optical material which is almost imnune to thermal shock failure. Windows made of this material could be used witiiout failure for observation during the thermal heating caused by reentry. Corning 7971 was chosen because it was readily available. Ill AESLARCH INSTITUTE

4 Specimens were obtained directly from the manufacturer, The sapphire specimens were obtained from the Crystal Products Division of Union Carbide Corporation as standard optical windows, 25.4 mm diameter and 2 mm thick. The optical glass specimens were also standard windows, 25.4 mm diameter and 3 mm thick,obtained from Melles Griott, Inc. , Irvine, California. The fused silica and ultra low expansion glasses were obtained from the Corning Glass Works as rough cut discs which were then ground and polished io 25.4 rn diameter and 3 mm thick. A1 1 the specimens were made of commercially produced materials. Standard industrial pitch polishing procedures were used to produce specitnens which were flat within 1/4 wavelength of light per 25 mn, parallel within 30 seconds of arc, and surface quality of 80-50 scratch and dig. Samples were cleaned,and approximately one half of the surface of each sample was coated with aluminum and then with silica by vacuum evaporation. The coating provided a back surface mirror for transmission optical measurement by double reflection. The samples of the selected glasses were furnished to Boeing Aerospace Corporation, who anchcted the simulated exposure. Brief descriptions of the glasses, tilt! number of specimens supplied, and the number of each type put into the exposure array, are indicated in Table 1.

TABLE 1. WINDOW MATERIALS FOR SPACE SIMULATION STUDIES

Number Number Glass Type Supplied Tested 7971 (low expansion glass) 3 2 BK-7 (crown glass) 3 2 7940 (fused silica) 2 1 Sapphire (alpha-alumina) 2 1

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5 3.2 TEST SIMULATION CONDITIONS Test irradiation parameters were determined on discussions with NASA/ Marshall and NASA/Goddard personnel, and were as follows:

1. Solar UV--one sun intensity, assuming earth orbit and air mass zero conditions; 0.2 0.4 micrometer spectral content from xerion arc continuum; infrared h > 1.4 micrometer sup- pressed by water column around xenon arc,

2. Vacuum UV--one sun intensity at hydrogen Lyman-a 0 wavelength (1216 A); radiation introduced into chamber through the VUV source's window.

3. Electrons--intensity (flux) 1 x lo9 e/cmZ-sec, at 50 KeV energy. 4. Protons--intensity 1 x 109 p/cm 2-sec, at 30 KeV energy; introduced into chamber after magnetic+ bending to separate mass-one (H') from Hp and other species.

5. Temperature--2O0C for a1 1 test samples,

6. Vacuum--average of at least 5 x loo8 torr during exposure, and 2 x torr during spectral reflectance measurement periods.

3.3 TEST SIMULATION FACILITIES

The space simulation test was conducted at the Boeing Aerospace Corporation, A description of their facilities is presented in Appendix A. During the course of this program, a visit was made to the facilities at TRW in Redondo Beach, California. A tour of the space environmental systems laboratory was made through the courtesy of Mr. G. Brown. These laboratories are described in the literature in Appendix B.

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6 3.4 SPACE SIMULATION EFFECTS ON GLASSES Examination of the specimens imnediateiy aftzr 503 hours of exposure showed that the optical glass and ultra low expansion glass specimens had darkened appreciably. All specimens web E! physically intact after exposure and had not suffered any gross mechanical damage. Heating and charge !-,i;ldup during exposure was probably very small, since the specimens were in good thermal and electrical contact with the water-cooled copper support. Graphs of the absolute reflectance data for representative specimens are shown in Fig. 1 (sapphire) and Fig. 2 (low expansion glass). 71- each of these figures the uppermost data curve was obtained before exposure bc,an. Small reflectance changes were measured at each subsequent measurem, it Dolnt. with the 503-hour data being represented by the lowest curve. Tabular dau of reflectance values vs. wavelength for the various samples at the difiirent levels of exposure are contajned in Appendix C.

The computer program also calculated solar absorptance coefficients (as) zorresponding to each measurement. Ti ese are tabulated below. TABLE 2. SOLAR ABSORPTANCE CHANGES If: SPACE SIMULATION STlJDIES .. Solar AbsorDtaupfficiPnt -- Pre-I rrad After Expasure for Increcise A*' Sanipl e in Vacuum 50 hr 154 hr 313 hr 50m 5C3 hr Sapphire 0.109 0.112 0.110 0.115 0.118 0.009 8,ZX) Crown No. 3 0.094 0.106 0.109 0.114 0.118 0.024 (25.5%) Crown No. 2 0.098 0.109 0,111 0.117 0.119 0,021 (21.4%)

Low exp. No. 1 0.085 0.092 0 094 0.101 0.108 0.~173 -. 1 Low exp. No. 2 0.083 0.083 0.090 0.097 0.108 0.020 ij Fused Silica 0.078 0.080 0.G81 0.088 0.087 u.009 (11.44

The change in the absorptance of the sapphire r-.J fused silica specimens was very small and close to the limit of accuracy af the measurenents.

A plot of these as values against exposure time (Fig. 3) shows that sapphire and fused silica have no pronounced initial increasc in as but, rather, have a somewhot' regular as increase rate. On the other hand, the iiT RLSEAUCH INSTITUTE

7 @ HMISPHERIL?. SPECTRRL ABSOWTFINCE

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9 crown glass samples and the low expansion glass samples underwent larger initial increases in as that tapered to slower rates of increase. Examination of the reflectance curves before and after various exposure times did not reveal the appearance of any new absorption regions. Rather for all specimens there was an increase in absorption with exposure time, particu- larly at the fundamental absorption edge in the near UV. Examination of the more detailed digital data in Appendix C, also did not reveal any particular pattern in reflectance chawjes. The absorptions at 1.3 and 2.2 micrometers of the glasses relate to the silica network and was most pronounced in fused silica. Sapphire being alumina did not, of course, show these. it was sur- prising that the modifiers present in the optical glass and ultra-low expansion glass did not result in new absorptance regic-s after exposure, although they do suppress the silica regions in optical glass. The change in solar absorptance values with time which was tabulated in the previous section and shown in Fig. 3, indicate that the a priori ranking of the optical materials was correct. Sapphire is the most stable material followed closely by fused silica, while crown optical glass and ultrzl-low expansion glass show stronger degradation. For sane applications fused silica may be preferred because of its lower initial absorption,since even after 503 hours of exposure it was still less than sapphire. The optical glass and ultra low expansion glass,on the other hand, while reasonably stable for 503 hours, may have limited usefulness for very iong space missions. However, the utility of any optical materials wiil be determined by the particular miss i on requ i remen ts . Other analyticel methods which were examined were thermoluminescence and chawes in microhardness. The measurement of microhardness was used in an attempt to determine any change in mechanical integrity. No significant differences were observed for exposed f 500 hr simul2ted exposure) vs. virgin materials. Apparently this test is not sensitive enough to show changes (if any) in the bonding of these glasses. Thermoluminescence measurement was thought to be another potential analytical method. However, it was determined that the assembly of the needed equipment,along with perform-ing of such experiments, !vas beyond the financial scope of this program.

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10 0.12

0.11

v) 3- Q, :: 0.10 (D c, n L 0 nv) U L m Low expansicn No. 1 .- 0 v1 0.09 -Low expansion No. 2

V Fused silica 0. C8

0.97 7-100 2@0 300 400 500 Exposure Time, hr

Figure 3. Solar absorpt'ince vs. exposure time to simulated space enviromicnt for qlass inaterials

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11 3.5 SPACECRAFT GLASS MATERIALS : STATE-OF-THE-ART Dialogue with engineers involved with spacecraft material, was carried on during this program. People who have been contacted include: Bill Carroll (JPL) Carl Maag (JPL) , Larry Fogdall (Boeing) , Gene Rusert (McDonnell Douglas) Me1 Clancy (Rockwell), A1 Rubin (Aerojet) and Jerry Brown (TRW). The consen- sus appears to be that fused silica is acceptable for all current missions including those in synchronous orbit. Problems or potential problems of any significanct with glasc materials have not yet surfaced.

Examples of glass material laminates used for viewing windows in space- crafts are shown in Fig. 4. This information was supplied by Mr 6. Rusert of McDonnell Douglas which was responsible rrrr these spacecraft.

SPACE ENVIRONMENT

Used on Mercury 0.330 in. C7913 Vycor I and Gemini space- 380 in. C7913 Vvcor \ craft 0.223 in. C1723 Tempered 1 Alumi nosi 1i cate SPACE ENV I RONMENT -t I Used on airlock module for Apollo- 0.420 in. c79\3 %yew Manned orbiting 0.240 in. C1723 A1:ininosiliciite I 1aboratory

Figure 4. Spacecraft Windows

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12 4. CONCLUSIONS AND RECOMMENDATIONS

Additional studies were conducted in this program in an effort to extract additional information from the results of a 500 equivalent hours space simulation experiment. These studies have shown the following:

1. Chczges in reflectance spectra do not show zny significant differences in absorption peaks which could be used to indicate degradation mechanisms. It is possible that the test period was too short to produce any such differences. 2. No significant changes were observed in microhardness of glass samples due to UV-vacuum- particulate irradiation.

3. Longer term testing is needed in order to obtain information applicable to longer term space missions of up to 10 years.

4. Fused silica and "Vycor," a high-silica (96%) glass, appear to be the preferred glass materials for viewing windows on spacecraft. Among the people interviewed with interests in this area, no significant problems were appar- ent.

5. For any future tests, the Boeing and TRW space simulation laboratories would appear to be signi f i cant 1y dif f erent .

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13 REFERENCES

1 Firestone, R. F.,and Harada, Y., "Evaluation of the Effects of Solar Radiation on Glass," IITRI Final Report No. 06139 for NASA/Marshall Space F1 ight Center (January 1979) . 2. Kurland, R. K, et al., "Properties of Metallized Flexible Materials in the Space Environment," Final Report, SAMSO TR 78-31, TRW/Defense and Space Sys tems Group (January 1978). 3. Zerlaut, G. A., Harada, Y., and Tompkins, E. H., "Ultraviolet Irradiation in Vacuum of White Spacecraft Coatings." in Symposium on Thermal Radiation of Solids, NPSA SP-55, Scientific and Technical Information Division, Washington, D.C., pp. 391-420 (1965).

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14 APPENDIX A Radiation Effects Laboratory Boeing Aerospace Company Seattle, Washington 981 26

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15 RAD I AT I ON EFFECTS LABORATORY BOEING AEROSPACE COMPANY

The simulated space exposure for this program was performed in a facility at Boeing known as the combined radiation effects test chamber (CRETC). This facility was originally designed after completion of a number of years of testing thermal control coatings for government and corporate customers. It has as high a quality of simulation of the space radiation environnent a; knowledge and the state-of-the-art will allow. Exposure capabilities include electron radiation and Lyman-a vacuum ultraviolet sources in addition to the usual proton accelerator and solar simulator.

The space simulation capabilities of the CRETC facility can be sunnnarized as follows: 1. Continuum ultraviolet radiation (xenon arc discharge) at selectable intensities ranging from less than one solar constant to 20 solar constants (1 AU) , simultaneously with: 2. Electrons with energies between approximately 10 eV and 200 keV and/or protons with energies from 0.5 to 85 KeV (kilo electron volts). Electrons of greater than ~15KeV are foil- scattered; protons are magnetically analyzed. 3. Vacuum ultraviolet radiation (VUV), primarily 0 the Lyman-a wavelength of 1216 A, from a con- tained discharge (no introduction of contami- nation into vacuum chamber). Intensity selectable up to and above one VUV sun at 1216 at 1 AU.

4. Control led temperatures for test and reference (standard) samples; temperatures range from

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16 -195°C (-320°F) to +180°C (+36OoF). Tempera- ture control is not interrupted for measurements in the chamber's integrating sphere. 5. hcuum pumping (both rough and final) without resorting to organic and other contaminating fluids. The sequence used is (a) dry nitrogen gas aspiration, (b) cryo-sorption, (c) large- surface LN2 cryogenic, and (d) ion pumping, to obtain a 5 x torr vacuum before testing begins. 6. Extensive automation, interlocks, and sequen- tial shutdown procedures during unmanned night-time operations, to allow as high a reliability as possible during long-term, continuous testing.

7. High-precision spectral refleL "ante data system with in-situ integrating sphere and doubl e-beam spectrophotometer coup1 ed to a data-logging module whose output is ready for computer processing. 8. Residual gas analysis from 1-100 amu with scan rates down to 0.1 sec/scan and minimum partial pressure detectability of 2 x loc1' torr. 9. Electrical discharge event counter with re1 a- tive magnitude indication.

The CRETC utilizes an integrating sphere reflectometer with detectors in situ. Only the measurement light sources, monochromator, and electronic and 1 ight chopping apparatus are external to the chamber. Sample reflectance measurements are made relative to the reflectance of the integrating sphere's magnesium oxide wall. Normalization to absolute reflectance (derived from National Bureau of Standards and other known reference measurements) is handled by computer since a1 1 original sample daLa are comp.iter-processed routinely.

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17 The thermophysical prb,perty of chief interest for this program, solar absorptance, is derived from solar reflectance, which is defined as

/Is ( X )R( A)dA Solar re;lectance, Rs = A)dA where Is(A) is the solar irradiance as a function of wavelength A, and R(X) is sample reflectance, generally a function of A. By measuring the reflectance of transmissive glasses with a suitable metal backing (aluminum), double sens tivity to radiation-induced changes is obtaitied. The measurement beam passing through the integrating sphere makes a double-pass through the sample, yielding a measurement proportional to change in transmission squared. This allowed a more exact determination of damage in the various optical quality glasses studied and compared during this program. In practice, change in transmission generally will be found to be half the measured change in reflectance. Radiation dosimetry systems are integral to the CRETC, and operating personnpl have long-time euperience in obtaining the pertinent irradiation and exposure parameters. Faraday cups and tabs are both employed for measuring oncoming particle radiation beans at the sarn?le plane.

For ultraviolet radiation paraiiieters, sun rates are determined from radiometer output levels taken with and without a UV-absorbing filter over the radiometer detector. Since the UV-absorbi ng f i1 ter a1 so excludes ten percent of the incident radiation at wavelengths longer than the ultraviolet (five percent reflection at each surface of the filter), a correction is made for radiometr readings taken with the filter over the radiometer sensor. For a total radiation reading T and a UV-filtered reading F,

10 T-TF Ultraviolet Sun rate = o.09,Y- = 1.37 (T - 1.11 F), where 8.0 is the radiometer sensitivity in millivolts per incident sun (~0.14 watts/cm 3 ) and 0.91 represents the ultraviolet content of the sun (at air mass zero). The uniformity of ultraviolet radiation intensity across the sample array is determined by "mapping" with the radiometer held in a precise jig.

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18 Spatial uniformity of ultraviolet radiation can be maintained within plus or minus 10 percent across the sample array. The F and T values indicate that the ultraviolet content of the long-arc xenon sources is approximately 10 percent of their total input. Characteristic of all xenon arcs, the shape of this ultraviolet content is somewhat more steep than the sun at air mass zero (AMO).

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19 APPENDIX B

SPACE ENVIRONMENT SIMULATION LABORATORY TRW Defense and Space Systems Group Redondo Beach, CA 90278

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20 TRCry- 2SSG SPACE 4??I IRONPIENT S I MU LA T I 9 ?l LA I! 0 RATI: R Y

TRIj DSSG-Vulnerabil ity and Hdrdness Laboratory of the Advanced Technology Division in conjunction with the Engineerinj Laboratory ot :he Sr xe Systems Division operates a Space Environmcntal Systems Laboratory for trle i!se of external as well as in company users.

The laboratory is located in TRW DSSG BUIWIXG 84 at 1452CJ av-idtion Boulevard, in Lawndale, California in close proximity to the main TF 5G cimplex in Redondo Beach, Cal ifornia.

The original design of this laboratory was primarily direc.-- Lowards the study of optical and mechanical property degradation of relatively thin film materials, solar cells, and other structural elements of spacecraft directly exposed to the natural space environmet.-.

A view of the laboratory in one configuration is shown in F-e 1. A number of vacuum irradiation stations/chambers are arranged around a 1.2 MV Van de Graaff electron accelerator in an irradiation vault. The electron beam from the accelerator can be directed into each of the chamber positions using a switching magnet and solenoidal focusing magnets. Each chamber is provided with additional simulation sources for simultaneous cperatic,is. The diver- sity of these sources is tabulated in TabZe 1'. Table! II shows the capabilities which presently exist at each chamber position.

Each chalnber has an available capacity of 18 inches diameter by 30 inches long. AI, are pumped by at least 400 l/sec vac-ion pumps. The "in-situ" optical chamber and "ex-situ" mechanical test chamber are also pumped by tita- nium sublimation pumps. The chambers can be easily modified for other types measurenients because of the 1arge number of instrumentation/entry ports available on each.

Figure Zld and fb) gives the simulation spectral capability of the ultraviolet sources in use. As listed in the table the "near" ultraviolet capability is available presently for all three chambers while the far ultraviolet capability is available only at the 50' position. The laboratory is designed for continuous semi-automated operation. At present all systems and er,vironments can be operated iattended except the Van de Graaff system which is being modified to allow unattended operation of it as well. During such operations key parameters are checked periodically and the latest data is available for remote telephone readout on demand. @ut of limit key pararr,eters can be alarmed to a central security system from whence repair personnel can be located immediately by beeper.

21 22 Tabh I - ENVIROMlErlTAL SIMULATION DESCRIPTION

Rdirtion

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P-i~i2 - SPECTRAL WRC’ ’ .idn OF ULTRAVIOLET SOURCES Trtbk Ir - LABORATORY CAFABILITIES

- ~ ~ U1 travfolst (0.10-0.40 urn) 2b pc (1k2 &tn) Olr. Spectral Rrflcctancc In-Sftu Optics1 Protons (1-30 kVI Olr. Diffuse Rrflrctrnrc E 1cc trons (1-10 kQb) 4 contrmtnatinn Btdlmttonsl Rrflcctrnce (60.1 E lectronr (30-1Wd CIV) YI tmss Slrter vacum (to sr10'*T) Specinens mvcd for total (to -100' C) mi ttrncc misw'etvnts.

U1 trrviolct (0 18-0.40 ,,n) 17 PC (1 in. ospe Tensile Stress-strain to In-Si tu Tcnsi le L let trons (5J-lOikl Lev) lrnath by C in. .'; strrtn level. (30') VICUUn (to 5\10-&l) width) Tensilr Strcss-strrin to Ircnperb ture (tz -:o* C) f allure.

~ Solar Test E ltctrons (5O-lOOO CCU) 12 pc (1x2 all) Spcclmra3 mvcd for 1 I Twperrtun (to -20' C) chrrrc trrist lcs tcs: 1 no. U!trrviblet (0.18 to 0.40 vm) fpelarns mrtdto drv Ex-Sltu kchanical (-M*) I vrcuun lalroth by $ In. testing. Uidth)

Additional inforration may be obtained by cnlling R. Kurland- 213/535-1?16 cr W. BWJCJS- 21 31536-31 35. lours of the facility can be arranged thrcugh the Building Operations Manager, P. Gullfoyle- 213/535-0056.

24 BPPENDIX C REFLECTANCE VS. WAVELENGTd DATA 500 Hour Space Simulation Experiment

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