(NASA-CP-3278) PROCEEDINGS OF THE 

13TH SPACE PHOTOVOlTAIC RESEARCH 

AND TECHNOLOGY CONFERENCE (SPRAT 

13) (NASA. Lewis Research Center) 

421 p 

Hl/33 

N9S-20S02 

--THRU--

N95-20543 

Unclas 

0038432 
NASA Conference Publication 3278 

Proceedings of 

the XIII Space 

Photovoltaic 

Research and 

Technology 

Conference 

(SPRAT XIII) 

- __ __ 0_ 

Proceedings of a conference held at 

NASA Lewis Research Center 

Cleveland, Ohio 

June 14-16, 1994 

eo"" 

, L ,;(;:00 -;/ . ..,.,.. ... . ~ "" . 

.... 

National Aeronautics and 

Space Administration 

Office of Management 

Scientific and Technical 

Information Program 

1994 

-_._0 __ 
FOREWORD 

David M. Wilt 

NASA Lewis Research Center 

The Thirteenth Space Photovoltaic Research and Technology (SPRAT) Conference gathered 

representatives from 26 commercial corporations, 8 universities and 7 governmental agencies, including 

Europe, for two and a half days of presentations and discussions regarding the status and future of space 

photovoltaics. The conference was well attended, with over 100 attendees, and included 38 technical 

papers, 6 program reviews and 5 workshop discussions. 

The effects of shrinking research and development budgets were evident in 'the focus and tone 

of this SPRAT. Most attendees appeared to be oriented toward near term, system oriented projects and 

fewer were involved in long term, high risk research. It was generally agreed that space power 

requirements would continue to move toward smaller «2 kW) power levels. Most future applications are 

believed to be in traditional orbits (LEO,GEO) although interesting opportunities may be found in high 

radiation, mid-altitude o·rbits useful for global communication networks. New solar cell devices and 

materials will be difficult to introduce unless they are mission enabling, or offer significant cost and/or 

performance benefits. The attendees were unable to come to a consensus regarding the type of array 

(eg. rigid, concentrator, flexible) suitable for specific missions. Many factors outside the realm of 

photovoltaics influence the selection process. These topics and many more are covered in the following 

pages of this record. 

Organizing and conducting the SPRAT conference requires a dedicated committee. The 

members of my team were: Navid Fatemi as logistics chairman, George Rybicki as publications chairman, 

Dennis Flood as conference mentor and Pat Nicewander as (the glue that holds everything together). 

The committee was also assisted by Karen Wester who organized the social functions. Their efforts and 

attention to detail were responsible for the success of this conference and the maintenance of my mental 

health. 

~ "AGi BLANK NOT fJlMED 

iii 
TABLE OF CONTENTS 

Goddard Space Flight Center Solar Array Missions, Requirements, 

and Directions 

Edward Gaddy and John Day, NASA Goddard Space Flight Center .. , .... ....... 1 

BMDO Photovoltaics Program Overview 

Leonard H. Caveny, Ballistic Missile Defense Organization, 

and Douglas M. Allen, W.J. Schafer Associates ............................... 13 

Air Force Activities in Space Photovoltaic Power System Technology 

Kelly Gaffney, U.S. Air Force ............................................... 23 

An Update on the Department of Energy's Photovoltaic Program 

John P. Benner and Mark Fitzgerald, National Renewable Energy 

Laboratory ... .. . ... ... .. ...... .. . ......................... . . .. . . .. . .... . 31 

Europe's Space Photovoltaics Programme 

Klaus P. Bogus, European Space Agency . . ..... .. , .............. , .. .. ...... 35 

SESSION 1-lnP CELL DEVELOPMENT 

High-Efficiency, Deep-Junction, Epitaxial InP Solar Cells on (100) and 

(111)B InP Substrates 

R. Venkatasubramanian, M.L. Timmons, and J.A. Hutchby, Research 

Triangle Institute, and R. Walters and G. Summers, U.S. Naval Research 

Laboratory .......... . ...................... . ............................ 43 

Chemical Beam Epitaxy for High Efficiency Photovoltaic Devices 

A. Bensaoula, A. Freundich, M.F. Vilela, N. Medelci, and P. Renaud, 

University of Houston ... . ................. . ................... . ........... 53 

Status of Diffused Junction p+n InP Solar Cells for Space Applications 

Mircea Faur, Cleveland State University; Maria Faur, NYMA, Inc.; 

D.J. Flood and D.J. Brinker, NASA; C. Goradia, Cleveland State 

University; N.S. Fatemi and P.P. Jenkins, NYMA, Inc.; D.M. Wilt and 

S. Bailey; NASA Lewis Research Center; M. Goradia and J. Moulot, 

Cleveland State University ... ..... ........ ....... .. . ...................... 63 

High Quality InP-on-Si For Solar Cell Applications 

Zane A. Shellenbarger, Thomas A. Goodwin, Sandra R. Collins, and 

Louis C. DiNetta, AstroPower, Inc. . ...... . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 81 

PIN InP Solar Cells on Ge Wafers 

Steven Wojtczuk and Stanley Vernon, Spire Corporation, Edward A. Burke, 

Private Consultant. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 91 

v 
SESSION II-THIN-FILM AND HIGH EFFICIENCY CELL DEVELOPMENT 

Status of Flexible CIS Research at ISET 

B.M. Basol, V.K. Kapur, A. Minnick, A. Halani, and C.R. Leidholm, 

International Solar Electric Technology (ISET) .. .. . ..... .. . ..... . . . .. .. . . . .. 101 

Development of a Lightweight, Light-Trapped, Thin GaAs Solar Cell for 

Spacecraft Applications 

Margaret H. Hannon, Louis C. DiNetta, Michael W. Dashiell, John R. 

Cummings, and Allen M. Barnett, AstroPower, Inc. ... . . ... .. .... . .. . . . .. . .. . 107 

Advances in Polycrystalline Thin-Film Photovoltaics for Space Applications 

Bruce R. Lanning, Joseph H. Armstrong, and Mohan S. Misra, Martin 

Marietta Corporation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 11 5 

Amorphous Silicon Thin Films: The Ultimate Lightweight Space Solar Cell 

G.J. Vendura, Jr., M.A Kruer, H.H. Schurig, M.A Bianchi, and J.A Roth 

TRW Space and Technology .. .. . . . . ... .. . .. ..... . . . ...... ... . .. . . . ... . .. 125 

Modeling of High Efficiency Solar Cells Under Laser Pulse for Power Beaming 

Applications 

Raj K. Jain, NASA Lewis Research Center, and Geoffrey A. Landis, 

NYMA, Inc . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135 

Calculation of Near Optimum Design of InP/lno.53Gao.47As Monol ithic 

Tandem Solar Cells 

P. Renaud, M.F. Vilela, A Freundlich, N. Medelci, and A. Bensaoula 

SESSION III-RADIATION DAMAGE AND ENVIRONMENTAL TESTING 

Performance, Defect Behavior and Carrier Enhancement in Low Energy, 

Proton Irradiated p+nn+ InP Solar Cells 

I. Weinberg and G.C. Rybicki, NASA Lewis Research Center, 

C. Vargas-Aburto, Kent State University, R.K. Jain, University of Toledo, 

141 

and D. Scheiman, NYMA, Inc. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 149 

Electron-Irradiated Two-Terminal, Monolithic InP/Gao.47Ino.53As Tandem 

H.L. Cotal, R.J. Walters, and G.P. Summers, U.S. Naval Research 

Laboratory, and S.R. Messenger, SFA, Inc ... .. .. . . . ...... . .. . ............ .. 159 

Radiation Effects on p+n InP Junctions Grown by MOCVD 

S.R. Messenger, SFA, Inc.; R.J. Walters, U.S. Naval Research Laboratory; 

M.J. Panunto, University of Maryland Baltimore County; and G.P. Summers, 

U.S. Naval Research Laboratory . .. .... . ... . .... . .. . .. . .. .. . . ..... . ...... . 167 

Design of High-Efficiency, Radiation-Hard, GalnP/GaAs Solar Cells 

Sarah R. Kurtz, K.A Bertness, AE. Kibbler, C. Kramer, and J.M. Olson, 

National Renewable Energy Laboratory . .. . ... . .. ... . .. . . .. . .... .. .. . . . .. . 181 

Investigation of the Stability and 1.0 MeV Proton Radiation Resistance of 

Commercially Produced Hydrogenated Amorphous Silicon Alloy Solar Cells 

Kenneth R. Lord II, Michael R. Walters, and James R. Woodyard, Wayne 

State University . . . . .. .... .. . . .. ......... ..... .. . .... . ..... .. . . . ..... ... . 187 

VI 
The Use of Displacement Damage Dose to Correlate Degradation in Solar 

Cells Exposed to Different Radiations 

Geoffrey P. Summers, U.S. Naval Research Laboratory; Edward A. Burke, 

Philip Shapiro, Richard Statler, and Scott R. Messenger, SFA, Inc.; and 

Robert J. Walters, U.S. Naval Research Laboratory ........ ... .............. 197 

Annealing of Irradiated n+p InP Buried Homojunctions 

R.J. Walters and G.P. Summers, U.S. Naval Research Laboratory, 

M.L. Timmons, R. Venkatasubramanian, J.A. Hancock, and J.S. Hills, 

Research Triangle Institute . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 203 

Qualification Test Results for Blue-Red Reflecting Solar Cell Covers 

W.T. Beauchamp, OCLI ............................ .. .................... 213 

UV Testing on INTELSAT-VII, VilA, and VIII Solar Cells 

A. Meulenberg, COMSAT Laboratories ...... . ....... ... .......... ... ...... 223 

Parasitic Current Collection by Solar Arrays in Leo 

Victoria A. Davis and Barbara M. Gardner, S-Cubed Division of Maxwell 

Laboratories . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 227 

SESSION IV-FLIGHT RESULTS 

Final Results of the Advanced Photovoltaic Experiment 

David J. Brinker, NASA Lewis Research Center, and John R. Hickey, 

The Eppley Laboratory, Inc. .......................... .. ..... .. ........... 239 

Preliminary Results from the Flight of the Solar Array Module Plasma 

Interactions Experiment (SAMPlE) 

Dale C. Ferguson and G. Barry Hillard, NASA Lewis Research Center 247 

Photovoltaic Electric Power Applied to Unmanned Aerial Vehicles (UAV) 

Jack Geis, Wright Patterson Air Force Base, and Jack H. Arnold, Rockwell 

International ................... . .............. .. ... ..................... 257 

Post-Flight Investigation Programmes of Recently Retrieved Solar Generators 

Lothar Gerlach, European Space Agency-ESTEC (XPG) .................... 269 

SESSION V-BLANKET AND ARRAY DEVELOPMENT 

High-Performance, Flexible, Deployable Array Development for Space 

Applications 

Russell N. Gehling, Joseph H. Armstrong, and Mohan S. Misra, Martin 

Marietta Astronautics ....... .. .... . .......... ... ... ........ . . ............ 287 

Static Stability Of A Three-Dimensional Space Truss 

John F. Shaker, NASA Lewis Research Center . ........ .................... 299 

An Update on the Development of a Line-Focus Refractive Concentrator Array 

Michael F. Piszczor, NASA Lewis Research Center, Mark J. O'Neill, 

Entech, Inc., and Lewis M. Fraas, JX Crystals. . . . . . . . . . . . . . . . . . . . . . . . . . . . . .. 313 

An Introduction to the Astro Edge Solar Array 

B.R. Spence and G.W. Marks, Astro Aerospace Corporation ............ . .... 323 

vii 

_I 
SESSION VI-NON-SOLAR ENERGY CONVERSION 

Rare Earth Garnet Selective Emitter 

Roland A. Lowe, Kent State University, and Donald L. Chubb, Serene C. 

Farmer, and Brian S. Good, NASA Lewis Research Center ..... .. ......... . . . 335 

InGaAs PV Device Development for TPV Power Systems 

David M. Wilt, Navid S. Fatemi, and Richard W. Hoffman, Jr., Essential 

Research, Inc.; Phillip P. Jenkins, David J. Brinker, and David Scheiman, 

NYMA, Inc.; Roland Lowe, Kent State University; and Donald Chubb, 

NASA Lewis Research Center .. . . .. . . ... .. .. . . ... . . . . . . . . . . . . . . . . . . . . . . . . 341 

Pulsed Laser Illumination of Photovoltaic Cells 

Jane A. Yater, NASA Lewis Research Center, Roland Lowe, Kent State 

University, and Philip Jenkins and Geoffrey A. Landis, NYMA, Inc. ... ......... 353 

PIN In(AI) GaAs Multijunction Laser Power Converters 

Steven Wojtczuk and Themis Parodos, Spire Corporation, and Gilbert 

Walker, NASA Langley Research Center .... . . .. .. . .. . . . ... . .. . ..... ...... 363 

High Efficiency GaP Power Conversion for Betavoltaic Applications 

Paul E. Sims, Louis C. DiNetta, and Allen M. Barnett, AstroPower, Inc. . .. ..... . 373 

SESSION VII-CELUMATERIAL PROPERTIES 

Review of Solar Cell Temperature Coefficients for Space 

Geoffrey A. Landis, NYMA, Inc .. ....... . . . . ....... .. . .. ... . .... .. . ........ 385 

Effects of Plasma Hydrogenation on Trapping Properties of Dislocations in 

Heteroepitaxial InP/GaAs 

S.A. Ringel and B. Chatterjee, The Ohio State University . .. ... .. .. . .... . . .... 401 

Nucleation, Growth, and Strain Relaxation of Lattice~Mismatched III-V 

Semiconductor Epitaxial Layers 

R.E. Welser and L.J. Guido, Yale University . . ... . . . . .. . .... . ..... . .. .. . .... 409 

WORKSHOP SUMMARIES 

Summary of Workshop on InP: Status and Prospects 

R.J. Walters, U.S. Naval Research Laboratory, and I. Weinberg, NASA 

Lewis Research Center .. .... . . . . ...... .. . . .... .. . . . . ....... .. . . ...... . . . 419 

Systems, Arrays and Applications Workshop 1 

G.J. Pack, Lockheed Missiles and Space Company, and Lothar Gerlach, 

ESAIESTEC . ... . ... . . . . . . . .. .. . . .... . . . . . . .. . ... . ....... . .... . .. . . .. ... 423 

Future Directions In PV Cell Development: Summary of the Workshop at 

SPRAT-XIII 

Geoffrey A. Landis, NYMA, Inc. . .. .. . . . . ........ .. . . .... . .. ... . .. . .... . .. . 425 

Array Technologies Workshop I 

Douglas Allen, W.J . Schafer Associates, and Andrew Meulenberg, Comsat 

Laboratories .. . . .... . . .. . . . . . . . . . . . . . .. ...... ... . .. .. . ...... . . . ...... . . . 431 

viii 
Workshop Summary 

Thermophotovoltaics and Non-Solar Energy Conversion 

Donald L. Chubb, NASA Lewis Research Center, and Larry C. Olsen, 

Washington State University ............................................. 437 

ix 

l --- ------------
N95-20503 

GODDARD SPACE FLIGHT CENTER SOLAR ARRAY MISSIONS, 397;;)-0 

REQUIREMENTS AND DIRECTIONS , 2 { ) t 

Edward Gaddy and John Day 

NASA Goddard Space Flight Center 

Greenbelt, Maryland 

ABSTRACT 

The Goddard Space Flight Center (GSFC) develops and operates a wide variety of spacecraft for 

conducting NASA's communications, space science, and earth science missions. Some are "in 

house" spacecraft for which the GSFC builds the spacecraft and performs all solar array design, 

analysis, integration, and test. Others are "out of house" spacecraft for which an aerospace 

contpctor builds the spacecraft and develops the solar array under direction from GSFC. The 

experience of developing flight solar arrays for numerous GSFC "in house" and "out of house" 

spacecraft has resulted in an understanding of solar array requirements for many different 

applications. This presentation will review those solar array requirements that are common to 

most GSFC spacecraft. Solar array technologies will be discussed that are currently under 

development and that could be useful to future GSFC spacecraft. 

BACKGROUND 

The GSFC both purchases and manufactures spacecraft. For the purchased spacecraft, GSFC 

supplies a spacecraft specification to a manufacturer who then purchases or fabricates the array. 

The spacecraft specification usually has in it a general specification which covers the power 

system and the solar array. Such spacecraft under development by the GSFC are: Total Ozone 

Mapping Spectrometer (TOMS), National Oceanic and Atmospheric Administration Spacecraft 

(NOAA)-J, NOAA-K, NOAA-L, NOAA-M, NOAA-N, and NOAA-N prime, Global Geospace 

Science (GGS)-WIND, GGS-POLAR, Geostationary Operational Environmental Satellite 

(GOES)-J, GOES-K, GOES-L, Tracking and Data Relay Satellite (TDRS) F-7, Landsat-7, 

Advanced Composition Explorer (ACE), Earth Observing Spacecraft (EOS)-AM, EOS-PM, 

SPEC, CHEM, TDRS H, I J, and Hubble Space Telescope (HST) Servicing Mission 

Replacement Array. Some of the characteristics of the arrays on these spacecraft is summarized 

in Table I. 

The GSFC also manufactures some·spacecraft. In these cases, GSFC and specifically the Space 

Power Applications Branch develops and purchases solar arrays. GSFC has a greater influence 

on the specifics of these solar arrays as opposed to the arrays on the out of house spacecraft. 

These arrays have provided us with knowledge of the requirements for spacecraft solar cells and 

some insight into what technologies will be most helpful for the future. Spacecraft in this group 

include the Small Explorer (SMEX)- 2 also known as the Fast Auroral Snapshot (FAST), the 

X-Ray Timing Explorer (XTE), SMEX -3 also known as the Submillimeter Wave Astronomy 

Satellite (SW AS), the Tropical Rainfall Measuring Mission (TRMM), SMEX 4, 5, and 6; and the 

Far Ultraviolet Spectroscopic Explorer (FUSE). Table IT summarizes some of the array 

characteristics of the spacecraft in tJ:ris category. 

The GSFC also operates spacecraft. Some of the spacecraft the center currently operates include: 

the International Ultraviolet Explorer (IUE), the International Sun Earth Explorer subsequently 

renamed the International Cometary Explorer (ICE), the Earth Resource Budget Spacecraft 

(ERBS), the Cosmic Origins Background Explorer (COBE), the Hubble Space Telescope 

(HST), the Gamma Ray Observatory (GRO), the Upper Atmosphere Research Satellite (UARS), 

the Extreme Ultraviolet Explorer (EUVE), and the Solar Magnetospheric Particle Explorer 

(SAMPEX).Table ill summarizes some of the array characteristics of the spacecraft arrays in this 

category. 
Table I 

GSFC Out of House Solar An·ays Currently in Development 

Life Array Cell Array/Cell 

SIC Launch (yrs) Orbit Type Type Manufacturer 

TOMS 1994 3 LEO rigid, deployable BSFR, TRW, ASEC 

silicon 

NOAA-J 1994 2 LEO ligid deployable, BSR, MM,ASEC 

tracking silicon 

GGS 1994 3 sun earth rigid, body mount, BSFR, MM,ASEC 

WIND libration electrostatically clean silicon 

GGS 1995 3 elliptical rigid, body mount, BSFR, MM,ASEC 

POLAR electrostatically clean silicon 

GOES-J 1995 5 GEO rigid, deployable, BSR, Loral, 

tracking silicon Spectrolab 

TDRS-7 1995 10 GEO rigid, deployable, BSR, TRW, ASEC 

tracking silicon 

Landsat 1998 5 LEO rigid, deployable, BSR, MM,TBD 

-7 tracking silicon 

ACE 1997 2 sun earth rigid, deployable BSFR, APL, Spectrolab 

libration silicon 

EOS- 1998 5 LEO flexible, deployable, 5.5 mil MM,TRW, 

AM tracking GaAs/Ge ASEC 

NOAA 1996- 2 LEO rigid, deployable, BSR, MM,TBD 

K,L,M 1999 tracking silicon 

HST 1999 5 LEO TED TED TED 

Servicing 

EOS-PM 1998- TED TED TED TED TBD 

SPEC 2003 

CHEM 

TORS 1999- TED TBD TED TED TBD 

H, I, J 2001 

2 
Tablell 

GSFC In House Solar Arrays Currently in Development 

Life Orbit Array Cell Array/Cell 

SIC Launch Cyrs) Clan) Type Type Manufacturer 

FAST 1994 1 350 x 4200 rigid, body mount, GaAs/Ge TRW,ASEC 

faraday cage 

XTE 1995 2 600 rigid, deployable, BSFR, Spectrolab 

tracking silicon 

SWAS 1995 3 600 rigid, deployable GaAs/Ge Spectrolab 

TRMM 1997 3 350 rigid, deployable, GaAs/Ge TRW,ASEC 

tracking 

SMEX 1996 - TED TED TED TED TED 

4,5, 6 1999 

FUSE 2000 TED TED TED TED TED 

REOUIREMENTS FOR SOLAR CELLS 

A solar cell of a given type must have undergone the following tests before we will consider it 

for use on a in house spacecraft solar array. Except as noted below, existing silicon and gallium 

arsenide and gallium arsenide on germanium solar cells meet these requirements almost 

flawlessly. These requirements are frequently not explicitly stated by GSFC specifications, but in 

one way or another they are present. 

1. The solar cell's beginning of life current versus voltage characteristics must be determined as a 

function of temperatures from -SOC to +SOC for low earth orbits. Ideally, this range is extended 

to -lS0C to +SOC to cover geosynchronous orbits. 

2. The solar cell's current versus voltage characteristics must be determined as a function of 

amount and kind of hard particle radiation. In other words, the solar cell must be exposed to 

electrons and protons of varying energies and of varying amounts in ground tests to determine 

suitability for use on spacecraft. This exposure must generally be thorough enough so that the 

radiation in space can be converted into an equivalent number of 1 Mev electrons and to 

determine the solar cell's end of life current versus voltage characteristics as a function of 

temperature from -l00C to +100C. 

3. The solar cell's performance must be determined as a function of its degradation due to the 

exposure to sunlight, particularly the ultraviolet component. The equivalent of a one year's 

exposure on an accelerated basis is considered to be satisfactory. 

3 
Table ill 

In - Orbit Spacecraft Currently Operated by GSFC 

Orbit Altitude Solar Cell Array/Cell 

Spacecraft Launched (km)and Solar Array Type Type Manufacturer 

Inclination (0) 

IDE 1/78 30,210 x rigid, deployable, BSR, silicon GSFC,ESA 

41,343, 33.8° [lXed AEG 

ISEE-3 8/78 sun earth rigid, body BSFR, silicon GSFC, 

later known libration point, mounted Spectrolab 

as ICE followed by 

lunar swing by 

and cometary 

encounter 

ERBS 11184 590, 57 rigid, fixed Ball, 

Spectrolab 

COBE 11/89 873 x 885,99 rigid, deployable, BSFR, silicon GSFC, 

[lXed Solarex 

HST 4/90 600,28.5 flexible, roll out, BSFR, silicon Marshall, 

tracking ESA, BAe, 

AEG 

GRO 4/91 370,28.5 rigid, deployable, BSR, silicon TRW,ASEC 

tracking 

UARS 9/91 570, 57 rigid, deployable, BSFR, silicon GE,ASEC 

tracking 

EUVE 6/92 520,28.4 rigid, deployable, BSFR, silicon GSFC, FSC 

Solarex 

SAMPEX 7/92 509 x 679,82 rigid, deployable GaAs/Ge GSFC, FSC 

fixed Spectrolab 

4 
4. The solar cell's perionnance must be determined after exposure to thennal cycling. Generally, 

a solar cell in a low earth orbit will be exposed to on the order of 5,500 cycles from -80C to 

+80C each year. Most GSFC spacecraft have a lifetime of two years or more, and the cell must 

be tested to such a flight environment. Further, the test showing competence in this area should 

have the solar cell fixed to a panel as this is the condition under which the solar cell must actually 

perionn. The condition of making electrical connection to the solar cell and mounting it to a 

substrate makes this test significantly more strenuous than just cycling the solar cell. If the solar 

cell is to be used in a geosynchronous orbit, it must be able to withstand very roughly eighty 

cycles per year as extreme as -18OC to +80C. 

5. The solar cell must not degrade due to the humidity in an air-conditioned room over many 

years. The solar cell must not degrade due to humidity exposer in an unconditioned atmosphere 

for several days, such as at launch and during shipment. The solar cell's resistance to humidity is 

traditionally proven by exposing the solar cells to 90% relative humidity at 45C for 30 days with 

the requirement that the solar cell not degrade more than 2% in peak power. This test is to some 

degree arbitrary. Exactly how well a cell must do in the test to show that it perionns well under 

the conditions of the real world is not well detennined. Consequently, this test could be 

weakened if it a cell manufacturer could show that it made an unduly pessimistic prediction for a 

new type of solar cell. 

6. The solar cell's absorptance must be measured. 

7. The solar cell's weight density must be detennined. 

8. The GSFC occasionally has missions with magnetic specifications that require no magnetic 

materials be used in the fabrication of the solar cell. 

9. A darkened solar cell must be able to withstand reverse bias to approximately 10 percent more 

current than its short circuit current or to a voltage, typically around 50 volts, that is detennined 

by a combination of the power system and array perfonnance. Whichever of the requirements is 

least severe governs. This requirement is not met flawlessly by gallium arsenide or gallium 

arsenide on germanium solar cells unless they are first screened. 

10. Although GSFC frequently does not specify cell size, sizes under 2 cm by 4 cm are not 

practical due to laydown cost. This requirement is of course flexible if the benefits of a small 

solar cell can be shown to outweigh the cost penalty. For example, we would gladly fly 50% 

efficient 2 cm by 2 cm solar cells. 

11. It must be possible to fix an interconnect strongly enough to a cell so that it can take small 

bumps and thermal cycling without coming off. To prove this, the contacts on the solar cell must 

withstand a 1.5N pull test before and after being exposed to humidity. As in the case of the 

humidity test, this requirement is somewhat arbitrary and can be modified if it can be shown that 

it is too severe for a paIticular solar cell. 

12. A completely new type of solar cell must be flown on a balloon to determine its output. 

13. Very likely. a completely new type of solar cell would have to be flown on a limited basis in 

space before using it as a primary source of power. 

15. Any organization manufacturing solar cells must have a significant quality assurance effort 

and be financially stable. In tenns of quality, this means that the organization should meet or 

come close to meeting the requirements of MIL-Q-9858. 

5 
REQUIREMENTS FOR GSFC IN HOUSE SOLAR ARRAYS 

In addition to the above for solar cells, the following requirements for solar arrays apply. These 

requirements are explicitly stated in GSFC specifications. 

The array must meet configuration, maximum weight, minimum beginning of life power, 

insulation resistance, cleanliness, outgassing, mechanical, and miscellaneous requirements such 

as temperature sensor accuracy. The array must also meet specifications on resistance to 

accidental damage, resistance to damage by storage temperature and humidity, and resistance to: 

ultraviolet radiation, atomic oxygen, micrometeroids and space debris and shadowing. 

The array's performance after thermal cycling must be proven in a life test which includes 

samples of every component which will be mounted on the flight arrays. Although solar arrays 

have been manufactured for many years, this requirement is still frequently not met flawlessly. 

The array's performance test must be proven after exposure to vibration or acoustic. This test is 

generally met without difficulties. 

The flight array must be acceptance tested by exposure to eight thermal vacuum cycles and 

exposure to acoustic. The thermal vacuum cycles again are frequently not flawless. 

We occasionally require that no magnetic materials be used on the solar array and/or that the array 

be electrostatically clean. This usually means that the upper surface of the array be conductive. 

The requirement is sometimes tightened to the extent that virtually every surface on the array be 

conductive. 

GALLIUM ARSENIDE VERSUS SILICON 

Table IV summarizes the array characteristics of the spacecraft for which GSFC or its contractors 

did a trade off between GaAs solar cells and silicon solar cells. The GaAs arrays cost 

approximately 70% more on a per watt basis than silicon solar arrays, but because of the benefits 

they provide, their system level cost is actually lower than silicon. For each spacecraft array in 

Table IV, but with emphasis on TRMM, we summarize below the factors used in determining 

whether to use GaAs or silicon. 

Table IV 

Summary of GSFC Solar AITays with a GaAs versus Silicon Tradeoff 

Array Cell Equivalent Array 

Area Lifetime Altitude 1 Mev Delivery 

Spacecraft Cell Type (m2) (yrs) (km) Electrons Date 

SAMPEX GaAs/Ge 1.7 3 450 x 830 l.l x 1013 Launched 

XTE Silicon 15.5 2 600 4.6 x 1012 Aug. 94 

FAST GaAs/Ge 2.6 1 350 x 4200 l.5 x 1014 May 94 

1RMM GaAslGe 18.1 3 350 l.7 x 1012 May 95 

SWAS GaAs/Ge 3.6 3 600 9.6 X 1012 Sept. 94 

EOS GaAs/Ge 35 5 705 5.4 x 1013 Feb. 96 

6 
SAMPEX 

The solar array for SAMPEX is very small. A silicon array could not meet the power 

requirements for the spacecraft. This solar array and its cells were supplied by Spectrolab. 

XTE 

The solar array for the X-Ray Timing Explorer is fully designed and a contract has been let to 

Spectrolab for its fabrication. The solar array consists of silicon solar cells on an aluminum face 

sheet honey comb core substrate. Silicon solar cells were selected for this spacecraft primarily 

because of the difficulties in obtaining accurate prices for gallium arsenide solar arrays. We had 

limited pricing experience with the cells and most of the panel manufacturers had limited 

experience working with the cells and hence a high price uncertainty. Had the array used gallium 

arsenide solar cells, it would have been the first GSFC spacecraft array of a moderate size, 

approximately 2,000 watts, to do so. The advantage of the gallium arsenide solar cells was that 

their use would have prevented the necessity of a tracking solar array which would have 

increased spacecraft reliability and removed the substantial costs associated with the tracking 

mechanisms. 

FAST 

The solar array for FAST has been fabricated through a contract with TRW. The solar array for 

the Fast Auroral Snapshot satellite is gallium arsenide. This array is area limited and the silicon 

solar array of the required size could not supply the needed power. The cells are on an aluminum 

face sheet honeycomb core that forms the outside of the spacecraft body. 

As an aside, this array is particularly interesting in that it has no magnetic materials, has magnetic 

compensation wiring directly under the solar cells and has a Faraday cage over its entire surface. 

The Faraday cage primarily consists of covers with conductive indium oxide coatings. The 

covers are interconnected in such a way that interconnects completely cover any insulating area 

on the array. 

TRMM 

The solar array for the Tropical Rainfall Measuring Mission is fully designed and a contract has 

been let to TRW for its fabrication. The solar array consists of gallium arsenide/germanium solar 

cells on an aluminum face sheet honeycomb core substrate. Both gallium arsenide and silicon 

solar cells were considered for the TRMM array. Table V provides a comparison of the resulting 

arrays. In Table V the delay actuator is a device which prohibits fouling by the premature 

deployment of one panel prior to deployment of another panel. The potentiometers are used to 

monitor the deployment of the various panels. 

The silicon solar cell array is, from Table V, approximately 45% larger than the gallium arsenide 

array, a figure which G. C. Datum and S. Billets have also reported. l The smaller area reduced 

the spacecraft's fuel consumption and increased the probability of meeting a three year life. This 

was a particularly important consideration. Table IT also shows that silicon array is 36% heavier. 

The areal and weight advantage of the gallium arsenide array results because the gallium arsenide 

solar cells offer approximately 40% more power on a per area basis at operating temperature. The 

GaAs solar cells greatly simplified the deployment of the solar array. This is important because 

deployables are historically among the less reliable components of spacecraft. Further, the array 

deployment would have to be tested on the ground and making a g-negation mechanism to allow 

the TRMM silicon array to deploy would have been difficult almost to the point of impracticality 

as each array wing would have consisted of six hinged panels. 

7 
Table V 

TRMM Silicon versus GaAs Technical Factors Comparison 

Parameter Si GaAs 

Weight of Cell Stack, Wiring, Connectors 

and Miscellaneous 48 kg 47 kg 

Array Area 26.2 m2 18 m2 

Array Operating Temperature 74C 87C 

BOL Efficiency @ Operating Temperature 11.3% 15.8% 

EOL Efficiency 9.4% 13.3 

Number of Individual Panels 12 4 

Number of Panel Hinges 20 4 

Number of Delay Actuators 2 0 

Number of Potentiometers 12 2 

Mechanical System Weight 144 kg 94 kg 

Total Weight 192 kg 141 kg 

SWAS 

The solar array for the Submillimeter Wave Astronomy Satellite is gallium arsenide. This array is 

area limited and a silicon solar array of the required size could not supply the needed power. The 

solar array for SW AS is fully designed and a contract has been let to Spectrolab for its 

fabrication. 

EOS 

Although, most of the conclusions we draw below follow from our experience with in house 

arrays, we here mention our most technologically advanced array which is for the Earth 

Observing System, an out of house project. The EOS carries a flexible deployable array powered 

with 5.5 mil thick gallium arsenide solar cells. This represents the first such use of these cells on 

a flexible array. The trade which drove this array to the gallium arsenide solar cells was that the 

array is on one side of the spacecraft and tended to rotate the spacecraft in flight. Using the 

gallium arsenide array thereby enabled the attitude control system to use existing reaction wheels 

rather than developing new ones. 

TRENDS 

The gallium arsenide solar cell offers a substantial improvement over silicon. Silicon solar arrays 

are generally on the order of 40% larger when the operating temperature of the array is taken into 

account. In most cases, the gallium arsenide solar cell offers a dramatic weight reduction 

8 
compared to silicon even though the gallium arsenide cells are heavier than silicon. This is 

because reducing the array area reduces the size of the substrate, harnessing, number of covers, 

amount of adhesive etc. This reduction more than compensates the weight difference between the 

solar cell types. This is true in the overwhelming number of spacecraft solar arrays because they 

use aluminum face sheet over aluminum honeycomb core substrates. As the weight of the 

substrate decreases, the weight advantage of gallium arsenide solar cells becomes less. It is only 

with extremely lightweight solar arrays that gallium arsenide on germanium solar cells result in 

an array which approaches the weight of a silicon array for the same power. This happens on 

lightweight deployable solar arrays using the thinnest commercially available silicon, 55 microns 

thick, compared to the same array with gallium arsenide cells, 115 microns thick. 2• Because 

these arrays represent advanced technology they do receive a great deal of attention in the 

literature, however there are only a handful of them flying and therefore their practical effect on 

cell technology is limited. 

The advantages of the GaAs solar cell, which derive primarily from its greater efficiency, suggest 

that spacecraft solar cells for future spacecraft will be driven primarily by a greater power 

density. With this statement in mind, the authors believe that the following solar cells offer the 

greatest opportunity to improve spacecraft performance. 

MUL TI JUNCTION SOLAR CELLS 

The authors believe that the qualification and development for production of these cells would 

provide the greatest benefit at the least cost in the shortest time. Such solar cells have been 

produced in the laboratory3,4,5 with AMO efficiencies of over 25%. If these solar cells are mass 

produced at 24 percent efficiency they represent over a 30 percent improvement in state of the art 

gallium arsenide solar cells. These cells will therefore multiply the considerable advantages the 

gallium arsenide cells have provided. Additionally, the method of manufacture of these solar cells 

suggests that the price will be competitive to gallium arsenide on germanium. This statement is 

based on the assumption that making the cascade cell primarily entails leaving it in the reactor 

which grows the various cell layers somewhat longer than is required for the gallium arsenide on 

germanium solar cell. During this time, the reactor will automatically control the flow of gasses 

to grow the additional layers. The time and labor involved in this additional processing is 

probably minimal. There will also be an additional yield loss. 

Because of the improved power density is so welcome, because the solar cell is in many ways 

similar to gallium arsenide solar cells, because the cell is probably not significantly more 

expensive and because several organizations have successfully produced versions of the cell, we 

believe that the next most probable step in improving array performance is with the multi-junction 

solar cell. 

IMPROVED GALLIUM ARSENIDE SOLAR CELLS 

Gallium arsenide solar cells have been fabricated with air mass zero efficiencies in excess of 21 

percent. 6 These solar cells offer a significant improvement to spacecraft power systems using 

improved versions of existing solar cells. These are therefore a very valuable asset, if they can be 

put into production. 

INDIUM PHOSPHIDE SOLAR CELLS 

These solar cells have been fabricated with air mass zero efficiencies of over 19 percent.7 The 

effect of radiation on these solar cells is significantly less than that for gallium arsenide cells of 

approximately the same efficiency. Tobin reports that cells of this approximate efficiency degrade 

4.7% after irradiation with 1014 1 Mev electrons. GaAs cells degrade 9% after the same 

9 

__ J 
irradiation. 8 At this radiation level, these cells will therefore show about a 4% advantage over 

gallium arsenide solar cells, provided both cells have the same initial efficiency. At 1015 1 Mev 

electrons, GaAs solar cells have a degradation of 26%. At these radiation levels, and presuming 

the InP cells retain an approximate 2 to 1 advantage over the GaAs cells in degradation means 

that the indium phosphide solar cell will have about a 13% advantage over the GaAs solar cell. 

However, the advanced GaAs solar cell has about a 10% advantage over the InP cell at beginning 

of life and the multijunction cell has an approximate 30% advantage over the indium phosphide 

cell at beginning of life. This means that only under the most extreme conditions of radiation will 

the InP cell show an advantage over an advanced gallium arsenide cell and that it will never show 

an advantage over a multi junction celL As can be seen from Table I which is typical for most of 

our spacecraft, the radiation damage is generally under 1014 equivalent 1 Mev electrons. 

ADV ANCED SILICON SOLAR CELLS 

A variety of high efficiency silicon solar cells have been developed.9 ,10 ,11,12 These cells may 

find application in space as competitors to GaAs and production silicon solar cells. To do this, 

they must be tested to the cell requirements mentioned earlier, particularly because the cells may 

be quite sensitive to radiation. Even if they are resistant to radiation or can be made so, the 

authors believe that they do not overcome the advantage of the higher efficiency multi-junction 

solar cells even considering the greater expense of the later technology. 

THIN FILM SOLAR CELLS 

Thin film solar cells offer the advantage of an enormous power to weight gains over any of the 

solar cells discussed previously. To utilize this advantage a mechanism must be developed to 

deploy them and this mechanism must also be light enough to not cancel the cell's inherent 

advantage. These cells are at a disadvantage at lower altitudes because of their larger area, which 

is almost twice that of state of the art GaAs arrays. At these altitudes the large area creates an 

adverse impact on the spacecraft attitude control system and on the ability of the spacecraft to 

maintain altitude. 

CONCENTRATOR SOLAR CELLS 

The authors believe that concentrator solar cells are not useful to NASA spacecraft. This is 

because they have a significantly detrimental effect on spacecraft reliability. If non concentrating 

solar cells are used on a spacecraft, the spacecraft can lose its ability to drive the arrays or it can 

tumble for many hours and still be recovered. This is because arrays will supply about 30% of 

their rated power in a random spacecraft tumble, about enough to keep a powered down 

spacecraft going indefinitely. If concentrator arrays are used, the arrays will supply only small 

fraction of their rated power in a random spacecraft tumble. Under these circumstances the 

typical spacecraft batteries will discharge after about four hours or three orbits. In short, the 

concentrator arrays impose very strict requirements on the short term pointing reliability of the 

spacecraft attitude control system and the solar array drive. These requirements would be very 

difficult to convincingly achieve. The likelihood of ever using these cells is therefore small. 

FLEXIBLE SOLAR ARRAYS 

The development of flexible, deployable solar array such as the Advanced Photovoltaic Solar 

Array, (APSA)13 , the Solar Array Flight Experiment (SAFE) and the Flexible Rolled Up Solar 

AlTay (FRUSA) have enabled NASA to enhance the capability of two of its larger spacecraft, 

namely EOS and HST. From the user's point of view these deployable arrays enable a large 

alTaY to be packed in a small volume on the spacecraft. Unfortunately these arrays are very 

difficult to analyze mechanically, particularly with respect to the effect they have on the 

10 
spacecraft's attitude control system, they are mechanically complex and they are impossible to 

end to end test. For these reasons, we avoid them until their advantages become very large. 

However, as we gain experience with them, the authors believe that they will become more 

popular. 

The weight advantage of these arrays is substantial, although not as great as the prototype arrays 

suggest. For example, the APSA has a power to weight ratio of 130W/kg. The EOS array, that 

derived from APSA but had to overcome various practical constraints, has a power to weight 

ratio of 32 W /kg. As an aside, this decrease is not due to the change from the thin silicon cells on 

APSA to the GaAs cells on EOS, if the EOS array had weightless cells and covers, it would have 

a power to weight ratio of only 43 W/kg. For reference the TRMM array has a power to weight 

ratio of 20W/kg, typical of many spacecraft. 

FLIGHT TESTING ADVANCED SOLAR CELLS 

To insure that advanced solar cells are flown as soon as practicable, it is necessary to fly them on 

spacecraft as soon as ground testing indicates that they are a promising candidate but before 

ground testing has fully qualified them. This is because even when solar cells are completely 

qualified through ground testing, many projects are reluctant to fly them unless they have flown 

before. The flight of small numbers of these cells on otherwise conventional arrays will provide 

experience to the manufacturers, will gain the confidence of spacecraft managers, and will 

complement ground based qualification. Such use of advanced solar cells will not generally 

enhance a given project's capability to meet its requirements and will therefore be resisted. 

However, the price to be paid by the project is small, provided only minimal telemetry is 

specified, and the benefits to the space program are large particularly in view of the several very 

promising advanced solar cells. The Space Power Branch at GSFC is therefore recommending 

that inexpensive low risk flight experiments be undertaken. 

CONCL USIONS 

The GSFC has responsibility for a large number of solar arrays, some of which are powered by 

GaAs solar cells. These gallium arsenide solar cells have provided a wide variety of benefits 

including the preservation of spacecraft fuel, the enhancement of missions that would be severely 

power limited without them, the simplification of array deployment mechanisms, and the 

reduction of solar array weight. Because of the benefits these cells have provided are so useful, 

we believe that the further test and development of high efficiency solar cells, particularly multijunction solar cells will further increase these already substantial benefits. These cells will 

provide additional power at a modest increase in price. We are recommending that the GSFC and 

other agencies start flying a small percentage of each of their state of the art arrays with advance 

solar cells so that experience can be gained with these cells even before they are fully qualified to 

be the primary source of power for spacecraft. 

We have made the point that some areas of research seem to us to be less useful. In particular, 

the future development of concentrator solar arrays does not appear to be a fruitful avenue to 

pursue, at least for NASA spacecraft. 

1G. C. Datum et at, "Gallium Arsenide Solar Arrays - A Mature Technology," Proceedings of the 22nd IEEE 

Photovoltaic Specialists Conference, October 7, 1991, Las Vegas, NV p. 1422. 

2p. M. Stella et at., "Thin Film GaAs for Space - - Moving Out of the Laboratory" Proceedings of the 23rd IEEE 

Photovoltaic Specialists Conference, May 10, 1993, Louisville, KY. Figure 1, p. 22. 

11 

-------
3B. C. Chung et al., "25.2%-Efficiency (I-Sun Air Mass 0) AlGaAs/GaAslInGaAsP Three Junction Solar Cell," 

Proceedings of the 22nd IEEE Photovoltaic Specialists Conference, October 7, 1991, Las Vegas, NV p. 54. 

4B. C. Chung et al., "27.6% (1 Sun Air Mass l.5) Monolithic Two Junction AlGaAs/GaAs Solar Cell and 

25%(1 Sun, Air Mass 0) Three Junction AlGaAs/GaAs/InGaAs Cascade Solar Cell," Proceedings of the 21st IEEE 

Photovoltaic Specialists Conference, May 21 1990, Kissimimee, FL, p. 179. 

5J. M. Olson et al., "A 27.3% efficient Gao.5InO.sP/GaAs tandem solar cell," Appl. Phys. Lett. 56(7) 12 

February 1990, p. 623. 

6S.p. Tobin, et al., "Advances in High Efficiency GaAs Solar Cells," Proceedings of the 21st IEEE Photovoltaic 

Specialists Conference, May 21 1990, Kissimimee, FL, p. 158. 

7c. J. Keavney, et al., "Emitter Structures in MOCVD INP Solar Cells," Proceedings of the 21st IEEE 

Photovoltaic Specialists Conference, May 21, 1990, Kissimimee, FL, p. 14l. 

8B. E. Anspaugh, Addendum 1 to the Solar Cell Radiation Handbook," 1989, p. 37. 

9M. A. Green, et al., "Progress in High Efficiency Silicon Cell and Module Research," Proceedings of the 23rd 

IEEE Photovoltaic Specialists Conference, May 10, 1993, Louisville, KY, p.8. 

lOR. A. Sinton, et al., "Large Area 21 % Efficient Si Solar Cells," Proceedings of the 23rd IEEE Photovoltaic 

Specialists Conference, May 10, 1993, Louisville, KY, p.157. 

11 J. Knobloch, et al., "High-Efficiency Solar Cells from FZ, Ca, and MC Silicon Material," Proceedings of the 

23rd IEEE Photovoltaic Specialists Conference, May 10, 1993, Louisville, KY, p.27l. 

12V. Garboushian, et al., "Radiation Hardened High Efficiency Silicon Space Solar Cell," Proceedings of the 23rd 

IEEE Photovoltaic Specialists Conference, May 10, 1993, Louisville, KY, p.1358. 

l3p. M. Stella et al., "Thin Film GaAs for Space - - Moving Out of the Laboratory," Proceedings of the 23rd 

IEEE Photovoltaic Specialists Conference, May 10, 1993, Louisville, KY. Figure 1, p. 22. 

12 

__ J 
BMDO PHOTOVOLTAICS PROGRAM OVERVIEW 

Leonard H. Caveny 

Ballistic Missile Defense Organization 

Washington, DC 

and 

Douglas M. Allen 

W.J. Schafer Associates 

North Olmsted, Ohio 

Outline 

BMDO advanced solar array program overview 

Brilliant Eyes type satellites 

Electric propulsion 

BMDO photovoltaic programs 

Why concentrators? 

loffee concentrator and cell development 

Entech linear mini-dome concentrators 

Flight test update/plans 

Conclusions 

BMDO Advanced Solar Array 

Program Overview 

N95-20504 

Primary need is for a better power source for Brilliant Eyes type 

satellites 

Lightweight and low cost systems are also needed to enhance 

BMDO electric propulsion programs 

Cooperation with other government agencies is important to address 

broader needs and leverage capabilities 

• NASA Lewis Research Center 

• Phillips Laboratory 

• Others 

Commercial spin-ofts to improve U.S. competitiveness in the global 

marketplace are also important 

13 
Brilliant Eyes Program Overview 

BMDO Program for a low orbit constellation of satellites for surveillance 

and tracking of ballistic missiles for theater and national defense 

missions 

Program is managed by the Air Force Space and Missiles Center for 

BMDO 

Competition ongoing to select prime contractor 

• Rockwell and TRW selected from four original contractors last 

year 

• Source selection in progress to select one contractor to build the 

first two satellites 

• Both contractor's programs will be continued for a later 

downselect for the full satellite constellation 

First flight -1997 

Brilliant Eyes Type System 

Constellation Specs 

Altitude <2000 km, high inclination & low inclination rings 

In lower end of VanAllen radiation belts 

Constellation of 10 to 40 satellites 

Deployed on an MLV, multiple satellites per launch 

EmphaSis on achieving mission objectives at minimum cost 

Conservative approach to non-critical satellite components 

14 
500 kg goal 

Brilliant Eyes Type System 

Spacecraft Specs 

1 kW solar arrays for 500 W of orbital average power onboard 

(end of life) 

5-7+ year lifetime 

Rad hard parts required 

Applications of Electric Propulsion 

Initial orbit positioning - primary application, major fuel savings 

at a cost of weeks extra time for the deployment 

• Extra satellitesllaunch vehicle (major $ savings) 

or 

• Extra margin in launch mass (risk reduction) 

Strategic maneuvers and repositioning 

Orbital maintenance 

Deorbit satellite at end of life 

Primary Drawback of EP is longer trip times for positioning and maneuvers -

, More solar array power (if affordable and mass efficient) helps a lot 

15 
750 

BRILLIANT EYES 

EARTH ORBIT TRANSFER 

A Case For High Specific Impulse And Efficiency - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - - '2--:;os I I e I I o-x.\{\ I 

B , . f L'f I \\'jC

f 

0) I 

.:::t. 

~7 0 

eg lnntng 0 1!3: I .,.. I 

- Power-= l.-5- ·kW- -aver-age· - - - - - - - _.1 - - - - - / - - - - - - _.J I I '; I Payload=500 kg I I .... .... I 

Parki ng Orbit=4GO km .) .... I .... I I 

II) 

II) 

o 

2 

650 _________ ~ ____ L _ _________ -' __ .1 _____ - - ______ .J 319b 

1 ,; ,; 1 500S -'+-

o 

L 

() 

I ,; I j ~ rh I .... .... 1 "s ~f C I '].5~5s1 

Q) 600 () 

I .... I '].0 ~ 'I.WS '/...6 

- - - - - - - - - - - - - - L ..... ~ - - - - - - - - - - - .1 - - - - , 3-c0I 

.... ,;: 1 r\\.l9n:

es o 005 51:11> 

0-

(f) 

- 550 - o 

.-

c 

/..... '!.Q' dO'{s sa I 60 70 78 oi "f...e \ £> 

,0 ........ '].0 ..]I a AB Sr I 

- io cO,{ - - - - - - - ~o. - ~ - - - -8~ - - -£r - - - - ~ 

8 [ 

500+-~.- --~~--~-- --~--~-+-- -- --~~ 

500 1000 1500 2000 

Higher Earth Orbit, km 

200 400 600 800 

I 1 I 1 

Delta Velocity, m/s (For Electric Propulsion) 

ep\spl5DO.in epl5il.grf 24Jun92 LHCoveny 

BMDO Photovoltaics Programs 

Concentrator Solar Arrays 

• Entech linear refractive concentrator 

• loffee reflective and refractive concentrators and cells 

• NASA Lewis agent for program management 

Multiple-band-gap planar arrays 

• Phillips Laboratory agent for program management 

Other SBIR programs 

Flight tests to demonstrate technology maturity 

16 
---- --------- - ------

Why Concentrators? 

Concentrator solar arrays provide a pathway to major advances in 

satellite solar array parameters of interest 

• Cost 2X lower than planar silicon 

• Specific power of 100 W/kg or better - even for small, satellites 

• Very small penalty for incorporating radiation resistance 

Technology development over the last decade has resulted in 

concentrator designs that are practical to integrate onto satellites 

• Linear concentrators only require one critical axis for sun 

pointing 

• Sun acceptance angles have been increased by nearly an 

order of magnitude 

• Innovative optics designs allow low cost manufacturing 

approaches 

Solar Array Technology Comparison 

for Brilliant Eyes Type Satellites 

GaAs/Ge Planar Solar Array . 

~ Linear Concentrator ~ 

~ Solar Array ~ 

I 

(note: drawings to scale) 

Array Specs: 

1278 W (EOl) 

1406 W (BOl) 

4.5 m2 

18 kg 

Array Specs: 

1278 W (EOl) 

2131 W (SOL) 

12 m2 

42 kg 

$6.39M (est.) 

$O.84M (projected) 

17 
loffee Concentrator 

and Cell Development 

Mirrored and flat fresnel concentrator array development 

• Linear version of reflective concentrator now being 

developed 

• Four fresnel modules being prepared for flight test 

Advanced tandem and multi-junction solar cells 

• GaAsAnGaAs ~----------------~ 

• GaAs/AIGaAs 

Array goals are: 

• 300 W/m2 

• 100 W/kg 

• 30% efficiency 

Entech Linear Concentrator 

Technology Description 

Concentrator solar arrays to provide power 

to satellites 

Single axis convex fresnel concentrator concept 

Technology nearing maturity for satellites 

• Materials, cells, and lenses already space flight tested 

• 2 axis mini dome concentrator module being prepared for 

launch on PASP+ 

• One axis version of concentrator ground demonstrated 

Performance of arrays provides many operational benefits 

• Efficiency = 1.5 x GaAs 

• Recurring Cost = 0.3 to 0.5 x Si 

• Array specific power = 1.5 to 2.5 x APSA for small arrays 

• Radiation tolerance same as InP 

18 
10" 

1 

Flight Test Update/Plans 

Lear Jet concentrator module performance tests 

LDCE-4, -5: Materials experiment, limited AO exposure 

EOIM III: Small area lens materials and cells 

Wakeshield (MATLAB-1): Materials experiment and two mini 

dome lenses (AO performance testing of lens coatings) 

PASP-plus: First active experiment of advanced refractive 

concentrator array technology, long duration (1-3 yrs.), 

radiation damage, and high voltage plasma interaction 

STRV-2: In planning stages, first active experiment of one axis 

concentrator module 

Molniya flight experiment: 1995 test of four 16 element loffee 

fresnel concentrators 

C-SA VE: Planned 1997 flight demonstration of a 1 to 1.5 kW 

linear concentrator array 

STRV-2 Experiment Layout 

Experiment Module 

, 1 2 I ~ I 4015 L:6 ·17 J .. I!'., L9~ J; 1U .. 1' n ." 

·' A 

,:. . .,< 

.' 

1 I 2 1;3' I ~ 1' 5' k 6 -i ." ~, Il ., 'O,~rn, Legend . " 

: ., : .. B . 

.' 

1 I l I a I 4 I ~ 1"6 ' IJ7 ';j r8 j 1;;-9 <-:1 '10 I ~ Triplet 

c 

.;. 

" - 1 I 2 1·3 I 4 I 5 1 6 1 "7 1 8 r 9 · UOJJ1': 

0 

:, .,' .. ' ,- . - •. '1' ': :~:-'. 1,; .... .: . . --

D Concentrator 

., 

I 1 I ~ I I '_ 1 ~ . I" ',- I ' :t' l §;.!.] 

E 

.......... -----10"-----_ .. ~ 

19 
Molniya Concentrator 

Flight Experiment 

Flight test of four loffee concentrator modules 

• Flat fresnel concentrators 

• 16 element modules 

• Tandem solar cells 

Mounted on outer wall of a Molniya communications satellite 

Launch planned in summer 1995 on a Molniya launch vehicle 

from Baikanour 

Molniya orbit provides high radiation environment 

Photo of Fresnel lens concentrator panel on the basis 

2 of 16 modules (i nput photoactive square 100 em ). 

20 

~I 
- _._------

Objective 

C-SAVE 

Concept 

• Demonstrate and quantitatively measure the 

performance of a linear photovoltaic concentrator 

array 

v' Array pointing 

v' Oft-axis tracking 

v' Thermal distortion effects 

• Space qualify concentrator solar arrays 

Description 

• Two solar arrays sized to approximately 500 w each 

(1 kw EOl total) - each array will be 1.5 m2 

v' Either reflective or refractive optics 

v' Sunlight focused onto a strip of high efficiency, 

multi-band gap photovoltalc cells 

• Arrays will be deployed in space and will track the 

sun in one axis (when the experiment Is operating) 

Conclusions 

~ .-

Concentrator solar arrays provide significant benefits to future 

SMOO missions 

• Low cost 

• Light weight 

• Radiation resistant 

SMOO approach is to develop flight test modules and arrays to 

demonstrate technology maturity 

Space concentrator arrays will provide significant cost and mass 

advantages to a range of commercial satellite programs 

21 

------- -- - ----- --
AIR FORCE ACTIVITIES IN SPACE PHOTOVOLTAIC POWER SYSTEM TECHNOLOGY 

Kelly Gaffney 

U.S. Air Force 

Kirtland AFB, New Mexico 

TECHNOLOGY FOCUS 

PHOTOVOL TAles 

COVERGLASS 

'x ELLS 

POWER MANAGEMENT AND 

DISTRIBUTION 

IPP BACKSIDE 

ENERGY STORAGE 

FULL TECHNOLOGY LIFE CYCLE DEVELOPMENT 

HIGH RISK/HIGH PAYOFF 

NEAR TERM SOLUTIONS 

MARKET AND USER DRIVEN 

'T H MED 

23 

--------- -
NEW AIR FORCE SATELLITES 

FOR THE YEAR 2000 

MILSATCOM: TECH DATE 1998 

• 3-4 kW 

• GEO 

• ADVANCED ARRAYS 

(CONNETRATORS/FLEXIBLE) 

• MULTI-JUNCTION CELLS 

ALARM: TECH DATE 1998 

• 2-4kW 

• GEO 

• ADVANCED GaAs AND 

MULTI-JUNCTION CELLS 

• ADVANCED ARRAYS 

GPS IIF: TECH DATE 1997 

• SAME MISSION NEED 

• DESIGN FOR AFFORDABILITY 

• ADVANCED ARRAYS 

• MULTI-JUNCTION AND THIN"FILM 

CELLS 

MISSION UPGRADES 

DSP-23 

DMSP BLOCK 6 

BE BLOCK CHANGES 

POTENTIAL FOR OTV 

TECHNOLOGY CHALLENGES 

SATELLITE/LAUNCH POWER TECHNOLOGY DRIVERS 

LAUNCH VEHICLE 

ARRAY VOLUME (150W/M2) 

BATTERY/ARRA Y MASS 

(15-25% OF SATELLITE MASS) 

SATELLITE LIFETIMES 

(NiCd BATTERIES--3 YEARS IN LEO) 

(SOLAR CELL DEGRADATION) 

PRIMARY BATTERY PERFORMANCE 

(10% OF LAUNCH PROBLEMS) 

24 

SATELLITE 

NEED FOR MORE POWER 

SOLAR CELL COST ($1000/w) 

POWER SYSTEM MODULARITY & 

STANDARDIZATION. 

COMPONENT RELIABILITY (LATCHING 

RELAYS, BYPASS SWITCHES, ETC) 

-l 

[ 
SYSTEM BENEFITS FROM TECHNOLOGY 

,-----, EPS AS % OF SPACECRAFT MASS 

NOW 

NiCd 

RIGIDSi 

28V PMAD 

2.0 Wllb 

$10,0001W 

EXPERIMEN 

STRV 

PASP+ 

NaSTEC 

CPV NiH2 

RIGID GaAs 

2.9 W/lb 

$10,0001W NaS 

FLEXIBLE MBG ARRAY 

HIGH VOLTAGE PMAD 

4.0 Wllb 

$S,OOO/W 

TECHNOLOGY OPTIONS/BENEFITS 

PHOTOVOLTAICS TECHNOLOGY IMPROVEMENTS 

STATE OF THE ART: GaAs RIGID ARRAYS 

18% EFFICIENT CELLS 

SOW/KG BOL, 2S0W/M2 BOL 

STOWAGE PROBLEM 

ADVANCED POWER OPTIONS: 

LOW RISK: GalnP/GaAs/Ge RIGID ARRAY 

27% EFFICIENT CELLS 

8SW/KG BOL, 400 W/M2 BOL 

LESS OF A STOWAGE PROBLEM 

MEDIUM RISK: CIS AND CdTe THIN FILM CELLS 

12% EFFICIENCY, RAD-HARD 

40% LIGHTER, LESS SUBSTRATE 

SIGNIFICANTLY LOWER COST 

MEDIUM RISK: ADVANCED FLEXIBLE BLANKET 

OR CONCENTRATOR ARRAY 

100-1S0W/KG BOL 

SOME HAVE STOWAGE VOLUMES 

OF 0.1SM3 

25 
PHOTOVOL TAles 

TECHNOLOGY OPTIONS/BENEFITS 

GaAs CELL IMPROVEMENTS: 

PREMISE: ULTRA LIGHTWEIGHT GaAs SOLAR CELL BY 

INCORPORATING LIGHT-TRAPPING FEATURES ON THE 

BACK OF THE CELL. HIGH CURRENTS CAN BE OBTAINED 

WITH A VERY THIN LAYER OF MATERIAL. GOAL IS A 24.5% 

CELL AT 1 SUN AMO 

SUCCESS: 

• IMPROVED LPE GROWTH LAYERS--2CM X 4CM CELL 

• OPTIMIZED ANTI-REFLECTIVE COATING 

• INCORPORATED LIGHT TRAPPING STRUCTURE TO THICK CELLS 

FUTURE PLANS: 

• CHOOSE METAL FOR USE IN BACK CONTACT 

• IMPROVE FABRICATION PROCESS 

BENEFITS: 

• VERY LIGHT, HIGH EFFICIENCY GaAs 

• GOOD RADIATION TOLERANCE BECAUSE OF THIN DEVICE 

STRUCTURE 

PHOTOVOL TAICS 

TECHNOLOGY OPTIONS/BENEFITS 

MULTI-JUNCTION SOLAR CELL: AR C 

PREMISE: DEVELOP 25% EFFICIENT, MONOLITHIC, MULTI JUNCTION 

SOLAR CELLS. CELLS SHOULD BE TWO-TERMINAL, STANDARD 

SIZE AND COST NO MORE PER WATT THAN GaAs 

SUCCESS: 

• 24.6% EFFICIENT, TWO JUNCTION GalnP/GaAs CELL DEVELOPED 

• READY FOR FLIGHT TEST ON MIGHTYSAT 

• 26.7% EFFICIENT GainP/GaAs/Ge , 2 BY 2cm LAB CELL READY BY 

CHRISTMAS 

FUTURE PLANS: 

• THREE YEAR PLAN TO IMPROVE RELIABILlTYIYIELD OF TANDEM CE 

COLLABORATE WITH MANTECH AND NASA TO DEVELOP 

6 BY 6, HIGH BATCH PRODUCTION 

• GROUND AND RADIATION TESTING IN LATE 1996 

• FLIGHT QUALIFICATION IN EARLY 1997 

BENEFITS: 

• 35-40% REDUCTION IN WEIGHT OVER GaA.s 

• 40% REDUCTION IN ARRAY CROSS-SECTIONAL AREA FOR LAUNCH 

• COSTS SIMILAR TO GaAs 

26 

ating 

LL 

Frt Contact 

n+ t;aAs 

n- AlinP 

n- GainP 

p- GainP 

n- AllnP 

n+ (.;::1A!'; 

n+ (.;::1A!'; 

n- AI(.;::1A!'; 

n-G::1As 

p-GaAs 

p-GaAs 
PHOTOVOL TAICS 

TECHNOLOGY OPTIONS/BENEFITS 

• IIII .. -FILM SOLAR CELLS 

PREMISE: DEVELOP HIGH EFFICIENCY (12-1 5%), LARGE 

AREA DEPOSITION THIN-FILM PHOTOVOL TAlC 

DEVICES. FOCUS IS ON CIS AND CdTe 

SUCCESS: 

• DEPOSITION OF CIS ON 4 X 41N SUBSTRATE 

• 12% CIS CELLS IN SMALL SIZES 

• DEVELOPMENT OF 10% CIS ON 24 CM X 24 CM 

FLEXIBLE METAL FOIL BY DEC 

FUTURE PLANS: 

NEW PROGRAM TO DEVELOP REEL-TO-REEL 

DEPOSITION ON FLEXIBLE SUBSTRATES IN THE 

12-15% RANGE 

BENEFITS : 

• SIGNIFICANT COST REDUCTIONS OVER CRYSTALLINE 

CELLS 

• 30-40% ARRAY WEIGHT REDUCTION 

• REDUCED STOWAGE VOLUME FOR ARRAY 

PHOTOVOL T AICS 

TECHNOLOGY OPTIONS/BENEFITS 

ADVANCED FLEXIBLE BLANKET ARRAY: 

PREMISE: TO DEVELOP AND FLIGHT QUALIFY TWO 

FOLD-OUT ARRAYS WITH AT LEAST 150W/Kg EFFICIENCY, 

0.15m3 STOWAGE VOLUME AND SYSTEM LEVEL COSTS 

OF LESS THAN $5001W IN 1-3 KW POWER RANGE 

SUCCESS: AWARDING OF TWO SEPARATE CONTRACTS 

• ROLL-OUT FLEXIBLE BALNKET USING CIS THIN-FILM CELLS 

AND SHAPE MEMORY CONTROL 

• RIGID, COMPOSITE, FOLD-OUT ARRAY USING ANGLED 

MIRRORS FOR SUN CONCENTRATION 

FUTURE PLANS: 

• DEVELOP AND FABRICATE ARRAYS 

• FLIGHT TEST IN 1998 

• POTENTIAL INTEREST BY POST-1999 SATELLITES 

BENEFITS: 

• 3-FOLD INCREASE IN ARRAY EFFICIENCY 

• SIGNIFICANT DECREASE IN ARRAY STOWAGE VOLUME 

• 50% COST SAVINGS ON ARRAY PROCUREMENTS 

27 
PHOTOVOL TAICS 

TECHNOLOGY OPTIONS/BENEFITS 

INTEGRATED POWER SYSTEM: LOW RISK 

PREMISE: TO DEVELOP AND DEMONSTRATE A MODULAR, 

STANDARDIZED POWER SYSTEM ARCHITECTURE USING 

SHUNT REGULATION MOUNTED TO THE GaAs SOLAR ARRAY 

USING A HYBRID PATCH AND RESISTOR STRIP. SYSTEM 

IS ABLE TO ACCOMADATE GROWTH BETWEEN 100W--

5KW WITH MINIMAL REDESIGN. 

SUCCESSES: 

PANEL DEVELOPMENT COMPLETE 

PRELIMINARY GROUND TESTING OF COUPONS COM 

FLIGHT TEST PROGRAM IN PLACE 

FUTURE PLANS: 

FLIGHT TEST ON STP MISSION IN LATE 1995 

BENEFITS 

MODULAR, STANDARD ARCHITECTURE MEANS LESS 

REDESIGN AND MINIMUM REQUALIFICATION: LOWER COST 

LESS SYSTEM WEIGHT: 12 W/KG 

DIELECTRIC 

IPP BACKSIDE 

FACESHEET 

BETTER THERMAL CONTROL 

PHOTOVOL TAICS 

FLIGHT EXPERIMENTS: NEW STARTS 

LABSAT 

FLIGHT TEST HIGH EFFICIENCY DUAL JUNCTION 

CELLS AND TRIPLE JUNCTION CELLS ON COMPOSITE SUBSTRATES 

HARDWARE DELIVERY DATE: 30 SEP 1995 

CONDUCT DESIGN ANALYSIS FOR FUTURE LABSAT 

POWER SYSTEM UPGRADES 

UoSAT 

COOPERATIVE EFFORT WITH UK, NASA, NRL, BMDO 

TO FLIGHT TEST POWER SYSTEM COMPONENTS 

PROBABLE AF EXPERIMENTS INCLUDE NaS LONG-TERM 

FLIGHT TEST, ADVANCED ARRAY DEMO, TRIPLE JUNCTION 

SOLAR CELL FLIGHT 

LAUNCH DATE IS JUN 1996 

28 
STRV 1B 

- - ----

PHOTOVOL TAles 

FLIGHT EXPERIMENTS 

FLIGHT TEST OF ADVANCED CELL TECHNOLOGIES. 

PRIMARY POWER PANELS ARE GaAs, 5.5MIL MANTECH 

CELLS. EXPERIMENTAL PANELS HAVE 20 DIFFERENT 

EMERGING CELL TYPES 

PANELS WILL BE TESTED OVER THREE MISSION FOR 

ELECTRICAL PERFORMANCE ANND RADIATION 

RESISTANCE. 

PASP+ 

FLIGHT TEST OF 12 ADVANCED ARRAY DESIGNS: 

FLEXIBLE BLANKETS, CONCENTRATORS, THIN-FILMS 

ARRAYS WILL BE TESTED FOR ENVIRONMENTAL AND 

PLASMA INTERACTIONS, HIGH VOLTAGE PERFORMANCE 

AND ELECTRICAL PERFORMANCE 

LABORATORY PROGRAM 

SOLAR CELL ASSESSMENT 

• FULL LABORATORY CAPABILITY TO 

CONDUCT MECHANICAL AND ELECTRICAL 

TESTING OF SOLAR CELLS 

• RADIATION DEGRADATION TESTING AT 

UK ORA HARWELL FACILITY 

• SOLAR CELL DATABASE TESTING WILL 

START THIS YEAR (12 CELLS ALREADY 

IDENTIFIED--TANDEM, THIN-FILMS, GaAs, 

ADVANCED Si) 

• DATABASE WILL BE OPEN TO ALL 

• WILL TEST ANY CELLS UPON SPO REQUEST 

AT NO COST 

29 

-- ----' 
CONCLUSIONS 

NEW AIR FORCE SYSTEMS ARE COMING ON LINE BY THE TURN 

OF THE CENTURY: LIFE CYCLE COST WILL BE THE BIG DRIVER 

POWER TECHNOLOGIES AS BOTH COST AND WEIGHT REDUCTION 

TOOLS ARE OF INTENSE INTEREST TO OUR CUSTOMERS 

TECHNOLOGY TRANSITION PATH FOR TANDEM CELLS AND ADVANCED 

ARRAYS IS IN PLACE 

LABORATORY WILL CONTINUE ON THE PATH TO LIGHTER, CHEAPER 

CELL AND ARRAY TECHNOLOGIES 

NEW PROGRAM INITIATIVES TO WATCH FOR: TANDEM CELL MANTECH, 

ADVANCED THIN-FILM CELL DEVELOPMENT, TANDEM CELL FLIGHT TEST 

SOLAR CELL ASSESSMENT, RENEWED INTEREST IN ARRAYS FOR 

OTV 

30 

. '" ~ 
.1 

I 

I 

l 

39?73CJ 

N95-20505 

AN UPDATE ON THE DEPARTMENT OF ENERGY'S PHOTOVOLTAIC PROGRAM 

John P. Benner and Mark Fitzgerald 

National Renewable Energy Laboratory 

Golden, Colorado 

Summary 

Funding for the terrestrial photovoltaics program is $78 million in 1994. This is more than 

double the minimum level reached in 1989 and runs counter to the general trend of decreasing 

budgets for Department of Energy (DOE) programs. During the past five years, the program 

has expanded its mission from research and development to also address manufacturing 

technology and commercialization assistance. These new activities are directed toward 

revital izing the market to reinstate the rapid rate of sales growth needed to attract investment. 

The program is approaching balance among efforts in each of the three areas. This translates to 

a reduction in some of the R&D activities of most relevance to the space power community. 

On the other hand, some of the advancements in manufacturing may finally bring thin-film 

technologies to reality for space arrays. This talk will describe the status and direction of DOE 

program with an eye toward highlighting its impact on technology of interest for space. 

Introduction 

During the latter part of the 1980's, the growth rate in photovoltaic module shipments was more 

than 20% . Industry shifted from a condition of shipping from excess inventory to back-orders. 

During the last several years, world wide sales growth has slowed to less than 5%. This 

condition is attributed to the world wide recession . This condition does not help the industry 

attract capital needed for expansion of production or introduction of new products. The 

Department of Energy has addressed this problem by adding two programs focussed at helping 

industry to increase sales. First, the program directed support to manufacturing technology 

research to further improve competitiveness and open new applications by reducing product 

cost. The second initiative increased interactions with envisioned customers to enhance 

awareness of photovoltaic technology for applications where it is competitive today. The result 

on the total program budget can be seen in Figure 1. 

00~---------------------------' 

80 . 

60 ...... . ...... .. .. ...... . . 

87 88 89 90 91 92 93 94* 95** 

Fiscal Year 

Figure 1. US DOE Photovoltaic Program Budgets 

31 

G Capit al Equip m en t 

tH Systems & Market Dev 

[IlJ Mfg Technology Dev 

o Gov ' t LabS.) S tra tegic 

[ffiJ Universities R&D 

• Industry 

• FY 94 Es t. p rior to resciss io llS 

•• FY9 5 Es t. based on DOE bud get reques t 

---~ 
Photovoltaic Manufacturing Technology 

The Photovoltaics Manufacturing Technologies (PVMaT) project is a partnership between the 

U.S. Department of Energy and the U.S. PV industry. By working together to reduce costs, the 

partners ultimately hope to extend the U.S. leadership in manufacturing and developing 

commercial PV systems. 

The PVMaT project is being conducted in three phases. Phase 1 identified and prioritized areas 

in the manufacturing processes where research and development were needed for major 

production cost reductions. Problem solutions began in 1992, under Phase 2A, while Phase 3A 

addresses generic module manufacturing needs common to the PV industry. Current industry 

partners, topic and award amounts are shown in Figure 2. 

Company Topic Principal Amount 

Investigator ($M) 

Phase 2A 

Siemens Soar Industries Cz Silicon Terry Jester 10.5 

Solarex Corporation Triple Junction a-Si Alloys Robert Oswald 10.0 

ENTECH , Inc. Linear Focus Conc. Modules Mark O'Neill 3.1 

Astropower , Inc. Silicon Film Technology Sandi Collins 7.1 

Utility Power Group a-Si Modules Michael Stern 4.7 

Energy Conversion Devices Roil-to-Roll a-Si Masatsuga Izu 10.8 

Phase 28 

Golden Photon, Inc. CdTe Steve Johnson 9.8 

Solarex Corporation Cast Silicon John Wohlgemuth 6.3 

Solar Cells, Inc CdTe Dan Sandwisch 7.4 

Texas Instruments Spheral Silicon Jim Skelly 16.6 

Phase 3A 

Springborn Laboratories Improved Encapsulants Bill Holley 1.5 

Spire Corporation Automated Cell Assembly Mike Nowlan 1.4 

Figure 2 PVMaT Industry Partners 

Systems and Market Development 

The DOE has established the PV Compact Coordinating Council as the focus of market 

development activities in the United States. PV Compact consists of utility representatives from 

the Utility Photovoltaic Groups (UPVG) and representatives from state energy offices and state 

Public Utility Commissions (PV4U). The DOE/NREL PV Program is supporting the 

development of a strategy to commercialize PV through UPVG. This is an organization of more 

than 70 utilities that work together to advance cost-effective and high-value uses of PV for 

utilities. UPVG activities include publishing a newsletter, sponsoring workshops, and supporting 

five active working groups working towards a goal to lead utilities into photovoltaic systems 

acquisitions of about 50 megawatts over five years with less than 30% government cost sharing. 

PV4U is a project to work with and assist state working groups involved with the development of 

PV systems by utilities. 

Building Opportunities in the United States for Photovoltaics (PV:BONUS) is a DOE/NREL 

project to evaluate and implement innovative ideas for integrating PV into building systems. 

Twenty-nine teams proposed concepts, and, in FY 1993, five teams began work on the planned 

5-year, $25-million (from DOE) cost-shared project. Each of the five teams includes a lead 

contractor and up to 10 other organizations representing building-materials manufacturers, 

building contractors, PV suppliers, utilities, colleges and universities, systems designers, 

architectural and engineering firms, and building owners. 

32 
... 

The DOE/NREL PV Program helps U.S. PV companies reach the growing international markets 

by leveraging funds and lowering institutional barriers to international trade in PV technology. 

DOE shares the cost of pilot installations, and PV Program personnel help perform site 

assessments, develop project ev~luation criteria, write technical specifications for projects, and 

conduct workshops and training seminars. 

In FY 1993, DOE and the Federal Republic of Brazil began working together to bring electricity 

to rural communities in Brazil using PV. The first phase began in December 1992 with the 

installation of lighting systems on 600 homes, schools, and public spaces in rural areas. Phase 2 

of this project is now underway, with DOE contributing approximately $1 .3 million toward the 

estimated $3 million cost of the project's second phase in FY 1994. Other project development 

efforts are underway in China, India, and Indonesia. 

The guiding principal behind all of the market development activity is the establ ishment of 

expanded markets that are sustained by the competitiveness of photovoltaic products such that 

ongoing Federal support, in tax policy or direct funding , is not required. 

Strategic Research and Development 

The budget for photovoltaic research has been held flat during the period in which the new 

initiatives in manufacturing and market development were added. However, within the research 

program, priorities have changed substantially. The support for thin-film technologies has grown 

to keep pace with inflation. Other projects needed to be reduced to afford to maintain thin-films. 

Much of this reduction came from support of concentrator and high efficiency options. Currently, 

the newest element of the research program is the active request for proposals to form a ThinFilm Partnership. This single solicitation will draw in proposals from both the major industrial 

interests and well as the research teams from industry, universities and not-for-profit 

laboratories. The goals of the procurement are to: i) support the successful introduction of U.S. 

thin film products; and, ii) support advanced thin-film R&D needed for future product 

competitiveness. Since this is an active solicitation, no more can be described in this paper. 

33 
3q??J~ 

N95-20506 

EUROPE'S SPACE PHOTOVOL TAICS PROGRAMME 

Klaus P. Bogus 

European Space Agency 

Noordwijk, The Netherlands 

SUMMARY 

The current space PV technology development programme of ESA is described. The programme is 

closely coupled to the European space mission scenario for the next 10 year period and has as main 

objective to make the most effective use of the limited resources available for technology in the 

present economical climate. This requires a well-balanced approach between concentration on very 

few options and keeping the competition alive if more than one promising technology exists. 

The paper describes ESA's main activities in the areas of solar array technology, solar cell 

technology, solar cell assembly technology, and special test and verification activities including the inorbit demonstration of new technologies. 

INTRODUCTION 

Europe has had a very successful record in the field of space photovoltaics and has achieved a 

competitive position in the world market. For a large variety of European spacecraft systems 

photovoltaic solar generators are the exclusive source of electrical power and additionally, Europe 

has succeeded in exporting solar arrays across the Atlantic (e.g. IUE, Hubble Space Telescope, CTSHermes and INTELSAT). 

The main purpose of the space photovoltaic technology development programme is to maintain the 

high standard of European solar array technology by carefully modernising the existing concepts in a 

step-wise approach. The smooth introduction of new technologies into flight programmes is achieved 

by close cooperation between technologists and project engineers. 

The close coupling of the technology programme to the European mission scenario for the next 10 

years is perfectly in line with the approach described above and leads to a fast application cycle of 

new technologies. A disadvantage of this approach is the resulting low priority for globally attractive 

new technologies which are not directly required in the ESA mission scenario with it's limited scope. 

DETAILS OF ESA'S SPACE PV TECHNOLOGY PROGRAMME 

The programme is described in the 7 tables below according to a systematic division into solar array 

technology , solar celt technology , assembly technology, technology verification activities and 

supplementary activities. 

The first column of each table contains the technology activity title, the second column indicates the 

present status of the activity. "Proposed" stands for a new activity which is not yet accepted in the 

budget planning, after acceptance it turns into "planned" , and after initiation into "running". The 3rd 

column identifies the frame-programme unter which the activity is funded : "Basic" is the general 

basic TRP(Technological Research Programme) of ESA, "ASTP" is the acronym for applied 

supporting technology programmes of the Telecommunications satellite area and "GSTP" is the newly 

introduced general supporting technology programme of ESTEC for all ESA missions. Additionally, 

several of the national activities in the ESA-member states are harmonised with ESA and run under 

common management arrrangements. Schedule indications can be found in the 4th column. The last 

column contains summary descriptions. 

SOLAR ARRAY TECHNOLOGY. 

ESA's Solar Array Technology is based on a two main design concepts: Advanced lightweight rigid 

panel arrays and advanced flexible blanket arrays. With both concepts a high degree of maturity and 

flight experience has been accumulated in recent flight programmes: OL YMPUS-1, HST and ERS-1 

35 
are operating very successfully with flexible blanket arrays, whereas EURECA, HIPPARCOS and the 

ECS/MARECS satellites are supplied with rigid panel arrays. 

The solar array technology development aims at the improvement of the present lightweigth carbonfibre face-sheet panels for specific future telecom-missions. Since the technology requirements are 

generally very mission specific it has become general practice do do these developments within the 

corresponding flight programmes and not in generic technology programmes. 

Unfortunately EUROPE has presently no new programme requiring flexible blanket solar arrays so 

that this technology which has already reached a high degree of maturity, can not be further 

improved. 

The development of alternative concepts was limited to studies on concentrator arrays (SARALouvre, Holographic Dispersive). 

SOLAR CELL TECHNOLOGY. 

This part of the programme comprises two main elements: Improvement of silicon solar cells and 

development of GaAs solar cells: 

- Hi-ETA silicon cells with 16-17% efficiency have been pre-developed and are now approaching the 

pilotline production stage. 

- Further improvement towards 18% efficiency are under pre-development using advanced 

passivation and optical confinement 

-a demonstration of a 20% efficiency silicon cell is foreseen in a basic R&D study on small area, low 

quantity samples 

-GaAs and GaAs-on-Ge cells have been pre-developed in Italy, Great-Britain and Germany. 

-The next step will be a pilotline production of GaAs-on-Ge cells 

- Further improvements of III-V-compound cells are expected in the area of ultra-thin (superstrate 

supported) GaAs cells and multi-junction/tandem cells. 

SOLAR CELL ASSEMBLY TECHNOLOGY. 

This part of the Programme covers : 

- the development of improved Aluminium solar cell interconnectors (ATOX-resistant, low-cost) 

- the development of ultrasonic welding for Si- and GaAs cells ( long cycling life, low-cost ) 

- advanced cover-glass bonding (Direct Electrostatic bonding, Teflon Pre-preg bonding) 

- Infra-red reflective coatings on cover-glasses (improved efficiency through lower Temp.) 

- Development of integrated solar cell modules (GaAs-thin film and Silicon superstrate concepts). 

TECHNOLOGY VERIFICATION TESTS. 

Apart from the activities performed at SPASOLAB (ESA's solar cell laboratory in Madrid), the main 

activities are related to the investigation of space environmental effects on solar arrays. This includes 

investigation of synergistic effects, thermal cycle induced fatigue, plasma and atomic oxygen effects, 

micrometeorite impact effects and particle radiation damage in solar cells. Radiation damage 

investigations are required for two different reasons: One is the evaluation of new solar cells (e.g. 

advanced GaAs and Hi-ETA silicon cells), the other the planning of missions in different orbits (e.g. 

the new Telecom-missions in high-inclination, medium altitude orbots with equivalent fluences of 

more than10E16 One-Mev-electrons Isqcm). 

SUPPLEMENTARY ACTIVITIES. 

This segment contains mainly the preparation and evaluation of flight experiments and the Post-flight 

investigation programmes on the HST and EURECA solar arrays reported in another part of this 

conference. 

36 

-. --------------------
SOLAR GENERATOR TECHNOLOGY PROGRAMME 

1994 STATUS 

1. ARRA Y SYSTEM DEVELOPMENT 

2. ASSEMBLY TECHNOLOGY 

3. COMPONENT TEC HNOLOGY 

4. TECHNOLOGY VERIFIC ATION TESTS 

S. SUPPLEMENT ARY ACTIVITIES 

SOLAR GENERATOR TECHNOLOGY PROGRAMME 

1. ARRAY SYSTEM DEVELOPMENT 

SOLAR ARR AY FOR RUNNING ASTP-4 94-95 ADAPTATIO OF RIGID ARRAYS 

HIGH POWER APPL. 9 TO 5-7kWTELECOM APPUCATIO S 

ADV ANCED RI GID RUNNING TELECOM 93-95 ARAFOM (FSS), HOTBIRD+ 

PANEL ARRAYS HARMO GSR(AS),MARK-3 (DASA) 

L1LPARRAYS RUNNING BA IC 91-93 LOW POWERILOW INTEN ITY 

4 ARRAYS FOR MARSNET 

SYSTEM A ALYSIS & PLANNED BASIC 95-96 EV ALUA TION OF NEW CELL TYPES 

A SESSM. NOVEL CELLS 9 ON ARRAY LEVEL 

SOLAR GENERATOR TECHNOLOGY PROGRAMME 

2. ASSEMBLY TECHNOLOGY 

ALUMI I MINTER ON. RUNNING BASIC 87-94 DEVELOPM. OF tILTASONICALLY 

TECHNOLOGY 6 WELDED ALU-INTERCONNEC,ORS 

SOLAR CELL ASSEMBL. PROPO ED ASTP-4 CONTINUE ASTP-3 DEVELOPM. 

TECIIN. (SI & GaAs) 9 FOR NEW CELL TYPES 

I T. ELEV. VOLTAGE RUNNING BA IC 93-95 DEV. OF INTEGRATED GAAS 

MODULE 6 lILTRATIllN MODlILES WITH D10DFS 

I-R-REFLECTING RUNNING ASTP-4 92-94 REDUCTION OF OPER. TEMPERATURE 

COATI G 9 11) REFLECTOR 0 COVER-GLAS 

TEFLON-BONDI G Rt !lN(, \' 11>·4 n -94 TU ·LO BONDED CELL-ASSEMBLIES 

OF GLASS ON CELL 9 wml ESD-PROTR.lION 

37 
SOLAR GENERATOR TECHNOLOGY PROGRAMME 

3. COMPONENT TECHNOLOGY-A 

ALTER AT. SOLAR RUNNING BASIC 89-94 EVALLJATlO OF "3RD GHffiRA'nO 

CELL ASSESSME T 9 CELLS r-OR SPACE 

GaAs·SOLAR·CELL RUNNING ASTP·3 87·94 DEY. 01- GaAS CEU.S BY MOCVD 

DEVELOPMENT &llARMO ON GaAs AND GI:-RMANTUM 

LILT SOLAR CELLS RUNNI G BA IC 91·95 DbY 01- CELLS WITH 25%-El-F. 

4 IN DEEP SPACE (ROSEITA) 

THIN·FILM CELLS RUNNING ASTP·4 92·94 PRE·OE'VELOPMENT 01- CIS·CFJ.LS: 

SYSIT,/\I·ASSESSMI· T 

SOLAR GENERATOR TECHNOLOGY PROGRAMME 

3. COMPONENT TECHNOLOGY-B 

PILOTLI NE HI-ETA RlfNNIN(; G~TP·I 94-96 END· DEVeLOPMENT Of- 16%-ER-. 

SI LlCO ' CELLS 9 CELI.S I CL. PILOT PROD. 

ADVA 'CEO LlGTHW. PLA ED GS·£1>· 1 94-% CASCADE'Il'ANDh I CI-1J.S:>22%. 

GAASCELLS 9 UI .TRA ·THlN 

GE-SLJBSTRATES FOR RUNNING G~TP 1 93-95 DEVELOP lMPROVID SUBSTRATES 

GAASCELLS 9 IN PILOTI.INE 

SOLAR GENERATOR TECHNOLOGY PROGRAMME 

4. TECHNOLOGY VERIFICATION TESTS 

S·!\. ENVI RONMENT AL Rl NG BASIC 93·94 lDENllF. AND TEST OF SYNERGISTIC 

INTERACTIONS 6 EFFECTS: MICROM. PROTONSATOX 

ESO·TEST Rl !Nn. BA;,IC 93 MODEl . OF E'>D INTERAc no S 

SIMULAT ION OF S·A', : RhPR. TEST 

SPASOI.AB RUNNING BASIC 93·95 TYPE APPROVAl . TESTS (l'S5·01·6O-I) 

ARCH IMEDES RUNNING TlASIC 94 p+ & c· TESTS 01- OVf'] . CEU .S 

I-OR M1SSIO S WlTlI ID(;H Fl.l IX 

38 I 

~ 
SOLAR GENERATOR TECHNOLOGY PROGRAMME 

5. SUPPLEMENTARY ACTIVITIES 

FLIGI-IT EXPERIMENTS RUNNING TDP 94 IN-ORBT VERlFlCA TION OF NEW 

CELL (STRV. HEALTH AT.ETC.) 

PORTABLE RUNNING GSTP-J 94-95 MINI-FLASHER FOR HELD TEST 

UN IMULATOR 

EURECASA RUNNING BASIC 93·94 POST·FLIGHT INVESTIGATIONS 

I-ISTSA-I RUNNING BASIC 94-95 POST·FLIGHT rNVESTIGATIONS 

39 
_._

..... . . 

SESSION I 

InP CELL DEVELOPMENT 

II /0 ~ PAQi tlLANK NOT fll En 
N95- 20507 

HIGH-EFFICIENCY, DEEP-JUNCTION, EPITAXIAL InP SOLAR CELLS ON 3 ?3¥ 

(100) AND (111)B InP SUBSTRATES /0 f' 

R. Venkatasubramanian, M.L. Timmons, and J.A. Hutchby 

Research Triangle Institute 

Research Triangle Park, North Carolina 

and 

R. Walters and G. Summers 

U.S. Naval Research Laboratory 

Washington, DC 

ABSTRACT 

We report on the development and performance of deep-junction (-0.25 11m), gradedemitter-doped, n+-p InP solar cells grown by metallorganic chemical vapor deposition (MOCVD). 

A novel, diffusion-transport process for obtaining lightly-doped p-type base regions of the solar 

cell is described. The I-V data and external quantum-efficiency response of these cells are 

presented. The best active-area AMO efficiency for these deep-junction cells on (100)-oriented 

InP substrates is 16.8%, with a J sc of 31.8 mAlcm2, a Voc of 0.843 V, and a fill-factor of 0.85. 

8y comparison, the best cell efficiency on the (111 )8-oriented InP substrates was 15.0%. These 

efficiency values for deep-junction cells are encouraging and compare favourably with 

performance of thin-emitter (0.03 11m) epitaxial cells as well as that of deep-emitter diffused cells. 

The cell performance and breakdown voltage characteristics of a batch of 20 cells on each of the 

orientation are presented, indicating the superior breakdown voltage properties and other 

characteristics of InP cells on the (111)8 orientation. Spectral response, dark I-V data, and 

photoluminescence (PL) measurements on the InP cells are presented with an analysis on the 

variation in Jsc and Voc of the cells. It is observed, under open-circuit conditions, that lower-Voc 

cells exhibit higher band-edge PL intensity for both the (100) and (111)8 orientations. This 

anamolous behaviour suggests that radiative recombination in the heavily-doped n+-lnP emitter 

may be detrimental to achieving higher Voc in n+-p InP solar cells. 

INTRODUCTION 

Epitaxial n+-p InP solar cells remain attractive for space photovoltaic applications as they 

have demonstrated a high AMO conversion efficiency (-18 to 19%) [1]. These cell efficiencies 

have been obtained through material improvements by epitaxy as well as the use of a thin emitter 

(-0.03 11m). However, the absence of photon-assisted annealing of radiation-induced defects and 

the consequent lack of superior radiation resistance of these epitaxial thin-emitter cells stand in 

sharp contrast to that observed in diffused, deep-junction (-0.3 11m) cells [2]. 

This work aims to develop high-efficiency InP solar cells by MOCVD that replicate the 

diffused-junction structure, especially the deeper junction and the emitter surface doping gradient. 

The radiation resistance data from these cells are reported separately in a companion paper [3]. 

The goal is to improve the understanding of the radiation resistance of InP solar cells. In this 

work, we also present the first reported performance of InP solar cells on (111 )8-oriented 

substrates. 

The motivation for the study of cell perfOFmance on (111) orientation stems from the 

predicted advantages for hetero-epitaxial InP solar cells on Si substrates. There is about 8% 

lattice mismatch between InP and Si. This lattice mismatch can potentially introduce about 1.9 x 

~P

43 

AGE 
1014 dangling bonds per cm2 at the interface of (100) InP-Si. However, the calculated dangling 

bond density for the (111) orientation is about 1.1 x 1014 per cm2. Thus, for similar growth 

conditions and defect control mechanisms, the (111) orientation potentially offers a 40% 

reduction in defect-density. 

OMVPE GROWTH AND CELL STRUCTURE 

The n+ -p cell structures were grown on p-type InP substrates, doped to mid 1018 cm3. 

Fig. 1 shows the schematic of the cell structure. The best results of epitaxial growth on (100)-

oriented substrates were obtained with use of no additional surface preparation to the customerprovided epi-ready InP wafers. However, it was found necessary to give a brief surface clean 

(consisting of a 1 min. etch in 1:1 :5= H20:H202:H2S04 solution followed by a thorough rinse in 

deionized water) for the (111 )B-orientation substrates prior to epitaxy. This cleaning step is 

critical to obtaining a smooth epitaxy of InP layers on the (111) oriented substrates. The growth 

was carried out in an atmospheric-pressure MOCVD system using ethyldimethylindium and 

phosphine as the Group III and Group V precursors, respectively. All the growths were carried 

out at 700°C, with typical growth rates of 0.05 }.lm/min. 

Diethylzinc (DEZn) bubbler source was used for obtaining Zn-doping. An important 

requirement for obtaining high-efficiency as well as radiation-resistant n+ -p InP cells is that the 

base region be lightly-doped to mid 1016 cm-3. This lightly-doped base also enables the 

evaluation of a larger portion of the quasi-neutral p-base by techniques such as deep level 

transient spectroscopy (by appliying a larger voltage-bias without breakdown of the junction) to 

study radiation-induced defects. 

Typically, a low concentration dimetylzinc-in-hydrogen gas source is used for obtaining 

lightly-doped p-base. At RTI and other laboratories [4], this gas source has been found to lead to 

very erratic doping levels. Hence, in this work, we have investigated a new approach to obtain 

lightly-doped p-lnP using a diffusion transport process. In this technique, the inlet of an 

organometallic-Zn bubbler is closed and the outlet is kept open. H2 gas flow is maintained 

through a bypass line, adjacent to the bubbler, to continously carry the organometallic-Zn that 

diffuses out from the bubbler to the growth zone. 

This approach avoids the possible residual moisture and oxygen contamination, 

frequently present in gas cylinders, that can scavenge the Zn , especially at low ppm levels. Also, 

this approach is likely to provide a constant molar ratio of organometallic-Zn to H2, as long as the 

organometallic Zn-source is kept at a constant temperature, unlike a high-pressure gas-mixture 

that can lead to inconsistent concentration levels over the lifetime of the source. 

We show in Fig.2(a) the polaron profile of a Zn-doped base region, obtained with a flow 

of 1 ccm through a diethylzinc (DEZn) bubbler causing a carrier concentration of -4x1 018 cm-3, 

suggesting Ultra-low flow rates to get a p-doping level in mid-1016 cm-3. In Fig. 2(b), we indicate 

that with a 10ccm H2 flow and using the diffusion-transport process, and using the vapor 

pressure of dimethylzinc (DMZn), we obtain a doping level of -6 x 1017 cm-3. In Fig. 2(c), we 

indicate that with the same 10 ccm H2 flow and using the diffusion-transport process with a DEZn 

bubbler, we obtain a doping level of -8 x 1016 cm-3. The lower doping level obtained with the 

DEZn source is consistent with the lower vapor pressure of DEZn, compared to DMZn, therefore 

leading to a lower concentration of Zn although the same 10ccm of H2 is used to transport the 

diffused species. 

44 

-l 
- - - - - - - - ----

H2Se gas source was used for obtaining the heavily doped n+ emitter regions. Se is the 

same Group-VI n-type dopant as in S-doped InP diffused junctions, as opposed to the use of Si 

in some of the work [1] of epitaxial InP cells. The emitter was nominally linearly graded with 

increasing H2Se flow rates, to obtain a surface doping of -3 x 1018 cm-3. 

The cell structure in Fig. 1 also indicates the use of Ti/Au (30nm/300nm) contact for the 

p+ InP substrate. This contact was sintered at 415°C for one minute. The front emitter contact 

was a non-alloyed AuGe/Ni/Au (50nm/10nm/300nm) metallization. The InP cells on (100)-

oriented substrates received a two-layer ZnS/MgF2 anti-reflection coating (ARC) as indicated in 

Fig. 1. The ZnS/MgF2 coatings were deposited bye-beam evaporation. However, the cells on 

(111 )B-orientation received a single-layer ARC of plasma-deposited silicon nitride because the ebeam evaporated ZnS/MgF2 coatings strongly deteriorated the performance of the cells on 

(111 )B-oriented substrates. The cause for this behaviour is not exactly clear at this point. 

CELL I-V AND SPECTRAL-RESPONSE CHARACTERIZATION 

The I-V and external quantum efficiency (spectral response) data of the cells were 

measured using an InP standard cell characterized at the National Renewable Energy Laboratory 

(NREL). Relative external quantum efficiency data for a deep-junction cell are shown in Fig. 3. 

The cell indicates a good red response and a small roll-off in the short-wavelength response. This 

roll-off is attributed to some recombination of photo-generated carriers either in the emitter region 

or at the surface of the emitter. The surface recombination velocity of n+lnP has been predicted 

to be low [5]. In addition, the electric field resulting from the graded doping in the emitter is 

expected to oppose the surface recombination of minority carriers (holes) and accelerate them 

towards the depletion layer. As discussed below, based on photoluminescence data, we believe 

that the roll-off is related to radiative recombination of minority carriers in the heavily-doped n+-

emitter regions. The radiative recombination is attributed to lower radiative-lifetimes in the quasineutral regions of n+-lnP. 

The best active-area AMO efficiency for these deep-junction cells on (100)-oriented InP 

substrates is 16.8%, with a Jsc of 31.8 mAlcm2, a Voc of 0.843 V, and a fill factor of 0.85. By 

comparison, the best cell efficiency on the (111 )B-oriented InP substrates was 15.0%, with a Jsc 

of 31 .3 mAlcm2, a Voc of 0.797 V and a fill factor of 0.815. These efficiency values for deepjunction cells are encouraging and compare favorably with performance of thin-emitter (0.03 ).lm) 

epitaxial cells as well as that of deep-emitter diffused cells. 

The cell (efficiency) performance of a batch of -20 cells, each on (100)- and (111 )Boriented InP substrates, are shown in Fig. 4. We can immediately observe the spread of cell 

efficiencies, for both the orientations, over a reasonably wide range. However, we note the tighter 

(and a more statistically expected Gaussian-like) spread of cell efficiencies on (111) substrates 

and a more random fluctuation on the (100) substrates. 

ORIENTATION-EFFECTS ON REVERSE-BREAKDOWN VOLTAGES 

Similar statistical behavior in the breakdown voltage of cells on (100) and (111)B 

substrates was obtained. Fig. 5 shows a histogram depicting the distribution of breakdown 

voltages of the cells on the two orientations. For a given base doping level, the breakdown 

voltages of nearly all the cells on (111) InP substrates were within 4.5-5.2 V , while the cells on 

(100) substrates had considerable variation. Two of the cells, from the same wafer of (100) 

45 
orientation, with very similar forward characteristics had completely different reverse-breakdown 

characteristics. In contrast, any two cells even from different wafers of (111) orientation had 

similar reverse characteristics. 

The cause for this variance in behavior of reverse-breakdown characteristics is thought 

to be related to the more electronically-active (111)B InP surface than the (100) surface [6]. 

leading to a more stablized surface from a more-accelerated native-oxidation process. Thus the 

junctions on (111 )-oriented substrates probably suffer from less surface-induced, soft-breakdown 

reverse-characteristics. In addition, it has been reported that under certain doping conditions, the 

(111) orientation can offer significantly higher breakdown voltages than the (100) orientation in 

GaAs [7]. 

In any case, the reproducibility of InP cell-efficiency and breakdown-voltage 

characteristics on (111 )B-oriented substrates, compared to (100)-orientation, is noteworthy. This 

portends well for the investigation of InP cells on (111 )-Si substrates. 

PL CHARACTERZATION OF InP CELLS 

Band-edge photoluminescence (PL) from the InP cells were also evaluated at 300K to 

understand the variation in Jsc and Voc of the cells. The PL excitation source was an Ar-ion laser 

operating at 515 nm. The PL measurements were conducted under open-circuit conditions and 

the laser radiation is expected to be fully absorbed within the 0.25-llm-deep emitter region. 

In Fig. 6, we indicate the PL data from two InP cells on (100) substrates, nominally grown 

with the same cell structure (shown in Fig. 1) and nearly identical Jsc values. The cell (1-1955-a4), with a significantly higher Voc (0.843 V) , shows a much smaller band edge PL signal than the 

cell (1-1961-a-4) with a lower Voc of 0.783 V. The same pattern is once again observed in cells 

grown on the (111)B orientation, and is shown in Fig. 7, i.e. , the lower Voc cells indicate higher 

band edge PL intensity. This leads us to suggest that the lower Voc of n+ -p InP cells is perhaps 

related to lower radiative lifetime in the n+-emitter regions. This argument would be consistent 

with the higher PL intensity and the well-documented low surface-recombination velocity of n-lnP 

and n+-lnP surfaces. 

It has been known for some time that the open-circuit voltage of state-of-the-art n+ -p InP 

solar cells are not as high as that one would expect for its bandgap [1]. in relation to that 

observed in GaAs, especially considering that the surface recombination velocity of native n+ -lnP 

surfaces are comparable to those at high-quality GaAs-AIGaAs interfaces in GaAs solar cells. 

This discrepancy has not been resolved [1]. We believe that the high n-type doping level (-high 

1018 cm-3 to low 1019 cm-3) used in these cells probably result in low radiative lifetimes, leading 

to lower-than-ideal Voc of cells. A lower emitter doping, in conjunction with a thicker emitter (like 

0.1 11m) to maintain low emitter sheet resistance, is worth an exploration. Recently at RTI , we 

have been able to improve the Voc of the InP cells from above-mentioned 0.843 V to 0.875 V, 

one of the highest-reported values for InP cells under AMO conditions, with the use of reduced 

doping levels in the emitter. 

It may be worth pointing out that the Jsc levels in n+ -p InP cells are less sensitive to any 

reduction in lifetime (from higher doping level) in the emitter-region as the combination of low 

surface recombination velocity of n+ -lnP and reasonable diffusion lengths can cause nearcomplete collection of photo-generated carriers from the emitter. This was evident in our deepjunction cells as well, where, the variation of Voc was present in cells with very similar Jsc values. 

This effect was noticeable in both the orientations of InP substrates. 

46 

J 
SUMMARY 

In summary, we have developed and characterized the performance of deep-junction 

(-0.25 I..lm), graded-emitter-doped, n+-p InP solar cells, grown by metallorganic chemical vapor 

deposition (MOCVD). A novel, diffusion-transport process for obtaining lightly-doped p-type base 

regions of the solar cell, was described. The best active-area AMO efficiency for these deepjunction cells on (100)-oriented InP substrates is 16.8%, with a J sc of 31.8 mAlcm2, a Voc of 

0.843 V, and a fill factor of 0.85. By comparison, the best cell efficiency on the (111 )B-oriented 

InP substrates was 15.0%. The cell performance and breakdown voltage characteristics of a 

batch of 20 cells on each of the orientation were presented, indicating the superior breakdown 

voltage properties of InP cells on the (111)B orientation. Spectral response, dark I-V data, and 

photoluminescence (PL) measurements on the InP celis were presented with an analysis on the 

variation in V oc of the cells. It is observed, under open-circuit conditions, that lower-Voc cells 

exhibit higher band-edge PL intensity for both the (100) and (111)B orientations. This behaviour 

suggests that radiative recombination in the heavily-doped n+ -lnP emitter may be detrimental to 

achieving higher Vae in n+-p InP solar cells. 

47 

--- --- .. -. - -------
REFERENCES 

1) C.J. Keavney, V.E. Haven, and S.M. Vernon , Proc. of 21 st IEEE Photovoltaic Specialists 

Cont., (IEEE, NY), 141 (1990). 

2) M. Yamaguchi and K. Ando, J. Appl. Phys., 63, 5555 (1988). 

3) R.J . Walters, G.P. Summers, M.L. Timmons, R. Venkatasubramanian, J.A. Hancock, 

and J.S. Hills, Paper Presented at this SPRAT Conference. 

4) R. Venkatasubramanian and S.K. Ghandhi, Unpublished Results of work on low p-type 

doping of GaAs at Rensselaer Polytechnic Insitute, Troy, NY, 1983. 

5) R. K. Ahrenkiel in Properties of Indium Phosphide, emis Datareviews Series No.6, 

INSPEC, p.80, 1991. 

6) S.K. Ghandhi, VLSI Fabrication Principles - Silicon and Gallium Arsenide, John Wiley, 

NY, p.13, 1983. 

7) M.H. Lee and S.M. Sze, Solid St. Electron. 23, 1007 (1980). 

Doping 

Level 

(cm-3) 

17 18 

7x10 -3x10 

18 

1 x 10 

5x1018 

n + InP 

p InP 

+ 

p InP 

p+ InP 

AuGe/Ni/Au 

Emitter 

Base 

BSF 

Substrate 

Thickness 

(11m) 

0.126 

0.063 

0.25 

3.0 

0.25 

Ti/Au 

Fig. 1 Schematic of a n+-p InP solar cell structure indicating the graded emitter doping. 

48 
20~------------------------------------~ 

19 

p) 18 'E 

.3-

z 

.3 17 

16 

19 

1 ccrn, DEZn (Conventional) 

--------------------- ",' 

____ ~~~~ 

10 ccrn, DMZn (Diffusion) ____________ j18 

10 ccrn, DEZn (Diffusion) 17 .................................................................. 

16 

15L-______ ~L-________ L_ ________ ~ ______ ~15 

o 2 3 4 

X (I1rn) 

Fig. 2 Polaron doping profiles for (a) conventional Zn-doping with 1 cern H2 bubbled through a 

DEZn bubbler, (b) diffusion-transport with 10ccm H2 using the vapor pressure of DMZn and (c) 

diffusion-transport with 10ccm H2 using the vapor pressure of DEZn. 

~ ~ 

>. 

() 

c 

Q) 

·u 

:t= 

w 

E 

:J 

C 

Cd 

:J 

a 

Q) 

> 

-ca 

ai 

cr: 

100r-------------------~----------~ 

80 

60 

40 

20 

OL-~ __ ~ __ ~ __ ~ __ ~ __ _L __ _L~LL __ ~ 

200 300 400 500 600 700 800 900 1000 1100 

Wavelength (nm) 

Fig. 3. External quantum efficiency of a deep-junction, n+-p InP cell measured at NREL. 

49 
EFFICIENCY COMPARISON 

ON <100> AND <111> SUBSTRATES 

10 ~------------------------------------------------------~ 

(j) 

....J 

....J 

8 

W 6 

o 

l1.. 

o 

0: 

W 

[Il 4 

~ 

:::l 

Z 

2 

o 

, _ <1 00> ~ <11 1>' 

8 

12-12.9 13-13.9 14-1 4.9 15-15.9 16-16.9 

AMO CELL EFFICIENCY (%) 

Fig. 4 Cell (efficiency) performance histograms for (100) and (111 )-6 InP substrate orientations. 

BREAKDOWN VOLTAGE COMPARISON 

20 CELLS ON <100> AND <111 > 

12 ,--------------------------------------------------------, 

, _ <100> ~ <111> ' 

10 ..... . ..... ...... ..... .... .... ... . .. . ....... . ..... . . ...... . ...... . .. ...... . .. .. . ..... .... .. . . 

~ 8 ....J 

W 

o 

l1.. 

o 6 

0: 

W 

[Il 

2 

:::l 4 Z 

2 

o 

0-.9 1.0-1.9 2 .0-2.9 3.0-3.9 4.0-4.9 

CELL BREAKDOWN VOLTAGE (VOLTS) 

>5.0 

fig. 5 Cell (reverse breakdown voltage) performance histograms for (100) and (111 )-6 InP 

substrate orientations. 

50 
I ~----

< 100> InP 10~------------------------------------' 

8 #1-1955-a-4 

S = 2.5 mV 

........ Voe = 0.843 V 

::::l 

~ 6 

~ . iii #1-1961-a-4 c 

Cl> S = 10 mV 

.f: - 4 Voe == 0.783 V -l 

Cl.. 

2 

10,200 

Wavelength (A) 

Fig.' 6. 300K PL data (under open-circuit conditions) from two InP cells on (100)-oriented 

substrates, with nearly identical Jsc values, but different V oc values. 

::::l 

~ 

~ 

'iii 

c 

Cl> 

C 

-l 

Cl.. 

10.-----------------------------------~ 

8 

6 

4 

2 

#1-1957-b-7 

S = 250 mV 

Voc = 0.79 V 

Voe == 0.815 V 

O~ ______ _L ________ ~ ______ __l ________ ~ 

8,000 10,200 

Wavelength (A) 

Fig. 7. 300K PL data (under open-circuit conditions) from two InP cells on (111 )B-oriented 

substartes, with nearly identical J sc values, but different V oc values. 

51 
397?3~ 

N95- 20508 

CHEMICAL BEAM EPITAXY FOR HIGH EFFICIENCY PHOTOVOL TAlC DEVICES 

A. Bensaoula, A. Freundlich, M.F. Vilela, N. Medelci, and P. Renaud 

University of Houston 

Houston, Texas 

INTRODUCTION 

InP-based multijunction tandem solar cells show great promise for high conversion 

efficiency (1\) and high radiation resistance. InP and its related temary and quatemary compound 

semiconductors such as InGaAs and InGaAsP offer desirable combinations of energy bandgap 

values which are very suitable for multijunction tandem solar cell applications. The monolithically 

integrated InPllno.53Gao.47As tandem solar cells are expected to reach efficiencies above 30%. 

Wanlass et al. (ref. 1) have reported AMO efficiencies as high as 20.1 % for two terminal cells 

fabricated using atmospheric-pressure metalorganic vapor phase epitaxy (APMOVPE). The main 

limitations in their technique are first related to the degradation of the intercell ohmic contact 

(IOC), in this case the Ino.53Gao.47As tunnel junction during the growth of the top InP subcell 

structure, and second to the current matching, often limited by the Ino.53Gao.47As bottom subcell. 

Chemical beam epitaxy (CBE) has been shown to allow the growth of high quality 

materials with reproducible complex compositional and doping profiles. The main advantage of 

CBE compared to metalorganic chemical vapor deposition (MOCVD), the most popular 

technique for InP-based photovoltaic device fabrication, is the ability to grow high purity 

epilayers at much lower temperatures (450·C - 530·C) (ref.2). In a recent report Yamaguchi et al 

(ref.3) have shown that cost-wise CBE is a breakthrough technology for photovoltaic (PV) solar 

energy applications. Through the research effort undertaken in our laboratory, we have seen a 

rapid progress in the energy conversion efficiency of InP-based solar cells fabricated using 

chemical beam epitaxy (ref.4). This communication summarizes our recent results on PV 

devices and demonstrates the strength of this new technology. 

GROWTH TECHNIQUE 

The realization of high performance solar cells and tunnel junctions requires extremely 

high purity layers, perfect control on hole and electron concentrations as well as low 

interdiffusion of dopant species across the junction during the device growth procedure. 

Therefore, a relatively low temperature growth process and a good control of the interface 

properties are required. CBE meets the above requirements since high quality GaAs, InP, GaP, 

InAsP, GainP and InGaAs layers can be grown at lower temperatures than those used in more 

conventional techniques such as LPE, MOCVD, and even MBE. Moreover, it combines features 

from both MOCVD and MBE, allowing the growth of semiconductor heterostructures with 

monolayer abruptness and thickness control as well as easy multiwafer scale-up. 

We have already demonstrated, using CBE, the growth of a variety of high quality 

heterostructures ranging from lattice matched (InGaAsllnP) and strained (InAsPllnP) to highly 

stained InAslinP (ref. 5). In these studies the effect of growth temperatures and interruption 

schemes at the interfaces was demonstrated to be crucial. To illustrate the growth control 

capability of CBE we mention here our group latest results in the fabrication of perfectly 

balanced highly strained heterostructures GaP(under tension)/GaAsllnP (under compression) 

superlattices with thicknesses up to 1 Ilm (ref.6). These heterostructures are highly ordered and 

show very low defect densities as seen in Figure 1. High resolution x-ray diffraction (400) from 

these layers exhibit a residual strain of less than 8x10·5

; by comparison residual strain in 

GaAS/AIAs structures is 1.38x 10-3

. 

CBE's flexibility in allowing the growth of such high quality complex heterostructures 

pemnits the implementation of novel PV design concepts with expected higher efficiencies at no 

added fabrication costs (sometimes lower). As an example we are investigating the use of 

internal Bragg reflectors and the addition of MaWs in a standard diode structure to enhance the 

photon absorption and increase the cell photocurrent, thus significantly improving the PV 

53 
efficiency. Preliminary results on adding MaWs in the intrinsic region of a P-I-N structures are 

very encouraging and will be published elsewhere. 

All our Epitaxial growth runs were accomplished in a Riber CBE 32 system using 

trimethyl-indium (TMI), triethyl-gallium (TEG) and pre-cracked arsine (AsH3> and phosphine 

(PH3) as growth precursors. While research to date has included the growth of InP-based PV 

devices on both GaAs and GaAslSi, we will restrict ourselves in this paper to homoepitaxial 

growths on InP substrates (mainly InP:S (100». Reflection high energy electron diffraction 

(RHEED) was used to monitor the surface morphology during the growth process as well as 

calibrate an optical pyrometer focused on the substrate. This calibration uses the oxyde 

desorption temperature (517 oc) as the reference point. When this temperature is reached, the 

RHEED pattem changes from diffuse bulk type spots to clear bright lines. Due to the strong 

dependence of the InGaAs alloy composition on the substrate temperature, lattice matching with 

re~rect to InP was studied in the 450-530 °c temperature range. Lattice matching to better than 

10 ,as checked by High Resolution Double Crystal X-ray Diffraction, was achieved reproducibly, 

demonstrating excellent composition control for our ternary compounds. 

DOPING STUDIES 

Both the Be and Si doping behavior of InGaAs layers fabricated on InP:Fe (100) semiinsulating substrates at relatively low temperatures (450 °c to 480 OC) were investigated. Be and 

Si dopings were achieved using solid source effusion cells. Doping studies were carried out in 

lattice matched conditions. Be doped p-lnGaAs with net hole concentrations (as determined from 

17 -3 20-3 

Hall measurements) varying from 2x10 em to 2x10 em were achieved. No surface 

20 -3 

degradation was observed, even at dopina levels as high as 2x10 em. Si doping was 

1'7 -3 19 -3 

investigated within a doping range of 10 cm to 2x10 cm . No noticeable electrical 

compensation was detected and good reproducibility was achieved from run to run. 

In order to investigate the cross-diffusion behavior and to further assess the cross-doping 

profiles, Secondary Ion Mass Spectroscopy (SIMS) experiments were carried out on InGaAs 

multilayered structures fabricated with a combination of different Be and Si doping levels and + 

undoped spacers. SIMS experiments were performed in a Cameca IMS 4F system and both 0 + 

and Cs· profiles were taken for Be and Si respectively. Figure 2 represents compilation of 0 

16 -3 

and es- SIMS data. As it can be observed, the memory effects are in the range of 10 cm or 

less for both Be and Si doping. Furthermore, at the interface between Be and Si doped layers, 

observed cross-diffusions are within the experimental error. The sharp diffusion profiles between 

Be and Si are consistent with solar cells and tunnel junctions low inter-diffusion requirements. 

Our results indicate that pin or nip structure should exhibit similar behavior. We have chosen 

deliberately to study pin type devices. 

SOLAR CELL AND TUNNEL DIODE DEVICES 

Figure 3 shows a diagram of a tandem solar cell structure. Besides achieving the proper 

electrical and structural quality for the individual layers, two issues stand out in the realization of 

these devices. The first relates to the fabrication of thin tunnel junctions with high current 

carrying capabilities and minimum resistivities. The second concerns the ability to grow a high 

quality InP solar cell on InGaAs at low enough temperature so as to preserve the characteristics 

of the underlying tunnel junction. Following we will describe the growth and fabrication of the 

individual tandem components. 

InO.53GaO.47As solar cell 

The cell fabrication process begins with the CBE growth of a structure as shown in 

Figure 3 (bottom cell). A summary of the characteristics under natural sunlight and without anti54 

I 

I 

J 
reflection coating is: Jsc=30mNcm2; Voc=0.27V; FF=57.8%; "11=5.1%, for an incident sunlight 

power of 91 .4mW/cm2. A double anti-reflection coating (ZnS/MgFv was deposited and the new 

electrical characteristics under natural sunlight are (Figure 4): Jsc=60mNcm2; Voc=0.295V; 

FF=54.2%; "11=10.2%, for an incident sunlight power of 94 .2mW/cm2 (total area, 0.25cm2). In 

this particular sample we note that a strong increase in the photocurrent after anti-reflection 

coating is observed. This increase cannot be explained solely by the addition of the antireflection layer. It is believed that this diode had a much lower shunt resistance before the 

dielectric deposition. A leakage current prior ARC coating was responsible for the lower 

photocurrent measured. Similarly after ARC the fill factor (FF) is smaller, a result of a higher 

series resistance. 

The Spectral Response (SR) characteristics after ARC of this InGaAs cell are presented 

in Figure 5. A minimum at 680nm is clearly observed. This minimum is related to reflections at 

the InP window layer (we remind the reader that the ARC is optimized for the InGaAs material). 

We note that for wavelengths higher than 1100nm, the SR decreases noticeably, a result of a 

non optimized base layer. Both the base layer thickness and the minority carrier lifetime in this 

layer must be investigated to better understand the cause for this decrease. The high response 

at smaller wavelengths shows however the high quality of the emitter layer. In conclusion, while 

further improvements in the cell characteristics are expected through a better optimization of the 

base layer, the high photocurrent density already demonstrated in these cells (60mAlcm2, the 

highest ever reported) is very promising. 

The dark I-V and short circuit current (lsC> as a function of the open circuit voltage {VoC> 

characteristics have been measured as well. The shunt resistance (Rsh) from dark I-V, the diode 

ideality constant (n) from Isc-Voc and the saturation current (JS> from Isc-Voc were calculated. 

These values are shown in Table I. An ideality constant of 2.2 is expected for a small bandgap 

diode such as InGaAs. Both the shunt resistance (42 o*cm2) and the saturation current (Js = 0.6 

mAlcm2) can be improved through proper mesa passivation. 

InP solar cell 

Figure 3 shows the doping levels and thicknesses for the InP solar cell (top cell). Its 

electrical characteristics are shown in Figure 6. The characterization results performed under 

natural sunlight without anti reflection coating are: Jsc = 20 mNcm2, Voc = 0.82 V, FF = 77%, 

and "11=13%, for an incident sunlight power of 97 mW/cm2. After the antireflection coating this 

solar cell presents 18% conversion efficiency under solar simulator. Its spectral response is 

shown in Figure 5. The flat SR from 450nm to 850nm shows the high quality of this material. The 

lower efficiency at smaller wavelengths «400nm) is expected since this InP solar cell does not 

have any window layer. By using a window layer or decreaSing the emitter thickness it is possible 

to improve the SR at lower wavelengths. In Table-1 the Rsh' Is and n for this solar cell are given. 

It is expected that Rsh and the InP solar cell effiCiency will increase after mesa passivation. A 

mesa passivation and a window layer should further increase the efficiency to 22%. 

InO.53Ga0.4 7As tunnel junction 

p+/n+ InO.53GaO.47As tunnel junction structures with different active layer 

thicknesses (ranging from 0.2 to 0.8 IJ.m) and growth temperatures (450-530 oC) were grown. 

De\(ices were made using a standard wet etching process with mesa openings of 100 and 200 

IJ.m in diameter. Evaporated Au was used for front and back contacts. No high temperature 

annealing of the metallic contact layers was performed. 

55 

J 
The net carrier concentrations as extrapolated from Hall measurement data for devices 

discussed here were varied from NA=1.3x1019 to NA=5x1019 cm-3 in the p-type layers, and 

from NO=7x1 018 to NO=1.3x1 019 cm-3 in the n-type layers. Thus, junctions with highly 

degenerate n type sides (> 10 kT) and degenerate p type sides (> kT) were obtained, resulting in 

devices with effective dopings N*=NANo'(NA+NO) ranging from 5.9x1018 to 8.3x1018 em-3. 

All our devices exhibit excellent I-V characteristics. Peak current densities obtained in 

this work are among the highest ever reported for epitaxial tunnel junctions. The best room 

temperature peak current density exceeds 1000 Ncm2 (Figure 7); and specific resistivities lower 

than 10-4 n.em2 were measured. Maximum resistivities Rmax (peak current to peak voltage 

ratios) are in the 2x10-4 to 5x10-4 n.cm2 range, making these junctions highly suitable as IOCs 

for the InPllnGaAs tandem solar cell. High peak to valley ratios are exhibited by many devices 

with room temperature peak to valley ratios of 9 and peak current densities greater than 550 

Alcrn2 . 

In order to investigate the evolution of the tunnel junction properties when incorporated 

in the InP/lnGaAs tandem solar cell structure, an InGaAs tunnel junction with an effective doping 

of 7x1 018 cm-3 was subjected to the growth of a thick (>3 Ilm) InP solar cell. Prior to the ~olar 

cell growth, the tunnel junction coated wafer was cut in two pieces. One piece was processe~ as 

a reference sample. The InP solar cell regrowth structure was realized with standard parameters 

(doping and layer thicknesses) on the second piece of the wafer (ref. 7). Following the growth 

process, the InP solar cell was selectively etched using a HCI solution and mesa processing was 

accomplished. Even after more than 2 hours at 560 °C, no degradation of the tunnel junction 

characteristics is observed. The device still exhibited very high peak current densities, up to 860 

Ncm2 as shown in Figure 8. Hence, our tunnel junctions are suitable for use in tandem 

structures either in a planar or the more demanding patterned electrical interconnect concept 

(ref. 8). With the above characteristics (tunneling currents and resistivities), the voltage drop 

accross a patterned tunnel junction aligned with the top solar cell grid (e.g. 5% shadowing) at 

concentration of 100 x AMO will be below 10-2 volts. 

InPllnGaAs tandem solar cell with planar tunnel diode 

A complete structure similar to that shown in Figure 3 was grown. In the following case 

the IOC for the tandem is realized with a planar tunnel diode. A very low photocurrent output for 

these type of structures under natural sun light illumination is characteristic. In Figure 9 we 

shown the electrical characteristics of this tandem under concentrated light stimulation. The 

tandem's Voc is equal to the sum of that from the InGaAs botton cell (Vrx1 ) and that from the 

InP top cell (Voc2) - Voc = Vrx1 + Vrx2 =1.2 V)-. The SR was perfonned in order to investigate its 

low photo sensitivity (Figure 10). Despite the low overall photocurrent ouput, the InP top cell has 

a relatively good response. While the tandem responds fairly well for wavelenghts less than 920 

nm, for photons with wavelengths between 920nm and 1650nm the SR is near zero. As expected 

the 0.21lm thick IOC tunnel diode absorbs nearly all the incoming photons thus totally blocking 

the irradiation of the bottom cell. The implications of this result is either the adoption of a tandem 

cell structure using a patterned tunnel concept, as proposed by Shen et al (ref. 8), or the use of 

ultra thin (-400A) planar tunnel diodes. Both of these concepts are possible using CBE and are 

currently being investigated. 

InPllnGaAs tandem solar cell with patterned tunnel diode 

A structure similar to that of Figure 3 was grown but this time in a two-step process. First, 

the growth of the InGaAs solar cell followed by the InGaAs tunnel diode was perfonned. The 

sample was removed from the growth chamber and the tunnel diode was pattemed with the 

same mask as the contact grid. Then, the sample was reloaded into the CBE chamber and the 

S6 
InP solar cell was grown on top. Figure 10 shows the SR (after anti-reflection deposition) of this 

tandem. We see a noticeable improvement in the tandem photoresponse. The tunnel diode 

patterning has allowed a much higher illumination of the bottom cell. The InGaAs bottom cell 

however displays a higher SR compared to that of the InP top cell. This is mainly due to 

inhomogeneous etching during the patteming step and regrowth issues resulting in an 

inhomogeneous and pOor quality InP material in the top cell. Nevertheless under light 

stimulations the patterned tunnel diode shows low resistivity and the tandem I-V characteristics 

display good ohmic IOC behavior even for high polarization. It is clear that regrowth procedures 

optimization will be necessary. 

CONCLUSION 

Chemical beam epitaxy has been shown to allow the attainment of high quality InGaAs 

and InP layers. InP and InGaAs solar cells have been obtained with high efficiency 

photoconversion, 10.2% for InGaAs solar cell and 18% for InP solar cell. The tunnel diodes 

obtained with this technique have very high peak current densities and show to be thermically 

stable during the top cell growth. This realization has permitted the fabrication of monolithically 

integrated tandem solar cells. The planar tandem device exhibits an open-circuit voltage equal 

to the sum of that of the individual sub-cells Noc = 1.2 V) demonstrating negligible voltage drop 

at the interconnect, but the InGaAs tunnel diode was shown to absorb highly thus blocking 

irradiation of the InGaAs bottom cell. By using pattemed tunnel junctions we observe a 

significant improvement in the tandem photoresponse, but a lower spectral response for the top 

InP cell. Work underway is focusing on optimization of the regrowth procedures to address this 

problem. 

** Present address LPSES - CNRS, Va/bonne - 06560 - France 

REFERENCES 

1. M. W. Wanlass, J. S. Ward, K. A. Emery and T. J. Coutts; Proc. 23rd IEEE -PV Conf.; 1993; 

pp 621. 

2. A. Freundlich, M. F. Vilela, A. Bensaoula, and N. Medelci; Proc. 23rd IEEE -PV Cont.; 1993; 

pp 644. 

3. M. Yamaguchi, T. Warabisako and H. Sugiura, Joumal of Crystal Growth; 136; 1994; pp.29. 

4. A. Bensaoula, N. Medelci, M. F. Vilela and A. Freundlich; 6th Intemational Conference on InP 

and Related Materials; Santa Barbara - CA - USA; IEEE catalog #94CH3369-6, (1994), pp.280. 

5. V. Rossignol, A. Bensaoula, A. Freundlich, and A. H. Bensaoula; 5th Intemational Conference 

on InP and Related Materials; Paris - France; IEEE catalog #93CH3276-3, (1993), pp.127. 

6. A. H. Bensaoula, A. Freundlich, A. Bensaoula, V. ROSSignol and A. Ponchet; J. Vac. Sci. 

Techno!. B 12(2),1994, pp. 1110. 

7. A. Freundlich, M. F. Vilela, A. Bensaoula, and N. Medelci, Proc. 23th IEEE-PV Cont., 1993, 

pp.644. 

8. C. C. Shen, P. T. Chang and K. A. Emery; Proc. 22nd IEEE -PV Cont.; 1991; pp 381. 

*This work was supported by the state of Texas advanced technology program 93-03652-260 , 

93-003652-236 and 91-03652-243 and advanced research program 93- 003652-224 and partially 

supported by 1ST Inc. The authors would like to thank A. H. Bensaoula, E. Kim and V. Rossignol 

for their help in this project. 

57 
r- . - -

• - .GoP 15100U 

GoA. ,27"'U 

G.aP tSUU 

TABLES 

Table L - Shunt Resistance (Rsh)' Saturation Current (JS> and Ideality Constant (n) for the solar 

cells studied in this work 

Sample InGaAs InP 

Rsh 4.2x1 0-5 0.25 

(Mo*cm2) 

Js 0.6 8x10-9 

(mNcm2) 

n 2.2 1.65 

FIGURES 

(a) 

Patterned Planar 

tandem tandem 

1.2x10-3 2 

0.7 8x10-4 

- -

10~ (004)C.A. 

Ii, 103 

.f' 102 

'" c 

~ 

c 

-7200 

~ p<riod 0 

-I 

2 

-2 .' 

4 

o 3600 

Diffraction Angle (Arcst'c) 

(b) 

Fig. 1 - GaP/GaAs/lnP superlattice. a 50-period GaP(5ML)/GaAs(32ML)/lnP(5ML)/GaAs(27ML). 

(a) cross-sectional TEM and (b)High resolution x-ray diffraction pattems (ref. 6). 

58 

-----------
Dopant cross diffusion 

I "0 

~ 1 "0 

X- 1. 1. X- 0 0 0 '"Q 0 "C 0 -8 0 "C C "0 C "0 "0 C :s .. :s (i; cZ (i; :s CQ 

1()20 

-.... 

E 

v 1019 --~ ~-- -, I 

E I 

0 I 

- 1018 I ~ I '-' 

t: I 

.g I 

-~ 1017 

-

"'" ", e Be ~ 

v C~ I . . 

e 1016 \ I 

0 , 

(,) "'Sf" ,,'OJ '\' I 

0 1 1 

Depth (microns) 

Fig. 2 - Secondary Ion Mass Spectroscopy of alternated dopant layers (Be and Si) and undoped 

spacers. 

Top cell 

Tunnel 

Junction 

Bottom cell 

~ 

"'-

,/" 

"-

,/ 

" 

InGaAs:Be 

Cap layer 

InP:Be 

Emitter 

InP:Si 

Base 

InP:Si 

BSF 

InGaAs:Si 

InGaAs:Be 

InP:Be 

Window 

InGaAs:Be 

Emitter 

InGaAs:Si 

Base 

InGaAs:Si 

BSF 

InP:Si 

Buffer 

InP:S 

Substrate 

0.5-1 x 1020 cm-3, 0.5 ~m 

1-3 x 1018 cm-3, 0.05-0.1 ~m 

0.5-1 x 1017 cm-3, 1-2 ~m 

1-5 x 1018 cm-3, 0.2-0.5 ~m 

1-5 x 1018 cm-3, 0.1 ~m 

1-5 x 1019 cm-3, 0.1 ~m 

3-5 x 1018 cm-3, 0.05-0.1 ~m 

1-3 x 1018 cm-3, 0.2-0.5 ~m 

0.5-1 x 10 cm-3, 2-3 ~m 

1-5 x 1018 cm-3, 0.2-0.5 ~m 

1-5 x 1018 cm-3, 0.5-1 ~m 

2 x 1018 cm-3, 400 ~m 

Fig. 3 - InPllnGaAs tandem solar cell structure. Doping levels and thicknesses are shown. 

59 
-l ·-t -- t- 0-

- , 

Voc- 295mV . ' / ' J~ = 60 mAlcm2 ;- ;- -

_.-J-""'" FF = ~ 2" '- "1 .. -

E!!iciency = 10 _ < _ ~ 

area = a 25 cm2 

::t .__ .. _ -._. 

U 

~_t--~-t-c"::.- - .- -- -

· '00 100 200 

Voltage (mY) 

Figure 4 : Electrical characteristics of the 

InGaAs solar cell under natural sunlight, 

with anti reflection coating. 

3~~~.-~~~-r~-r~-r~-' 

2.= :..-=- L-= =:=:=:=r=:=: = =.rb ~ , = =:..: =:=1=:.=:=:=: = = = - o 1-' _ .. -:- - < - '- : - r- ' -. --, - -- -

<" 

::::-

E -, --

_ 

-'-

-

_._ '- Jsc 

Voc 

= 

= 

20 

820 

mAlcm2

mV --

_

_. . 

_. -' 

_ . 

-- - -

_ _

1--. _ 

~ .2 - ~- .. - FF = 77% _ . - -- __ 

§; - - - Etfic,ency = , 3~ -' - - ---

u-3 ~ =:= . rea=025 cm2 =: =: -:' _. _ -4 - - - ---'- _._-- - ---

- --.- - ,-.-.- .- . - - -.,.- -5 -- ~--+-------~--=-

-200 a 200 .aD 600 800 1000 

Voltage (my) 

Figure 6: Electrical characteristics of the InP 

solar cell under natural sunlight (without anti 

reflection coating). 

InP 

Solar Cell 

Removal 

oflnP 

Tunnel 

Diode 

: 0 

~ 

~ 0 

.. i 04 

1 

.. 

~ 0 

200 400 800 k)() 1000 ' 200 10400 1800 1800 

W.""leng1h (nm) 

Figure 5 : Spectral Response of the InGaAs 

solar cell (opened squares) and InP solar 

cell (solid circles). 

Figure 7: Room temperature I-V 

characteristics of InGaAs tunnel junction 

(mesa diameter = 200 )lm) with a peak 

current of 1015 Ncm2

. 

Figure 8: I-V characteristic of a InGa A~ 1'..Jnnel junction (mesa 200 )lm in diameter) after solar cell 

growth. J p = 860 Ncm2

, Rmax = 2.9 X 10-'1 O.cm2

. 

60 

J 
2 

! i 

I I J 

0 

1 ! i 

i' -2 - r--r ~ 

<.3 i i 

~ -+~ 

! 1 ! j I I ! ! ! I 1 

I 

Voe = 1.2 V (= Voel • Voe2i . -r- ,-I. _ ,-

Voel = InGaA. botton cell ' ! :' : 

. Voe2 = InP lop cell 

area = 0 05 cm2 -: T'} -,- :- . -; , ' j' - ,...... 

- , -I ...;.. ~L-+ +- ,- -t- i..J-. - ,.-

I i j 4 : ii I ;-d-~ ~ .. ~ 

~+-~-+~: ~~1 -+~~~~-r~ -200 0 200 ~ 600 600 1000 1200 1~ 

Voltage (mY) 

Figure 9 : Electrical characteristics of the 

tandem solar cell with a planar tunnel diode 

under concentrated light stimulation . 

61 

O+-~,-...--.-~-"---'-~r-""-'~--r--I 

200 ~ 600 600 l00012001~16001600 

Wavelength (nm) 

Figure 10 : Spectral Response of the 

InGaAs/lnP tandem solar cell with a planar 

tunnel diode (solid circles) and InGaAsllnP 

tandem solar cell with a pattemed tunnel 

diode (opened squares) _ 
--------------

3 Cf 7/"13 

N95-20509 

STATUS OF DIFFUSED JUNCTION p+n InP SOLAR CELLS FOR SPACE APPLICATIONS 

Mircea Faur 

Cleveland State University 

Cleveland, Ohio 

Maria Faur 

NYMA, Inc. 

Brook Park, Ohio 

D.J. Flood and D.J. Brinker 

NASA Lewis Research Center 

Cleveland, Ohio 

C. Goradia 

Cleveland State University 

Cleveland, Ohio 

N.S. Fatemi and P.P. Jenkins 

NYMA, Inc. 

Brook Park, Ohio 

D.M. Wilt and S. Bailey 

NASA Lewis Research Center 

Cleveland, Ohio 

M. Goradia and J. Moulot 

Cleveland State University 

Cleveland, Ohio 

ABSTRACT 

Recently, we have succeeded in fabricating diffused junction p+n(Cd,S) InP solar cells with measured 

AMO, 25 °C open circuit voltage (V oc) of 887.6 m V, which, to the best of our knowledge, is higher than 

previously reported Voc values for any InP homojunction solar cells. The experiment-based projected 

achievable efficiency of these cells using LEC grown substrates is 21.3 %. The maximum AMO, 25 0C 

efficiency recorded to date on bare cells is, however, only 13.2%. This is because of large external and 

internal losses due to non-optimized front grid design, antireflection (AR) coating and emitter thickness. This 

paper summarizes recent advances in the technology of fabrication of p + n InP diffused structures and solar 

cells, resulted from a study undertaken in an effort to increase the cell efficiency. The topics discussed in this 

paper include advances in: 1) the formation of thin p + InP:Cd emitter layers, 2) electroplated front contacts, 

3) surface passivation and 4) the design of a new native oxide/ A1203/MgF2 three layer AR coating using a 

chemically-grown P-rich passivating oxide as a first layer. Based on the high radiation resistance and the 

excellent post-irradiation annealing and recovery demonstrated in the early tests done to date, as well as the 

projected high efficiency and low-cost high-volume fabricability, these cells show a very good potential for 

space photovoltaic applications. 

INTRODUCTION 

Owing to their excellent radiation resistance and annealing properties (refs. 1, 2), InP solar cells hold 

great promise for space power applications. In 1990, within a few short years after the renewed interest in 

the development of InP solar cells began in 1985, driven by early reports from NIT, Japan (ref. 4), an 

AMO, 25 0C efficiency of 19.1 % was already achieved on a 4 cm2 InP cell (ref.5). These n+pp+(Si,Zn) 

InP cells, developed by SPIRE Corp., were fabricated by MOCVD growth of the active layers on a heavily 

doped p + -InP:Zn suhstrate. The achievement of such a high efficiency in such a short period of time is 

particularly important since no special efforts were made to reduce external losses associated with an 

unpassivated surface and with the use of a ZnS/MgF2 AR coating, developed for GaAs based solar cells. An 

unpassivated surface is suggested both by the low Voc value of 876 m V, and the low blue response for 

an emitter thickness of only 30 run. By passivating the front surface, the projected AMO, 250C efficiency of 

22 % for these cells seems very realistic. Given this unprecedented short term success in developing high 

efficiency InP solar cells, and the relatively large accepted potential of these cells for high radiation 

environment space solar cells applications, it is rather odd that research funds from the space agency have 

continuously declined after 1990. A possible explanation of this, besides shrinking research funds, is the high 

cost of presently developed all-epitaxial high efficiency InP solar cells. For small power applications, where, 

beside radiation resistance, high BOL efficiency is required, the use of all-epitaxial InP solar cells grown on 

63 
InP substrates might be justified . However, their high wafer and processing costs, prohibit their large scale 

use in space solar cell arrays. 

The cost of InP cells is due, to a significant extent, to the high cost of InP substrates. Hence, a large cost 

reduction might be achieved through heteroepitaxial InP solar cells grown on cheaper substrates such as Si or 

Ge. The increase in their AMO, 25 °C efficiency above presently achieved 7.1 % (ref.6) for Si and 9% (ref. 

7) for Ge substrates respectively is expected to come through the growth of GaAs based intermediate layers 

to accommodate the large lattice mismatch between InP and Si or Ge. Should these heteroepitaxial cells 

exhibit a BOL efficiency in excess of that of presently used Si space solar cells while retaining the radiation 

tolerance of homojunction InP solar cells, then, due to their lighter weight, lower substrate cost, and lower 

fragility, as compared to all-lnP solar cells, they would have a very good potential for large scale use in 

space solar arrays, provided that the processing costs could be kept within reasonable limits. 

A significant cost reduction over epitaxy can be achieved through the use of diffused junction InP cells. 

Until recently, the only InP solar cells fabricated by thermal diffusion were of the n+p configuration. Cells 

fabricated by closed ampoule diffusion have yielded maximum AMO, 25 0C efficiency of 16.6% for a (S ,Zn) 

cell (ref.8), 14.35% for non-optimized (S,Cd) cell (ref. 9) , and 15.2% for an open tube diffusion (S,Zn) cell 

(ref. 10). The n +p (S,Zn) InP cells made by closed ampoule diffusion have high radiation resistance (ref. 1), 

and independent studies (e.g. refs. 11, 12) have shown much higher annealing rates after irradiation, under 

cell operating conditions, than high efficiency n + + In + Iplp + + (Si ,Zn) all-MOCVD InP cells (ref.5). 

The drawback of the n+p (S ,Zn) diffused InP cells is that a large number of defects are present after 

diffusion both in the n + emitter and in the p-base, which, as shown above, makes the solar cell efficiency 

lower than that of all-MOCVD n+p (Si,Zn) InP cells. Since, neglecting surface effects, of all solar cell 

performance parameters, Voc can be regarded as the best measure of how low is the defect density within a 

given cell structure from among the different cell structures, its value can give useful information about the 

quality of each cell structure. For example, for the best n+p (S ,Zn) diffused cell with AMO, 25 0C 

efficiency of 16.6%, measured at NASA LeRC, the Voc was only 828 mV as compared to 876 mV for the 

higher AMO efficiency (19.1 %) of the MOCVD-grown cell. 

At the 11th SPRAT Conference, we predicted that for diffused solar cells, the p + n configuration has a 

higher potentially achievable maximum efficiency than the n +p configuration due especially to an increased 

Voc (ref. 13). The prediction was based on AMO, 25 °C Voc values of 860 m V we measured for p + n (Cd,S) 

InP solar cells as compared to experiment-based projected maximum V oc of only 840 m V for our n + p 

(S,Cd) InP cells. For our thermally diffused p+n and n+p structures, the ranking for projected maximum 

efficiency, in decreasing order, is: (1) p+n (Cd ,S); (2) n+p (S ,Cd); (3) n+p (S ,Zn), and (4) p+n (Zn,S). 

The large structural and electrical-type defect density we found in the emitter and the base of structures (2) to 

(4) as compared to structure (I), explains Voc, Isc , and 11 limitations of the last three structures. This also 

explains why, although a large experimental effort was made by NTT to improve V oc and efficiency of 

diffused n+p (S ,Zn) cells, the independently confirmed maximum efficiency was of only 16.6%. However, 

more recent work performed on NTT fabricated n + p cells (ref. 14) has shown that the AMO efficiency of 

one such cell could be increased from 14.8% to 17.5% by plasma hydrogenation at 150 0c. The large gain in 

11 is due to a sensitive increase in Jsc, which is thought to be due to a decrease in H2 and H3 trap center 

densities in the p-InP:Zn base, and the formation of a n + -p-p + structure. A small increase in Voc is 

thought to be due to a decrease by hydrogenation in surface and space charge recombination. For n + p 

structures, we found Cd-doped substrates to be inherently superior to Zn-doped substrates (ref.! 0), and the 

experiment based practically achievable AMO efficiency of diffused n + p (S ,Cd) InP solar cell using 

relatively large EPD (- 5xl04 cm-2) substrates is about 18.8% . 

A preliminary investigation of p + nand n + p diffused structures and solar cells, prior to and after 

irradiation with 1013 cm-2 3 MeV protons indicates that the same ranking should hold for structures (1), (2), 

and (4) with respect to radiation resistance as mentioned above for maximum efficiency. The p + emitter of 

p+n (Cd,S) InP diffused structures exhibits very low radiation induced carrier removal rates (ref. 16), which 

is thought to contribute to the increased radi ation resistance of these cells as compared to other InP cell 

structures for which 3 MeV proton irradiation studies are available. Another interesting characteristic of 

64 
---- ,--- - -

-------

these cells is a remarkable annealing property at room temperature in the dark, which might be very 

attractive for other applications such as long-life alpha- or beta-voltaic batteries. 

Recently, using a P-rich passivating layer grown by chemical oxidation, we have succeeded in fabricating 

diffused junction p + n(Cd,S) InP solar cells with measured AMO, 25 °C V oc of 887.6 m V (ref. 16), which to 

the best of our knowledge is higher than previously reported Voc values for any InP homojunction solar 

cells. Although the experiment based projected achievable efficiency of these cells using LEC grown 

substrates is 21.3% (ref.17), the maximum AMO, 25 0C efficiency recorded to date on bare cells is only 

13.2 %. This is because of large external and internal losses due to non-optimized front grid design, 

antireflection coating (ARC) and emitter thickness. This paper summarizes recent advances in the technology 

of fabrication of p + n InP diffused structures and solar cells, resulting from a study undertaken in an effort to 

increase the cell efficiency. These advances include: 1) the formation of thin p+ InP:Cd emitter layers, 2) 

electroplated front contacts, 3) surface passivation and 4) the design of a new native oxide/Al203/MgF2 

three layer AR coating using a chemically-grown P-rich passivating oxide as a first layer. The paper will 

also discuss some light instability problems associated with high Zn content front Au-Zn contacts and 

preliminary radiation resistance and post-irradiation annealing studies. 

EXPERIMENTAL 

Cd diffusion into n-InP:S (NO-N A = 3.5 xl016 to 3.1 xl017 cm-3) was performed by a closed ampoule 

technique using Cd3P2 as diffusion source (ref. 18). Diffusion temperatures were from 560 to 660 0C. The 

substrates were Czochralski (LEC) grown with EPD of about 5 x 104 cm-2. Diffusions were performed 

through either P-rich chemically grown (ref. 13) or MOVPE grown InGaAs cap layers. 

The surface quality of diffused samples was monitored by either Nomarski or SEM microscopy. 

Electrochemical techniques (ref.19) were used for step-by-step characterization of these diffused structures 

during fabrication and after irradiation with high energy protons using a Polaron profiler PN4200, 

manufactured by BIORAD. We have recently developed a new electrolyte, which we call "UNIEL", for ECV profiling of InP and GaAs based structures (ref. 19). In order to accurately locate the position of different 

defect levels derived from low frequency EG-V measurements, we have began, in parallel, 

photoluminescence measurements at 5 K on several Cd diffused samples. The excitation wavelength was 514 

nm at a power density of 160 mW/cm2. Luminescence was dispersed in a 1.26 meter spectrometer and 

detected with cooled CCD array. The system resolution was about 0.5 meV. 

Small area (0.48 cm2) p + n InP solar cells were fabricated using Zn- and Cd-diffused structures. Au was 

used for the back contact. The Au-Zn-Au (0.18 to 1.5 11m thick) front contact grid was deposited by 

evaporation and defined using existing photolitographical masks, designed for nip cell configurations. Au 

based contacts are known to penetrate into InP during sintering at 430 °C up to depths which are over three 

times the initial thickness of the evaporated Au-Zn-Au layer. Hence, we kept the thickness of the emitter at 

quite a high value (up to 5 11m) while keeping the thickness of the evaporated contacts below 0.2 11m. After 

sintering, the thick emitter was thinned down over the uncontacted area using a chemical etch (PNP), 

specially developed for this purpose (ref.20). Recently we were able to fabricate thin p + emitters using Cddiffusion, by using either thicker P-rich oxides or InGaAs cap layers. New front contact schemes, employing 

electroplated Au and Au-Zn developed for thin emitters will be discussed in the following paragraph. We 

also propose a new optimized three-layer ARC for InP solar cells, which uses a P-rich chemical oxide as a 

first layer in a In(p03)3/AI203/MgF2 structure. This P-rich chemical oxide, which is described in more 

detail elsewhere (ref. 16), is primarily designed as a surface passivation layer. 

Unless otherwise mentioned, the performances of solar cells in this paper refer to non optimized single 

layer ARC using only the passivating layer. Dark and illuminated solar cell performances and their variation 

with illumination time or temperature were recorded at CSU using a computer controlled facility and an ELH 

lamp (assumed AM1.5 spectrum) as the light source. For selected cells, dark and illuminated I-V 

characteristics, and their variation with time, as well as reflectivity and spectral response measurements were 

performed at NASA LeRC. 

65 
RESULTS AND DISCUSSION 

At the last SPRAT conference we reported an AMO, 25 oc, Voc value of 880.3 mV for a diffused p+n 

(Cd,S) InP solar cell. Since the maximum efficiency was only 12.57% as compared to projected achievable 

maximum efficiency of 21.3 %, over the last year we have concentrated our efforts to reduce the large 

external and internal losses due to non optimized front grid contacts, AR coatings and emitter thickness, 

while further improving the diffused structure quality, by reducing the defect density in the emitter and 

junction area. 

Since, of all solar cell parameters, Voc can be regarded as the best measure of how low is the defect 

density within a given structure from among the different cell structures, its value can give useful information 

about the quality of each cell structure. Recently, using a P-rich passivating layer grown by chemical 

oxidation, we have succeeded in fabricating diffused junction p + n(Cd,S) InP solar cells with measured AMO, 

25 0C Voc of 887.6 mV (Fig.!), which to the best of our knowledge is higher that previously reported Voc 

values for any InP homojunction solar cells. The achievement of such a high V oc value for a diffused 

junction cell with no AR coating, except for the thin passivating layer, is remarkable if one takes into 

consideration that the InP:S LEC grown substrates used have had a rather large EPD of 5-7 x 104 cm-2 . 

From the dark I-V characteristic (Fig.2), the dark saturation current density 10 (A= 1) has a record low value 

of 1.38 x 10-17 A/cm2, which explains the high Voc value. 

The cell was fabricated by thinning the emitter from its initial thickness of about 4.5 11m to about 0.45 

11m, after sintering the Au-Zn front contact. The relatively low short circuit current density (Jsc) of 26.3 

mA/cm2 and the low external quantum efficiency (EQY) of this cell can be explained by the large 

thickness of the emitter (0.45 11m, compared to the optimum thickness of - 0.25 11m) and the absence of an 

AR coating. Figure 3 shows the EQY of this cell. Worth noticing is the relatively high blue response for a 

p + emitter as thick as 0.45 11m, suggesting a large diffusion length in the Cd doped emitter and a well 

passivated surface. The cell had a high Rs of about 3 .5 Q-cm2 due to relatively high contact and sheet 

resistance and non optimized front grid design (an existing photolitographical mask designed for nip cell 

configurations was used), resulting in low FF of 69 % and efficiency of onl y 11.98 %. 

Next paragraphs summarize recent advances in p + n InP diffused structures and solar cell technology, 

undertaken in an effort to increase the cell efficiency. This includes the formation of thin p+ -lnP:Cd emitter 

layers, electroplated front contacts, surface passivation and the design of a new native oxide/ A1203/MgF2 

three layer AR coating structure using a chemically-grown P-rich passivating oxide as a first layer. Also 

discussed are some light instability problems associated with high Zn content front Au-Zn contacts and 

preliminary radiation resistance and post-irradiation annealing studies. 

Emitter Layer 

As mentioned above, for cells such as that shown in Fig.l, thick emitters have been used. This adds a 

troublesome fabrication step, which affects the reproducibility of cell performances. Furthermore, as seen 

above, using thick emitters such as in Fig.4, thinning from over 4 11m to below 0.5 11m has the drawback of 

reducing the surface hole concentration in the thinned emitter, thereby increasing the series resistance (Rs) 

and lowering the fill factor (FF). A more step-like diffusion profile such as in Fig.5, was possible by 

optimizing the thickness of the In(P03)3-rich chemical oxide (- 10 nm), used as a diffusion cap layer. To 

further decrease the Rs values of these cells, two avenues were undertaken : 

(a) Since the optimal emitter thickness is estimated at about 0.25 to 0.3 11m, we were able to 

fabricate thin emitters while maintaining the high surface acceptor concentration (see Fig.6). 

(b) Use low band gap diffusion cap layers such as InGaAs. Example: The EC-V diffusion profile 

in Fig.7(a) was realized using an intentionally undoped InGaAs cap layer. Curve (b) refers to 

a diffusion carried out under similar conditions (600 °C, 25 min., and same amount of Cd3P2 

source material) using a thin In(P03)3-rich chemical oxide cap layer. As expected, the 

InGaAs cap, after Cd diffusion becomes highly doped, which should improve the contact 

66 
.

resistance. However, in this case, since InGaAs cap is not stable at higher temperatures, the 

hole concentration in the InP:Cd emitter is rather low, which should increase the sheet 

resistance. 

Electroplated Front Contacts. 

For thin emitters (0.5 to 0.75 j.lm) we first tried thin (0.1 j.lm) Au-Zn evaporated contacts, with an intent 

to then deposit thicker electroplated contacts after sintering. However, after sintering the contacts at 430 oC, 

for 2 minutes, the contacts penetrated at depths greater than the emitter thickness, short-circuiting it. For 

lower sintering temperatures the contacts lifted during subsequent chemical treatments in PNP etch (ref.20) 

we are using both for surface passivation and thinning the emitter. For these samples, after removing the 

evaporated contacts, Au-Zn and Au-Cd front contacts were fabricated using conventional UY lithography 

and electroplating. The positive photoresist (- 5 j.lm thick) was deposited on clean and chemically oxidized 

emitter surfaces. In both cases about 0.5 j.lm Au-Zn or Au-Cd were first electrodeposited by pulse plating at 

pulse current density of 0.5 to 2 rnA/cm2, then 5 to 18 j.lm Au was deposited at a constant current density of 

0.2 to 0.3 mA/cm2. When using clean surfaces the width of the contact grid fingers became up to 3 times the 

designed values, while their width have not increased significantly when a 20 to 50 run chemical oxide was 

used. Electroplated Au-Zn or Au-Cd front contacts we found are well suited for deposition on thin emitters 

since they do not require sintering. For example, using - 8 j.lm thick electroplated Au-Zn contacts, grown on 

a 0.6 j.lm thick emitter, using an oxidized surface, we recorded Rs values as low as 1.28 n-cm2, and FF 

values of over 80% , after thinning the emitter to about 0.3 to 0.4 j.lm. Since for pin configuration the sheet 

resistance is a major contributor to Rs' we estimate that by using an optimized front grid mask, Rs values of 

less than 0.5 n-cm2 and FF greater than 84% can be achieved after thinning the p+emitter such as in Fig.6 

to 0.25-0.3 j.lm. 

Surface Passivation 

One of the key factors limiting the performance of InP solar cells is their high surface recombination 

velocity (SRY), which is estimated, even for epitaxially grown cells to be as high as 107 cm/s (ref.21). 

Although not near to such an extent as the n + p InP structures, p + n InP structures fabricated by thermal 

diffusion have their surface stoichiometry destroyed. Therefore, it is important in the fabrication of highperformance InP solar cells in general and diffused InP cells in particular, to remove in a controlled manner 

the high defect density surface layer of the emitter and to passivate the surface. Calculations have shown that 

SRYs higher than 5 x 105 cm/s drastically reduce the efficiency of InP solar cells by lowering their blue 

response (ref.22). Simple chemical treatments of InP surfaces using HN03 and HF based etchants (ref.23) 

were found to decrease the SRY to below 5 x 105 cm/s, e.g. 1.7 x 105 for n+ -lnP and 4.7 x 105 cm/s for 

p+ -lnP, after rinsing the substrates in a HN03 (15%) solution (ref.24). 

Using the PNP etch, based on HN03, o-H3P04, and H202, we developed for thinning after contacting 

the p + -lnP emitter (ref.20), from low frequency EG-V measurements, we recorded a surface state density 

minimum (Nss) at the Cd-diffused p+-InP/passivating layer interface as low as 2 x10 10 cm-2 ey-l. About 

40 run was removed from the surface of a the p + n InP structure diffused at 660 0C (surface acceptor 

concentration: - 4 x 1018 cm-3). Such a low Nss value is in good qualitative agreement with the high 

measured Yoc and blue response values of solar cells fabricated on these structures. 

AR Coating 

The residual oxide grown on p + -lnP using the PNP etch is composed of a thick In-rich outer layer and a 

P-rich layer at the interface with the emitter (Fig.8). From XPS investigation (ref.20), the interfacial oxide is 

rich in In(p03)3. Since this oxide, as seen above, passivates the surface quite well, and it has a bandgap of 

6.8 ±0.2 eY (ref. 25), we proposed that it be used as a first layer AR coating (ref. 16). The transparency of 

67 
this oxide over the measured 1.8 to 5.2 eV range (ref.25) and its low blue reflectivity, as compared to SiO, 

Sb203, shown in Fig.9, and optimized ZnS/MgF2 double layer AR coating (not shown here), seem to make 

this oxide a very attractive candidate, indeed , for use as a first layer AR coating. In addition, when we 

deposited on our bare p + n InP cells either SiO or Sb203 or a double layer of ZnS/MgF2, the Voc dropped 

by as much as 50 mY, indicating a large increase in SRV. As shown in Fig. 10, the two layered oxide (- 130 

nm) reduces the reflectance of an p+n InP solar cel1 from an average of 40% to slightly less than 20%. In 

this particular case, after removing the In-rich outer-oxide layer, the reflectance of the remaining thin 

In(p03)3 oxide is about 25 %. 

Although the overall reflectance of the double-layered chemical oxide in Fig.9 is lower than that of SiO, it 

is still too high for use as a single layer AR coating. Furthermore, the outer In-rich oxide is unstable and 

quite conductive, which caused for our cel1s a noticeable drop in Rsh' and Voc. Therefore, we removed it, 

and in our best design we add Al203 and MgF2 as second and third layers of the three-layer coating. For 

the example in Fig.10, a three-layer AR coating composed of In(p03)3 (45 nm) I Al203 (62 nm) I MgF2 (41 

nm), reduces the overall reflectivity (no grid fingers) to less than 2 %. Details of this design wil1 be given 

elsewhere (ref.26). 

Progress in p + n InP Diffused Solar Cells 

As mentioned above, our efforts over the last year or so were concentrated on designing: 1) thin p + -lnP 

emitters, 2) front contacts, 3) passivating layer and 4) AR coating, so as to minimize the large external losses 

present in our cells. Therefore, since these efforts were made simultaneously, solar cells were only 

fabricated to check our progress and to correlate other measured parameters to solar cell parameters. Table 1 

shows some preliminary results. As an example, by using a more step-like diffusion profile of Fig.5 for cel1 

#3, instead of the more graded profile of Fig.4 for cell #1, the Rs value could be noticeably decreased, using 

the same n-type front grid design. An increased FF made it possible to increase the AMO efficiency from 

11.25% to 13.2 %. The relatively lower Voc for cell #3 is due to current leakage which occurs through the 

In-rich oxide (- 110 nm) outside the mesa etched active area. This is confirmed by the data of cel1 # 4. As 

seen, upon removing the In-rich outer layer, although the Jsc value decreased, as expected from the 

reflectivity measurements such as shown in Fig.10, the Voc value increased. The small increase in FF was 

due to an increase in Rsh. By depositing a non-optimized SiO (- 85 nm) second AR coating layer, the cell 

current density increased by about 12 %. To make sure that no current is collected from outside the active 

area of the cell, this particular cel1 was cleaved around the mesa etched defined lines, and no noticeable 

changes in cel1 AMO parameters were observed. 

From correlations between measured cell parameters, reflectivity, spectral response, dark saturation 

current densities and J sc-V oc plots, the projected AMO~ 25 0C performance parameters of p + n (Cd,S) InP 

solar cel1s, using our state-of-the-art newly developed thin emitters, and optimal front grid (6% coverage), 

and the newly designed three layer AR coating, are: Voc=910 mY, Jsc =37.85 mA/cm2, FF=84%, and 11 

=21.2 %. These performances are predicted for an emitter thickness of 0.3 ~m, a surface acceptor 

concentration of 3.5x1018 cm-3, base electron concentration of 7.5xlOI6 cm-3, front SRV of 105 cm/s, and 

using LEC grown InP:S substrates with EPD=5x104 cm-2 . Higher efficiency is possible by using better 

quality substrates, further improving the diffused structures and the cel1 fabrication sequences. 

Some light degradation effects have been observed in our earlier p + n (Cd,S) cel1s, which we have 

attributed to excess Zn content (> 10%) in the Au-Zn front contacts. Indeed, as seen in Fig.II by lowering 

the Zn content to less than 10%, the relative degradation of cel1 parameters decreases from about 10% to less 

than 1 %. We believe that by using lower Zn content (0.1 to 1 %) Au-Zn front contacts, which recent studies 

indicate to offer a lower contact resistance, this problem can be eliminated. We observed similar degradation 

effects in our n+p (S,Zn) diffused cells with high base acceptor concentation (> 1017 cm-3). However, as 

seen in Fig.11 no light degradation is observed for a cell with lightly base doping (2.4 x 1016 cm-3). 

Preliminary Radiation Resistance and Annealing Studies. 

68 
Preliminary results of radiation resistance studies of diffused p + n (Cd,S) InP solar cells, such as shown in 

Table 2, indicate that the percent of remaining power (39%), after irradiation with 1013 cm-2, 3 MeV 

protons is higher than that of other InP cell structures, including the all-MOCVD fabricated n+p (Si,Zn) and 

diffused n +p (S,Zn) InP cells, for which irradiation data using 3 MeV protons are available (ref. 15). 

Furthermore, this cell shows a remarkable annealing property at room temperature (RT) in the dark. The 

AMO, 25 oC, performance parameters of this cell prior to irradiation and after about I year at RT in the 

dark, are shown in Table 3. Subsequent light soaking of this cell for I hour under AMI.5, 25 oC, raised its 

efficiency by about 2.5% (See Fig.ll), indicating good annealing properties under illumination. 

Preliminary radiation resistance studies and annealing studies of these cells have been started at Spire 

Corp., after irradiation with high energy alpha particles. For one such cell, for which data are available, after 

irradiation at an equivalent IMeV e- fluence of 1.06x1017 e-/cm2, which corresponds to over 100 years in 

GEO, the remaining power output is 32 % of the initial power (fable 4). As a result of a significant recovery 

of about 6%, after only 4 days at RT in the dark, and an expected higher recovery rate under the cell 

operating conditions (RT, under illumination), these cells are not expected to degrade significantly in high 

radiation environment orbits, even after such large fluences. 

We attribute the high radiation resistance of diffused p + n (Cd,S) InP cells to a a very low carrier removal 

rate (Fig.12) in the emitter (ref. 15). It is note worthy that the carrier removal in the InP:S base, after 

irradiation with 1013 cm-2, 3 MeV protons has decreased by more than an order of magnitude. For a thick 

p+ emitter, most of the cell current is not expected to come from the base or space charge region. Since, as 

seen, the more heavily doped Cd-diffused emitter degrades less than the low doped base, the superior 

radiation resistance of these cells, as compared to a thin emitter n +p configuration with a thin emitter, 

should be expected . 

Low Cost Processing Scheme for High Efficiency Radiation Resistant p + n InP Diffused 

Solar Cells. 

For InP solar cells to be commercially useful for practical space mission applications, their cost must be 

significantly reduced, and they should achieve high BOL and EOL efficiencies. For small to medium power 

requirements, the weight might not be a prime requirement, as long as the cells are intended for high 

radiation environments. Diffused structures InP cells in general, and p + n (Cd,S) cell structures in 

particular, appear to be more radiation resistant than cells fabricated by epitaxy. A simplified processing 

scheme is proposed in Fig. 13, for fabrication of high efficiency, radiation resistant p + n InP diffused junction 

cell. The scheme we propose ensures not only a low fabrication cost, but also high throughput and 

reproducib il ity. 

As shown in the previous paragraphs, preliminary results show that cells using Cd-diffused emitter have 

not only a good potential for achieving high BOL efficiencies, but they also appear to be more radiation 

resistant and to have better post irradiation annealing properties than other diffused cell structures. Since for 

the p + n configuration, the Cd-diffused cells are more radiation resistant than the Zn-diffused emitter cells 

fabricated using similar S-doped substrates, it will be very interesting to see how the radiation resistance of 

epitaxial and diffused n +p cells, with similar structures, e.g. (S,Zn) do compare. Also it would be 

interesting to compare p + n InP homojunction or heteroepitaxial cell structures with the emitter diffused into 

a thin base grown epitaxially on heavily doped InP or cheaper substrates such as GaAs, Ge or Si with allepitaxial grown similar cell structures. 

CONCLUSIONS 

• We have found the ranking in decreasing order of projected maximum efficiency of diffused junction InP 

solar cells to be: 1) p+n (Cd,S), 2) n+p (S,Cd), 3) n+p (S,Zn), and 4) p+n (Zn,S). Preliminary 

investigation indicates the same ranking holds for these structures with respect to radiation resistance. 

69 
• If further developed, the p + n (Cd,S) InP solar cells, developed by CSU/NASA LeRC, offer the 

following significant advantages for space power applications: 

• Low processing cost 

• High throughput 

• Good reproducibility 

• High projected efficiency 

• High radiation resistance 

• Self annealing during operation 

• The radiation resistance of these cells, combined with their annealing behavior under operating 

conditions, may eliminate the need for shunt circuits used in conventional satellites to dump excess power 

early in their missions. 

• Although InP has about twice the density of Si, or Ge substrates, the ability of diffused InP to anneal 

under operating conditions allows the thickness of the protective cover glass to be reduced, compensating 

for the difference in the substrate weight. 

• The technology of InP diffused structures, electroplated contacts, In(p03>31 Al203/MgF2 ARC, and 

improved electrochemical characterization techniques, developed in this work, can be applied to 

fabrication of other III-V solar cell and opto-electronic devices. 

References 

1. M. Yamaguchi and K.Ando, "Mechanisms for Radiation Resistance of InP Solar Cells," 

J.App!. Phys. 63, 5555 (1988). 

2. LWeinberg, C.K. Swartz, R.E.Hart, and R.L.Statler, "Radiation and Temperature 

Effects in Galium Arsenide, Indium Phosphide, and Silicon Solar Cells," Proc. 19th IEEE 

PVSC, 1986, p. 548. 

3. I. Weinberg , C.K.Swartz and R.E.Hart, "Potential for Use of lnP Solar Cells in the 

Space Radiation Environment," Proc. 18th IEEE PVSC , 1985, p.2l. 

4. M.Yamaguchi, C.Uemura, A.Yamamoto and A.Shibukawa, "Electron Irradiated 

Damage in Radiation-Resistant InP Solar Cells," Jpn. J. App!. Phys. 23, 30 (1984). 

5. C.J.Keavney, V.E.Haven and S.M.Vernon , "Emitter Structures in MOCVD InP Solar Cells," 

Proc. 21st IEEE PVSC, 1990, p.141. 

6. C.J.Keavney, S.M. Vernon, V.E.Haven, S.J.Wojtczuk and M.M.AI-Jassim, "Fabrication of n + Ip 

InP Solar Cells on Silicon Substrates, " App\. Phys. Lett., 54, 1139 (1989). 

7. S.J.Wojtczuk and S. Vernon, "pin InP Solar Cells on Ge Wafers," This Conference. 

8. H.Okazaki, T.Takamoto, H.Takamura, T.Kamei, M.Ura, A.Yamamoto and M.Yamaguchi, 

"Production of InP Solar Cells for Space Power Generation," Proc. 20th IEEE PVSC, 1986, 

p.886. 

9. M.Faur, M .Faur, C.Goradia, M.Goradia, N.Fatemi, D.J.Brinker and R.Thomas, "InP Solar Cells 

made by Closed-Ampoule Diffusion of Sulfur into Cadmium-Doped Substrates: Dependence of 

Cell Characteristics on Diffusion Temperature and Time," Proc.lst Int'l IPRM Conference, 1989, 

p.459. 

10. S.Bothra, H.G.Bhimnathwala, K.K.Parat, S.K.Ghandhi and J.M.Borrego, "Characterization and 

Modeling of Open Tube Diffused n+p InP Solar CelIs," Proc. 19th IEEE PVSC, 1987, p.261. 

11. R.J. Walters, G.P.Summers and J.Bruening, "A Detailed Study of the Photo-Injection Annealing 

of ThermalIy Diffused InP Solar CelIs," Proc. 12th SPRAT Conference, 1992, p.l. 

12. S.R.Messenger, R.J.WaJters and G.P.Summers, "High Temperature Annealing of Minority 

Carrier Traps in Irradiated MOCVD n +p lnP Solar Cell Junctions," ibid., p.8. 

13. M.Faur, M.Faur, D.J.Flood, LWeinberg, D.J.Brinker, C.Goradia, M.Goradia and W.Thesling, 

"A Comparative study of p + nand n + p InP Solar CelIs Made by Closed Ampoule Diffusion, " 

Proc. 11th SPRAT Conference, 1991, p. 3-1. 

70 
L 

14. 

15. 

16. 

17. 

18. 

19. 

20. 

21. 

22. 

23. 

24. 

25. 

26. 

S-K.Min, H.Y.Cho, W.C.Choi, M.Yamaguchi and T.Takamoto, "Hydrogenation Effects on n+-p 

InP Solar Cells," Proc. 23rd IEEE PVSC, 1993, p.639. 

M.Faur, M.Faur, D.J.Flood, I. Weinberg, C.Goradia, M.Goradia and c.Vargas-Aburto, "Low 

Carrier Removal Rates and Annealing Behavior of Thermally diffused p + n (Cd,S) InP Structures 

After 1013 cm-2 3 MeV Proton Irradiation," ibid., p.1437. 

M.Faur, M.Faur, D.J.Flood, D.J.Brinker, C.Goradia, S.Bailey, I.Weinberg, M.Goradia, 

D.T.Jayne, J.Moulot and N.Fatemi, "Effective First Layer Antireflecting Coating on InP Solar 

Cells Grown by Chemical Oxidation," Proc. 6th Int'l IPRM Conf. , 1994. 

M.Faur, M.Faur, C.Goradia, M.Goradia, D.J.Flood, D.J.Brinker, I.Weinberg, C.Vargas and 

N.S.Fatemi, "Progress in p+n InP Solar Cells Fabricated by Thermal Diffusion," Proc. 12th 

SPRAT Conf., 1992, p.23. 

M.Faur, M.Faur, C.Goradia, M.Goradia and I.Weinberg, "High Quality Thermally Diffused 

p+n loP Structures," 3rd Int'l IPRM Conf., 1991, p.304. 

M.Faur, M.Faur, M.Goradia, C. Vargas and D.M. Wilt, "Electrochemical Characterization of 

n+p and p+n InP Diffused InP Structures," Proc. 12th SPRAT Conf. , 1992, p.33. 

M.Faur, M.Faur, D. T.Jayne, S.Bailey and M.Goradia, " Etchant for Chemical Thinning of InP 

and its Applications in the Fabrication of InP Diffused Junction Solar Cells," Surface and 

Interface Analysis, 21, 110 (1994) 

C.Keavney, " Improvement in InP Solar Cells," Proc. 20th IEEE PVSC, 1988, p.654. 

C.Goradia, W.Thesling and I.Weinberg, "Key Factors Limiting the Open-Circuit Voltage ofn+pp+ 

InP Solar Cells," Proc.22nd IEEE PVSC, 1991 , p.386. 

M.Faur, M.Faur, P.Jenkins, M.Goradia, S.Bailey, D.Jayne, I.Weinberg and C.Goradia, "Study 

of Surface Passivation of InP," Surface and Interface Analysis, 15, 745 (1990) 

P.Jenkins, M.Goradia, M.Faur, S.Bailey and M.Faur, "Measurement of Surface Recombination 

Velocity on Heavily Doped Indium Phosphide," Proc. 21st IEEE PVSC, 1990, p.399. 

J.Joseph, Y.Robach, G.Hollinger, P.Ferret and M.Pitval, "The Passivation of InP by In(P03h 

for MISFET Applications," Proc. 1st Int'I IPRM Conf. , 1989, p.268. 

J.Moulot, M.Faur, M.Faur, C.Goradia, M.Goradia, S.Bailey and N.Fatemi, "Design of a ThreeLayer Antireflective Coating for High Efficiency Indium Phosphide Solar Cells using a Chemical 

Oxide as First Layer, " Review Abstract submitted to the 1st World Conf. on Photovoltaic Energy 

Conversion, Hawaii, Dec. 5-9, 1994. 

20 

16 

::? 

.s 12 

IZ 

UJ 

(( 

(( 

:::> 

<.) 8 

.2 .4 

rvf\S/\ 

.6 .8 

Lewis Research Center 

I" = 12.832 mA 

v., = 887.6 mV 

1m.. = 11 .68 mA 

V m .. = 673.3 mV 

Pm .. =7.868mW 

F.F. = 69 

Efl. = 11.98 % 

Cell : 81 

Dale : 13 Oct 1993 

Reference Cell: A·181 , 

Area : .48 cm 

Temperature : 2S·C 

Air Mass Zero 

File Name : 81A.I_V 

1.2 

VOLT AGE (volts) 

1.4 

Fig. I. AMO, 250C illuminated I-V characteristic. Emitter thickness: - 0.45 Ilm; 

AR coating: - 30 nm In(p03)} . 

71 
Table 1: AMO, 250C perfonnances of selected diffused junction p+n (Cd,S) InP solar cells measured at NASA LeRC 

Diffusion AR coating Approx. Rs Jsc Voc FF T\ Cell # Profile emitter (Q-cm2) (mA/cm2) (mY) (%) (%) 

such as thickness 

in: (Jlm) 

1 fig. 4 In(P03})b (- 30oA) 0.35 4.92 27.6 886.9 62.8 1l.25 

2 fig. 5 In(P03})b (- 400A) 0.45 3.24 27.5 884.6 73.7 12.95 

3 fig. 5 In203a (900A)/ 0.4 3.35 28 .2 881.7 72.6 13.2 

In(P03})b (- 30oA) 

In203a (1100A)/ 29.4 877.2 61.7 11.63 

In(P03)3b (- 400A) 

4 fig. 4 In(P03})b (- 400A) 0.3 4.38 27.6 886.6 62.8 11.25 

SiO (- 850A)/ 30.95 887.3 61.5 12.36 

In(P03})b (- 400A) 

The resIdual cherrucal oXIde after dlssolutmg the p+ InP emItter usmg the PNP etch [4], has two components: (a) a thick In203-

rich surface layer, and (b) an interfacial In(p03h-rich layer. 

Table 2: AM1.5, 250C performance parameters of a diffused p+n (Cd,S) InP solar cell before (b) and after irradiation 

with 1013 cm-2, 3MeV protons. 

Time after Voc Jsc FF T\ Rs 

irradiation (mY) (mA/cm2) (%) (%) (n-cm2) 

(hours) 

o (b) 871 27.42 72.64 12.65 3.27 

8 664 15.48 64.00 4.86 5.33 

90 676. 16.37 65.10 5.25 5.19 

206 682 16.85 65.50 5.49 4.88 

857 708 18.47 66.00 6.29 4.85 

(*) The cell was kept at RT In the dark except for 6 short exposures to lIght dUrIng IllumInated I-V measurements 

72 
L _ _ 

---------

Table 3: AMO, 250 C performance parameters of a p+n (Cd,S) InP 

solar cell (same as in Table 2), before (b) irradiation and 

1 year after irradiation (a) with 1013 cm-2 3MeV protons. 

Voc Jsc FF TJ 

(mY) (mA/cm2) (%) (%) 

(b) 880.3 26.81 73.1 12.57 

(a) 713.9 18.02 67.0 6.31 

(") Measurements performed at NASA LeRC. After irradiation, 

the cell was kept at RT in the dark except for 6 light exposures 

during the AM1.5, 250C illuminated I-V measurements in Table 2. 

Table 4: AMO, 250C performance parameters of a CSU p+n (Cd,S) InP solar cell prior to and after alpha irradiation 

at IMeV equivalent electron fluence of 1.06x1017 cm-2. (*) 

Voc Jsc FF T) JOI (nl=l) J02 (n2=2) Rs Rsh 

(mY) (mAJcm2) (%) (%) (AJcm2) (AJcm2) (!l-cm2) (x 105 !l) 

Before irradiation 888 27.18 69.21 12.17 2.8510-17 7.710-11 3.66 5.1 

Post irradiation 652 13.64 60.14 3.90 2.72 10-14 3.4 10-08 6.04 2.8 

After 4 days at RT 659 14.32 60.14 4.13 2.1 10-14 3.05 10-08 5.91 2.6 

in the dark 

(*) The measurements and irradiation were performed at SPIRE Corp., and presented here with permission from Dr. 

C. Blatchley. 

73 
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o 

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·2 $ : 

'0 . Cl - a - .-J -3 _ 

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Jo 1.38E- 17 A / cm2 ( A 

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. ./1 / 

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·~ .O 0.1 0.2 0.3 0.4 0.5 0.6 0.7 O.S 0.9 1.0 1.1 

0.9 

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0.7 

u 

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E 

OJ 

C 

'" 0.5 OJ 

o 

'iii 

~ 0.4 

x 

w 

0.3 

0.2 

0.1 

Voltape (V\ 

Fig. 2. Dark I-V characteristic of cell # 4 in Table 1. 

0.0 L-.!......l---"-'-' .1..' ....l'--'.. .....l..!...!...!...!...l...!..L!..!...l...!..J....LLJLLL!...!...LJ...!..l.J...L;'-..L!..~J...!..l-:-'-L!..!..:-::--'-..L!..J::-'-'-'--'-:::=--:-."-J .3 .35 A .45 .5 .55 .6 .65 .7 .75 .8 

Wavelength (iJm) 

Fig. 3. Typical EQY plot of a cell with a thin In(p03h (- 30 run) as an AR coating. 

Overall reflectivity: - 30%. Emitter thickness: - 0.45 Ilm. 

74 
(T) 

I 

C5550201 (CD/P/NINP(NINP :5/7 . 1E15) 20~ -- -' -- -- ' -, - - - -r -- - ,_.- I _. -- I 

I 

r 

; 

1St: 

[ p+-InP 

27 Oct 1993 - , -- ~ 

~ ""..-.--------------

I 

U 

-:f-

~ ~W 

z i 

18~ 

r 

I 

r 

I 7~ .- -- - -1 . 1 _ _ _ I __ 

2 3 

Depth (u m ) 

__ J 

4 

, 

I 

::l 

-; 

J 

5 

Fig.4. EC-V profile of a p+n (Cd,S) InP structure diffused at 660 0C. 

Diffusion cap: In(p03)3 (- 3 run ). 

NIGR13 (l NP/ P/N ) ISr- ---'- ' ------.-•. -- --,.- -00- --- '''- -.- -.-", -._-.----,- .- II Oct 1993 00 -,--.- - - --'-- ------l 

"1 ~ 1 r 

r 

r 

[ 

I 

7 I E l8i- t) ... 

~ r 

z r 

r 

~ 

~ 

l 

p+-InP 

_. '-. _ ___ . . _ L _______ __ 1 _ __ _____ . 1. _. _ 0 •• • • _ I 

2 3 

Depth (urn) 

.., 

-l 

~ 

j 

i 

I 

, 

i 

I 

I 

I I 

l u 

] 1--

-iw 

-! 

, 

"1 

"1 I 

I 

l 

I 

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4 5 

Fig.5. EC-V profile under similar conditions as in Fig. 4, but using - 10 run 

thick In(p03h diffusion cap. 

75 
p+-InP 

18 

(T) I I U E 

u ~ ~ l.JJ 

z n-InP 

17 

160 .1 .2 .3 . 4 .5 .6 .7 .8 .9 

J 1 

Depth ( um ) 

Fig.6. Thin emitter p + n (Cd,S) loP EC-V profile, at 660 0c. using 

In(p03h (- 3 om thick) diffusion cap. 

(T) 

I 

E 

(J 

Z 

NIGR60025 (INGRS / INPDIF/ UNIEL l 

201' I 

17 

p++ -lnGaAs 

(b) 

I 

I 

(a) 

p+-InP 

\ 

MF 11 Ma y 1994 

----:J 

j 

J 

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: I 

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::Jf--

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Dep th ( um) 

Fig.7. EC- V profiles of p + n (Cd,S) loP structures, diffused in similar conditions 

at 600 °C through: (a) undoped InGaAs, and (b) In(p03) 3 cap layers. 

76 
-.l 

-.l 

70001-

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5S01-

----

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4SOJ7 4000L 

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28 

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--

o LI 

300 400 500 600 700 800 900 

Wavelength (nm) 

Fig.9, Surface reflectance of p + n InP using a single layer 

AR coating. 

Fig.S, XPS profiles of an oxide grow

PNP etch, 

" u c 

S 

c:: 

" 

70 

60 

50 

40 

30 

\1\(P

20 

\ n20 

10 

In(p03}) (45 11m) I AI203 (6

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300 400 500 600 

Wavelength (

Fig,IO, Surface reflectance of p 

double or single layer chemical o

bare surface and optimized three lay
5 I 

P + n (Cd,S)' 1 013cm-2, 3MeV protons 

-

.... 

w n+E!(S,Zn) ~ : - > 10% Zn 

~ .. / '

, 

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- • • ~ -5 'V 'V ''::: 

~ 

> 

'V 

* • • Efficiency 

V Isc 

• 0 Voc 

-10 

0 5 10 15 20 

Cell No. 

Fig.l1. Relative degradation ofp+n (Cd,S) loP (cells # 1-17), and n+p (S,Zn) loP (cell #18) 

parameters after 1 hour AM 1.5, 25 °C illumination. For cells # 1-16, the Zn-content of 

evaporated Au-Zn front contacts is shown. 

19.L------~------~------~------._------._----_c 

= Sample 4. 

, 

IO~----

r 

, 

1. Prior to irradiatiol1 

2. 8 hours after 1013 CI11"2 3MeV protol1s 

-

('T) 

I 

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:I 

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i 

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, 

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Fig.12. EC-V profiles of a p+n (Cd,S) loP structure, prior to and after 1013 cm-2 3 MeV 

protons irradiation. 

78 
-...l 

\0 

t. Diffusiol1 up 

2. Diffusion 

3. Dead l:tycr rClUov:a1 

.• :. -.ziif 

., 

4. Photoresi't 

.. 

5. nack SUrf;lCC oxide :lIId diffused l:lyel' 

.• .i.C:.:,: 

Fig. 13 . Processing scheme of high efficiency p + n (Cd,S) InP diffused-junction solar cell. 

6. Back contart II (a). Lift off 

7. rellloval 

12(a). CootaclS sintoring 

8(.). Photolithography 13(0). Electroplating 

9 (a). Etching 8(b). flack contact sintcring 

10(a). Evaporated thin contacts 

9(b). Photolithography 

10(b). Etching 

11(b). Electropillting 

tl(b). Resist removal 

14(0) or 13(b). Pbotolithography 

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CJ 

p+ InP:Cd diffused lay

Metal 

n-InP:S substrate 

A12031MgF2 
J9?1¥c, 

N95-20510 

HIGH QUALITY InP-on-Si FOR SOLAR CELL APPLICATIONS 

Zane A. Shellenbarger, Thomas A. Goodwin, 

Sandra R. Collins, and Louis C. DiNetta 

AstroPower, Inc. 

Newark, Delaware 

SUMMARY 

InP on Si solar cells combine the low-cost and high-strength of Si with the high efficiency and 

radiation tolerance of InP. The main obstacle in the growth of single crystal InP-on-Si is the high residual 

strain and high dislocation density of the heteroepitaxial InP films. The dislocations result from the large 

differences in lattice constant and thermal expansion mismatch of InP and Si. Adjusting the size and 

geometry of the growth area is one possible method of addressing this problem. In this work, we 

conducted a material quality study of liquid phase epitaxy overgrowth layers on selective area InP grown 

by a proprietary vapor phase epitaxy technique on Si. The relationship between growth area and 

dislocation density was quantified using etch pit density measurements. Material quality of the InP on Si 

improved with both reduced growth area and increased aspect ratio (length/width) of the selective area. 

Areas with etch pit density as low as 1.6 x 104 cm-2 were obtained. Assuming dislocation density is an 

order of magnitude greater than etch pit density, solar cells made with this material could achieve the 

maximum theoretical efficiency of 23% at AMO. Etch pit density dependence on the orientation of the 

selective areas on the substrate was also studied. 

INTRODUCTION 

The material quality of InP films on silicon has not yet achieved acceptable levels. The critical 

issue is the reduction of dislocation density and residual strain. Yamaguchi (ref. 1) has estimated the 

material quality characteristics required for obtaining high performance devices using III-V films on silicon. 

A dislocation density of 106 cm-2 and a residual strain of 109 dyne/cm2 are the estimated requirements for 

a solar cell. The lowest reported dislocation density of 3 x 107 cm-2 (ref. 2) for InP-on-silicon has not yet 

reached this level. Hence, state-of-the-art accomplishments are not within the high performance solar 

cell realm. 

Both dislocation density and residual strain can be decreased by using a reduced growth area. It 

may be possible to improve the quality of InP on Si solar cells by exploiting this property. Instead of planar 

growth, a mosaic structure of closely spaced selective area growths can be grown. These mosaic units 

can then be monolithically interconnected to form a large area InP on Si solar cell structure. 

Reductions in dislocation density resulting from reduced growth area have been reported by many 

groups for many different heterostructures. Noble (ref. 3) best described the reduction in dislocation 

density by the illustrations shown in Figure 1. Limiting the film area limits the effective dislocation length, 

which in turn reduces the probability of dislocation interaction and multiplication, and reduces the 

dislocation density of the material. Figure 1 illustrates the reduction in dislocation density as a function of 

effective dislocation length alone (dislocation interaction and multiplication are not illustrated). 

MI~ PAGe: SLAM( NOr f lLMW 81 
Reductions in residual strain through limiting growth area have been reported by several groups 

for GaAs-on-silicon (refs. 4,5,6). Strain-relief in selectively grown films can be expressed by the bi-metal 

model as: 

e = eo (1 - exp [-k(w-x)]) 

where e and eo are residual strain for selective area growth and planar growth, respectively, w is half the 

patterned width, x is the distance from the center of the patterned film and k is the interfacial compliance 

parameter (ref. 4). This equation shows that a reduction in film area (i.e., a reduction in w) causes a 

reduction in residual strain (e < eo)' For selective GaAs films on silicon, residual strain has been reduced 

to zero by limiting film area to a 10 11m by 10 11m square (ref. 5). 

The goal of this research was to quantify the benefits of reduced area InP films on silicon using a 

combined vapor phase epitaxy (VPE) and liquid phase epitaxy (LPE) growth process. Etch pit density 

measurements quantified the relationship between the growth area dimensions and dislocation density as 

well as the relationship between dislocation density and the orientation of the selective areas. On the 

basis of the demonstrated reduced area and combined technology benefits of GaAs films on silicon, we 

expected a significant reduction in etch pit density that would bring the quality of InP films on silicon within 

the device realm. 

APPROACH 

Single crystal n-type InP was grown on selectively masked on-axis (111) n-type Si substrates by a 

proprietary VPE technique. The average thickness of the VPE buffer layer was .5 11m. The growth areas 

were defined by chemically etching selective areas in a thermally grown Si02 masking layer. The growth 

geometry was evaluated using two different selective area patterns. Figure 2 displays the mask used to 

evaluate growth area size and aspect ratio. This pattern had selective areas ranging in size from 60 11m 

by 60 11m up to 4000 11m by 4000 11m. Figure 3 displays the mask used to test the orientation. This 

pattern had 400 11m wide selective areas oriented every 30°. On a (111) wafer the < 110> and < 112> 

equivalent directions are aligned in alternating 30° intervals. The masks were aligned with one edge 

parallel to the <110> edge. Therefore, on the mask in Figure 2 the selective areas were aligned with one 

dimension in the <110> direction and the perpendicular dimension in the <112> direction. On the mask in 

Figure 3 the selective areas were aligned in alternating <110> and <112> directions. The initial vapor 

phase growth produced stray InP crystals on the Si02 surface that created melt carryover problems during 

the LPE overgrowth. To prevent this, the selectively grown areas were masked with photoresist and any 

excess InP crystals were chemically etched from the substrate surface prior to LPE. In preparation for the 

LPE overgrowth, the substrates were cleaned in organic solvents and etched in H2S04:H202:H20 

(2: 16: 1 000) for 30 sec. 

The LPE growth system used for these experiments consisted of a quartz reactor tube, a 3-zone 

moveable furnace, a mechanical vacuum pump, and gas sources of nitrogen and palladium-diffused 

hydrogen. A graphite multi-well horizontal slider boat transported the substrate to the various growth 

melts. In order to protect the VPE InP layer from thermal degradation, a Sn-In-P overpressure melt as 

well as an InP polycrystalline cover wafer were used over the substrate prior to LPE growth. To enhance 

wetting of the growth area, the first growth melt contained In:Sn (3: 1) solvent. Successive melts contained 

pure In solvent. The method of growth used was a two-phase ramp cool. Best results were obtained with 

a growth temperature of 694°C, a supercool of 6°C, and a cooling rate of O.loC/min for the first layer and 

0.250C/min for each subsequent layer. It is necessary to grow multiple layers to distribute the strain 

associated with the lattice mismatch of InP and Si. Growths with 3 and 5 layers exhibited good crystal 

quality, free of cracks. The average total thickness of the LPE growth layers was 5 11m. Figure 4 is an 

illustration of the growth structure. Figure 5 shows the surface of a typical growth. 

82 
The dislocation density of the growth surface was quantified with etch pit density (EPD) 

measurements. To reveal the etch pits, the growth surface was etched in HBr:H20 2:HCI:H20 (10: 1: 1 0:40) 

for 10 sec. The etch pits are visible as triangular wells or depressions as in Figure 6. The EPD was then 

determined by counting the etch pits in a known area on a photomicrograph. By analyzing several areas 

across the growth region, an average EPD was accurately determined. 

RESULTS AND DISCUSSION 

The first area of focus of this research was to determine the relationship between the selective 

area dimensions and film quality of the LPE overgrowth layers using the pattern shown in Figure 2. Etch 

pit densities of the selective area growths were determined using the method previously outlined. Most 

growths had EPD between 4 x 105 and 2 x 106 cm-2 for the different selective areas. The lowest etch pit 

density of 1.6 x 104 cm-2 was obtained for a selective area with dimensions of 80 J..lm by 2000 J.Lm . Figure 

7 shows a graph of EPD versus selective area for a typical growth. As expected, the EPD decreased with 

decreasing area. The sets of two data points connected by vertical lines represent selective areas with 

the same dimensions but aligned perpendicular to each other. There was no substantial difference 

between selective areas oriented in the two different directions. The numbers next to the data points are 

the aspect ratios (length/width) of the selective areas. As seen in the graph, the data fall into two distinct 

groups. The group having lower etch pit densities consists of those selective areas with large aspect 

ratios (length/width). All of these selective areas have a short dimension of 500 J..lm or less. Figure 8 is a 

plot of the etch pit densities versus the short dimension of the selective area for three different long 

dimensions. Above a width of 500 J.Lm there is little dependence on the short dimension. Below 500 J.Lm 

there is a strong linear dependence on the short dimension. The etch pit density reduces at a rate of 

about 2000 cm-2 for every 1 J.Lm reduction in the short dimension. Reducing the short dimension from 500 

J.Lm to 100 J.Lm causes an order of magnitude reduction in etch pit density. Holding the short dimension 

constant and varying the long dimension shows almost no change. These results indicate that a long thin 

selective area will have substantially better material quality than a square selective area with the same 

total area. We believe the explanation for this lies in the fact the material quality improves near the edge 

of the selective area due to lateral overgrowth onto the masking layer. As the selective area becomes 

thin, this effect starts to become significant. 

The second area of interest was dependence of material quality on substrate orientation. Si 

substrates were masked with the pattern shown in Figure 3. InP was then grown in these selective areas 

using our vapor phase InP overgrown with LPE InP. After growth, etch pit density measurements were 

performed on two of these samples. One had an average EPD of 4.05 x 105 cm-2 in the <110> directions 

and 7.64 x 105 cm-2 in the <112> directions. The other sample had an average EPD of 1.39 x 106 cm-2 in 

the <110> directions and 9.74 x 105 cm-2 in the <112> directions. This result, along with the lack of 

orientation dependence of the quality of growth in the first experiment, indicates that the material quality is 

independent of the orientation. 

While the etch pit density was not dependent on the orientation, the growth morphology was. The 

selective area edges oriented in the <110> directions had jagged overgrowth while the selective area 

edges oriented in the <112> directions had very smooth growth. The photograph in Figure 9 shows this 

effect. We believe this is due to the fact that the InP growth nucleates in a triangular pattern that has its 

flat side in the <112> direction. Figure 10 shows a photograph of an InP growth with only partial 

nucleation. The triangular nucleation areas have a flat side parallel to the edge of the selective area 

oriented in the <112> direction. For most applications, having the long edge of the selective area in the 

<112> direction would be more desirable. 

83 
CONCLUSIONS 

Using selective area growth is one method for improving the quality of heteroepitaxial InP on Si. 

Using a vapor phase buffer layer overgrown with LPE, etch pit densities as low as 1.6 x 104 were 

obtained. The following conclusions about the effects of selective area growth were reached. 

1. The etch pit density of the InP/Si heteroepitaxial growth can be significantly reduced by reducing one 

dimension of the selective area below 500 !Am. 

2. The etch pit density is not strongly dependent on the orientation of the selective area on the Si 

substrate. 

3. Jagged overgrowth occurs on the selective area edges that are aligned in the <110> directions. 

These results show it is possible to grown InP on Si by LPE with sufficient quality to support high 

performance solar cells. In order to make use of the benefits of selective area growth for high 

performance solar cells, methods to interconnect the individual areas must be developed. 

REFERENCES 

1. Yamaguchi, M., et al. : Dislocation Density Reduction in Heteroepitaxiallll-V Compound Films on Si 

Substrates for Optical Devices. Journal of Material Research, vol. 6 (2), 1991 , pp.376-384. 

2. Wanlass, M.W.; Coutts T.J .; Ward, J. S.; Emery, K.A. ; and Horner, G.S.: High-Efficiency, Thin Film InP 

Concentrator Solar Cells. Third International Conference on Indium Phosphide and Related Materials, 

April 1991 , pp. 40-47. 

3. Noble, D.B.; Hoyt, J.L. ; King, C.A. ; Gibbons, J.F.; Kamins, T. I.; and Scott, M.P.: Reduction in Misfit 

Dislocation Density by the Selective Growth of Si1 _xGex/Si in Small Areas. Applied Physics Letters, 

vol. 56 (1), 1990, pp. 51-53. 

4. Yamaguchi, M.; Tachikawa, M.; Sugo, M.; Kondo, S.; and Itoh, Y.: Analysis for Dislocation Density 

Reduction. in Selective Area Grown GaAs Films on Si Substrates. Applied Physics Letters, vol. 56 (1 ), 

1990, pp." 27-29. 

5. Yacobi, B.G.; Zemon, S.; Jagannath, C.; and Sheldon, P.: Strain Relief in Patterned GaAs Grown on 

Mismatched Substrates. Journal of Crystal Growth, vol. 95, 1989, pp. 240-244. 

6. Sakai, S.; Matyi, R.J .; and Shichijo, H.: Selective liquid phase epitaxy and defect reduction in GaAs 

grown on GaAs-coated silicon by molecular beam epitaxy. Applied Physics Letters, vol. 51 (23), 1987, 

p. 1913. 

84 

I 

I 

I 

I 

I 

I 

J 
.111 II 

it 

a) Planar Film b) Selective Film 

Figure 1: The effect of growth area on dislocation density (ref. 4). Heavy lines represent oxide stripes, 

narrow lines represent dislocation segments, and stars represent dislocation nucleation 

sources. 

Figure 2: Mask design for film area vs. film quality study (scale = 10X). 

85 
Figure 3: Mask design to determine optimum selective area orientation (scale = 10X). 

InP LPE Overgrowth 

InP VPE Layer 

~"" ======-" __ .L=====:J"_./ Si0

2 

Mask 

f Si Substrate T 

Figure 4: Growth structure. 

86 
Figure 5: Typical LPE overgrowth surface (scale = 200X). 

Figure 6: Typical LPE overgrowth surface showing etch pits (scale = 1000X). 

87 
2.0 X 106 

I 

4 

1.5 X 106 

8 

"i 1.67 • 

· 1333J 

E I 6.67 .s. 

z. 1.0 X 106 I ·iii 

c 5 Q) 50 

Q • 40 

- 20 Aspect Ratios I a. I 

6.25 .r. 5.0 x 105 U 

W 25 

o 

1.0 x 10~ 1.0 X 10.

3 

Figure 7: Graph of etch pit density vs. film area for growth J11605. 

1.60 X 106 

1.40 x 106 

• • 

"i 1.20x106 • 

E 0 0 

~ 

:?: 1.00x106 

·iii 

C 

Q) 

0 

B.00x105 

a:: • Long Dimension 4000 IlIT1 

.r. • 0 Long Dimension 2000 IlIT1 

6.00 X 105 0 W ~ 

• • Long Dimension 500 11m 

4.00 x 105 

2.00 x 105 ~ 

• 

c1 

~ 

0 

0 500 1000 1500 2000 2500 3000 3500 4000 4500 

Short Dimension (11m) 

Figure 8: Graph of etch pit density vs. short dimension for growth J11707. 

88 
. . , . 

• ' J.: .... • 

" ... .. ·Y·:.. , " . • ' " . .. ~..; " ". ~ ..... : :. e.. . : . ~ I . 

.... 

-, -f,. ... " .. --, 

•. \:~:. ~ 

r .'I.. .. '1' 

Figure 9: Edges of growth areas in <11 0> and <112> directions (scale = 100X). 

' . t (112) . 

. . ~~ .. . . '

. <: 

'. "' " I .' .,. ... ~ ';' 

-

Figure 10: Nucleation areas near <112> oriented edge (scale = 100X). 

OfUQ!NAl PAGE . 

OF ~R QUAtrr"Y 

89 

" 
PIN InP SOLAR CELLS ON Ge WAFERS 

Steven Wojtczuk and Stanley Vernon 

Spire Corporation 

Bedford, Massachusetts 

and 

Edward A. Burke 

Woburn, Massachusetts 

SUMMARY 

397'?¥f 

N95- 20511 

Indium phosphide (InP) P-on-N one-sun solar cells were epitaxially grown using a metalorganic 

chemical vapor deposition process on germanium (Ge) wafers. The motivation for this work is to 

replace expensive InP wafers, which are fragile and must be thick and therefore heavy, with less 

expensive Ge wafers, which are stronger, allowing use of thinner, lighter weight wafers. An 

intermediate InxGa1.

xP grading layer starting as Ino.49GClos1P at the GaAs-coated Ge wafer surface and 

ending as InP at the top of the grading layer (backside of the InP cell) was used to attempt to bend 

some of the threading dislocations generated by lattice-mismatch between the Ge wafer and InP cell so 

they would be harmlessly confined in this grading layer. The best InP/Ge cell was independently 

measured by NASA-Lewis with a one-sun 25°C AMO efficiency of 9.1 %, open-circuit voltage of 790 mV, 

fill-factor of 70%, and short-circuit photocurrent 22.6 mAlcm2

. We believe this is the first published 

report of an InP cell grown on a Ge wafer. 

Why get excited over a 9% InP/Ge cell? If we look at the cell weight and efficiency, a 9% InP 

cell on an 8 mil Ge wafer has about the same cell power density, 118 Wlkg (SOL), as the best InP cell 

ever made, a 19% InP cell on a 18 mil InP wafer, because of the lighter Ge wafer weight. As cell panel 

materials become lighter, the cell weight becomes more important, and the advantage of lightweight 

cells to the panel power density becomes more important. 

In addition, although InP/Ge cells have a low beginning-of-life (SOL) efficiency due to 

dislocation defects, the InP/Ge cells are very radiation hard (end-ot-lite power similar to beginning-otlife). We have irradiated an InP/Ge cell with alpha particles to an equivalent fluence of 1.6 x 1016 

1 MeV electrons/cm2 and the efficiency is still 83% of its SOL value. At this fluence level, the power 

output of these InP/Ge cells match the GaAs/Ge cell data tabulated in the JPL handbook. Data are 

presented indicating InP/Ge has more power output than GaAs/Ge cells at fluences in excess of this 

value. 

INTRODUCTION 

The cost, weight, and fragility of InP wafers have impeded InP cell use in space. Therefore, 

InP cells on light, strong, inexpensive silicon (Si) or Ge wafers are of great interest (ref. 1). This 

paper reports the results of a Phase I Small Susiness Innovative Research program seeking to replace 

the InP wafer on which the InP solar cell is epitaxially grown with a Ge wafer which has better 

properties (Table I). Since InP is fragile, thick -16 to 20 mil wafers are required tor strength. Ge has 

become a leading substrate for GaAs space solar cells, and is inexpensive if bought in large quantities. 

Although the density of Ge (5.3 g/cm3) is similar to InP, Ge is stronger and therefore a thinner Ge wafer 

than InP wafer can be used to increase the power density (W/kg) and lower the launch weight. 

91 

PAGE BlANK NOT FILMeD 
Table I Comparison of JII-V Space Cells on InP and Ge wafers. 

InP Cell InP Cell GaAs Cell 

on Ge Wafer on InP Wafer on Ge Wafer 

High-Volume Wafer Cost 20-50 180-250 20-50 ($/2" wafer) 

Wafer Density 5.32 4.81 5.32 (g/cm3

) 

Wafer Strength Strong Fragile Strong 

Typical Thickness 8 18 8 (mils) 

Wafer Weight 0.108 0.220 0.108 (g/cm2

) 

One-sun AMO SOL 9% 18% 18% 

Efficiency (NASA) (Typical) (Typical) 

SOL Cell Power Density 114 112 229 (W/kg) (Phase I) 

AMO EOL 8% 12% 6% (1016 1MeV e/cm2

) (Spire data) (Spire data) (JPL Handbook) Efficiency 

EOL Cell Power Density 102 75 76 (W/kg) 

InP cells have more defects when grown on Ge wafers, due to differences in crystal lattice 

constants of InP (5.87A) and Ge (5 .66A). This lattice constant mismatch is 3.7% between InP and Ge. 

Mechanical stress in the InP film grown on the substrate is relieved through the formation of dislocation 

defects in the InP film. The dislocations act as recombination centers, and resulting in a lower 

beginning-of-life (SOL) efficiency for InP heteroepitaxial cells (-10%) on Ge wafers than InP cells 

(-20%) on InP wafers. However, the lower SOL efficiency of InP heteroepitaxial cells is compensated 

by the lighter weight of Ge wafers compared to InP wafers (Table I), so that the SOL power densities 

for InP/Ge and InP/inP are similar, and we expect a higher end-of-life (EOL) power density from these 

InP/Ge cells compared to GaAs or Si cells at high fluences. 

The PIN InP/Ge cells have lower SOL efficiencies than Spire's InP cell record efficiency 19% 

NIP InP cells grown on InP wafers. However, we predict EOL efficiencies after high fluences will be 

similar"for these two InP cell types, since radiation damage will dominate the diffusion lengths in both 

cell types, instead of dislocations dominating the diffusion length in InP/Ge cells, leading to similar EOL 

efficiencies in both types. Therefore, because of their lighter weight, InP/Ge cells should have an EOL 

power density about twice that of InP homojunction cells. 

N-on-P InP-on-Si cells were investigated by Spire (ref. 2) with one-sun AMO efficiencies 

of 9.9% the highest reported to date (ref. 3). Our Phase II goal in this program is to achieve 

15% SOL efficiencies for P-on-N InP-on-Ge cells. We plan to work on lowering the dislocation density 

through an improved InxGa1.

xP grading technique between the Ge wafer and the InP cell. Sy using a 

PIN design, the need for a tunnel junction in NIP heteroepitaxial InP cell designs is eliminated. The I 

92 
tunnel junction is necessary in NIP designs due to outdiffusion of Si or Ge, N-type MOCVD dopants, 

from the Si or Ge wafers into the back P-Iayers of the NIP cell. P-on-N InP cells may more radiation 

resistant than even InP N-on-P cells, leading to higher EOL efficiencies (ref. 4). InP cells were 

grown on GaAs (ref. S) at NREL with efficiencies of 14% BOL at one-sun AMO (ref. 6). 

The NREL work shows that a 14% BOL efficiency can be achieved in a heteroepitaxial cell with 

significant dislocations. The dislocation density in the 9.9% AMO Spire InP/Si cells is 10x higher than in 

the NREL work. If we can reduce the dislocation density in our InP-on-Ge cells to the level achieved 

by NREL in its InP-on-GaAs work, a similar efficiency to the NREL work should be achieved, since Ge 

(S.66A) and GaAs (s .6SA) have similar lattice constants. The 1S% Phase II InP/Ge goal therefore 

seems reasonable. 

The power density of P-on-N InP/Ge cells at this early stage is already similar to the best N-onP InP cells on InP wafers, which have undergone much more development. Beginning-of-life efficiency 

(9% AMO) of the PIN InP/Ge cells is limited partly by the new P-on-N InP cell technology used for the 

first time in Phase I. We are now making 17% PIN InP cells on InP wafers, but at the time of this 

program effort the P-on-N InP control cells on InP wafers, our first, had reached only 12% efficiency, 

mainly due to too thick (-2000A) an emitter layer, compared to the 19% N-on-P InP cells on InP wafers 

with 300A emitters achieved after long development. As the performance of the PIN InP cell on InP 

wafer baseline technology increases, so should the InP/Ge cell performance since we would be starting 

out with a higher efficiency InP cell on the Ge wafer. It is important to realize that the best efficiency 

the Phase I InP/Ge cells could possibly have is the -12% efficiency of the control PIN InP cells on InP 

wafers. The BOL efficiency of the InP/Ge cell at time of the program was substantially limited by the 

PIN InP cell technology. This PIN InP cell technology should be improvable to levels approaching 20%. 

CELL STRUCTURE 

Table II shows the target epilayer structure used for the Phase I InP/Ge cells. 

Table II Epilayer structure of P-on-N InP cell on Ge substrate. 

Layer Material Doping Thickness Comments cm-3 11m 

Contact InGaAs is selectively etched from 

Cap InGaAs P, _1019, Zn 0.3 photoarea, but left under front grid 

metal to form ohmic contact 

Emitter InP P, _1018, Zn 0.2 Thickness tradeoff - thin better for 

QE; thick better for low resistance 

Base InP N, -1017 , Si 1.S 1.S 11m absorbs> 9S% of AMO light 

Back Surf. InP N, _1019, Si O.S Reflects minority carrier holes, 

Field enhances QE 

InxGa1_xP InP to N, 1019, Si 8 Lowers dislocation density due to 

Grading Ino49GCla51 P InP/Ge lattice-mismatch 

Nucleation GaAs N, 1019, Si 1 Easier to grow on GaAs 

Substrate Ge N, 1017, Sb 300 Eagle-Picher epi-ready Ge wafers 

93 
The epilayers are grown by MOCVO using trimethylindium, triethylgallium, and phosphine at 

76 torr and a low temperature, 600°C, to limit zinc diffusion and emitter junction depth. Oimethylzinc is 

being used for all P-type and silane for N-type doping, respectively. With an InP PIN cell design, we 

desire a thin emitter to limit surlace recombination loss and increase photocurrent; on the other hand, 

the emitter must be thick enough so that along with increased cell gridline coverage (4% shadow loss), 

a reasonable emitter sheet resistance 12R loss is obtained, compensating for the low maximum P-lnP 

emitter doping (-1 O,a cm·

3

, - 10X lower than N-lnP) and mobility (- 20X lower than N-lnP) of the P-lnP 

emitter. 

A PIN design is used to avoid a tunnel junction present in NIP designs. Germanium is an 

N-type dopant in III-V semiconductors in the metalorganic chemical vapor deposition (MOCVO) process, 

outdiffuses from the Ge wafer during epigrowth into the back of the cell. In an NIP design, Ge would 

create an opposing PIN junction in the P-type back layers of the cell. This parasitic junction must be a 

tunnel junction for the cell to pass current, increasing complexity in NIP cell designs. In a PIN deSign, 

the back N-type doping is simply increased by the Ge diffusion and is of little concern; no tunnel 

junction is needed. In addition, various NASA-Lewis papers indicate PIN InP cells may eventually be 

more efficient than NIP cells due to higher obtainable open-circuit voltages and amenability to surlace 

passivation. 

A high density of defects, mainly dislocations, form in the material to accommodate the latticemismatch (3 .7%) between InP and Ge. If these defects thread upward into the cell through the 

junction, they increase the dark current and act as minority carrier recombination sites, lowering the cell 

efficiency. For lattice-mismatched heteroepitaxial cells, grading layers are used to attempt to bend the 

threading dislocations harmlessly away parallel to the plane of the cell junction. In this program we 

used an InxGa,.xP grading layer starting with Ino.49Gaos,P lattice-matched to the GaAs-coated Ge wafer 

and ending with InP. This grading layer will be discussed more fully in future publications conceming 

InPISi solar cells. 

PRE-IRRADIATION CELL DATA 

Table III shows verified (courtesy of I. Weinberg and o. Brinker of NASA-Lewis) preirradiation 

InP/Ge cell data of similar PIN InP cells on InP, GaAs, and Ge wafers. The 11.9% control cells on InP 

wafers represent an upper limit of what the InP cells on Ge could achieve at the time of the program. 

Recently 17% PIN InP cells on InP were made, so that if the new InP cell growth parameters were 

used, a higher InP/Ge cell efficiency would be obtained than presented in this paper. Series resistance 

from I-V data for all Phase I cells was -0.5 n -cm2

, causing - 10 mV drop in VMAX for the -20 mNcm2 

photocurrent. The series resistance is dominated by the emitter sheet resistance. If we half the emitter 

thickness, we will double the series resistance . 

Table III AMO one-sun data of Spire Phase I cells. 

All one-sun AMO 25°C pre-irradiation '11 Vae Jse FF Cell 

Comments % mV mNcm2 % 10 

PIN InP cell on InP wafer (control) 11.9 848 22.9 84.2 5668-2626-2-8 

PIN InP cell on 8 ~m InGaP grade on 

GaAs-coated Ge wafer 9.1 792 22.6 69.8 5714-2795-1 -8 

(NASA-Lewis verified measurement) 

94 
CELL IRRADIATION 

A second PIN InP/Ge cell, of a slightly lower BOL efficiency (7.5%) than the best cell shown in 

Table III was mounted in a special test fixture for the destructive alpha irradiation test (equipment 

courtesy of C. Blatchley and C. Colerico of Spire). An Am-241 alpha particle source was used to 

irradiate the cells to explore how performance varies with radiation damage. Equivalent 1 MeV electron 

exposures were determined by a non-ionizing energy loss (NIEL) calculation (ref. 7) for the alpha 

source for a 1.5 cm separation between the 1 cm2 InP/Ge cell centered under the alpha source in 

vacuum. The calculation included effects of the angular incidence of particles at the cell edges offset 

from the centerline, so that the divergence of the alpha particle beam from its 0.5 cm aperture was 

taken into account. The alpha particle energy in the InP at the depth of interest is 3.643 MeV and the 

calculated NIEL value is 0.29 MeV-cm2

/g. The equivalent 10 MeV proton flux is 4.18 x 107 protons/ 

cm2

-s. The equivalent 1 MeV electron flux is 3.55 x 1010 electrons/cm2-s or 1.28 x 1014 electrons/ 

cm2

-hour. AMO efficiency data and quantum efficiency data were taken before irradiation, and after 1, 

8, 32, and 126 hours. The final 126 hour data set was equivalent to a fluence of 1.6x1016 1 MeV 

electrons/ cm2

. One of the advantages of using the alpha source for these experiments is that it is 

possible to obtain high equivalent electron fluences in short times due to higher damage rate of the four 

heavy nuclei (two proton, two neutron) of the alpha particles versus the lighter electrons. AMO data for 

this cell is at various fluences is shown in Table IV and Figure 1. 

Table IV InPIGe cell (7.5% BOL) one-sun 25°C AMO data at various equivalent electron fluences. 

(Alpha particle irradiation) II Vae Jse 

Equivalent # of 1 MeV electrons/cm2 % mV mAlcm2 

1.3 x 1014 

1.0x1015 

4.1 x 1015 

1.6x 1016 

"E 

41 

~ 

20 

~ 10 o 

-o 

.r:: 

7.5 774 

7.4 768 

7.0 751 

6.3 708 

PIN InP/Ge 7.5% BOL Cell: AMO Data 

(l. 

5 After 4.7x1 011 a/cm2 

(- 1.6x1 016 1 MeV ele/cm2) 

w....u..w...wu....L.l.~.l..U..L.LJ....o..........w~'-'-'-'.u..lLu...U.J..L.u ........... u..w 

23.4 

23.5 

22.7 

21 .6 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 

Photovoltage (V) 

FF 

% 

57.1 

56.2 

56.8 

56.8 

Figure 1 AMO I-V curves of an InPIGe cell before and after alpha irradiation. 

95 
Figure 2 shows measured absolute quantum efficiency of the cell before irradiation and after 

the maximum irradiation. As expected, most of the photocurrent loss seen in Figures 1 and 2 is due to 

lower quantum efficiency at the longer wavelengths , which are absorbed further from the cell junction 

and must diffuse to the junction to be collected . The radiation damage lowers the base (hole) diffusion 

lengths in these PIN cells slightly, as seen in Figure 2 near the InP cutoff wavelength (920 nm). The 

quantum efficiency at shorter wavelengths is affected very little, since, even though the diffusion lengths 

are lower in the emitter also, the emitter thickness (0.2 ~m) is still small compared to the emitter 

(electron) diffusion lengths (-1 ~m). 

Figure 2 

1.0 rr"""~~~~~~~~",,,,~~I~J 

t 0.9 

z 

!!! 0.8 

Co) 

~ 0.7 

III 

:I 0.6 ::::» 

~ 0.5 

CC 

; 0.4 

... 

CC 0.3 

z 

a:: III 0.2 ... 

>< III 0.1 

7.5% BOl PIN InP/Ge Cell 

After 4.7x1 011 ex Icm2 

( equiv. - 1.6x1 016 1 MeV ele/cm2 ) 

0.0 L.J......L--L...J.....l......L....l....L....LL.L....J...J.....J.....I.....1....L....L...J.....J.....J...J......L....I.....L.L..JL....L.i.....l......L..~ ... 

300 400 500 600 700 800 900 1000 

WAVELENGTH (nm) 

Absolute quantum efficiency of 7.5% BOL InPIGe cell before and after irradiation. Drop 

in longer wavelength QE indicates lower base diffusion lengths. 

CONCLUSIONS 

We have presented experimental data on PIN InP/Ge solar cells which indicate that this 

technology is promising for space power use in long space missions or in very high radiation orbits. 

The use of Ge wafers eliminates the need for costly, fragile, heavy, InP wafers. Even in its primitive 

current state of development, these 9% BOL InP/Ge cells appear to have higher power output and cell 

power density than either 19% BOL InP cells on InP wafers or 18% BOL GaAs cells on Ge wafers after 

a fluence of _1016 electrons/cm2 (Figure 3). Future work would center on increasing the InP/Ge cell 

performance so that it could compete with GaAs/Ge cells for space missions that do not require 

extreme radiation resistance. 

96 
Figure 3 

-

24----~ 

22 

9,d 105.d 

GaAs/Ge (JPl Handbook) 

('\I 20 E t----__ Phase IIlnP/Ge Goa 

u 18 

iE 16 

-t- 14 ~ 

e: 12 

5 10 

ffi 8 

~ 6 

D. 4 

I 

7.5% BOl Phase IlnP/Ge Cell 

Points: Measured Data 

2 Lines: Regression Fit 

o L-~~~ __ ~~~ __ ~~~~~~~ 

1013 1014 1015 1016 

FLUENCE (# of 1 MeVelectrons/cm2 ) 

Measured AMO power output versus electron fluence for PIN GaAsl Ge cells (from JPL 

Solar Cell Handbook) and PIN InPIGe cells (this work). Points are measured data; lines 

through pOints are simple 2nd-order polynomial regression fit. Line labeled Phase /I is 

simply a goal and is not measured data. This plot shows that even the primitive 

Phase I InPIGe cells have more power output after a fluence ot 1016 electronslcrrt than 

the current mainstay GaAslGe cells . However, this fluence is very high, and is likely 

only in high radiation (van Allen belt) orbits or tor very long (-10 year or more) missions 

in more standard orbits. The Phase /I goal shown indicates the point where these 

InPIGe cells could compete with GaAsl Ge tor more standard, low radiation missions. 

REFERENCES 

1. I. Weinberg and D.J. Flood, "Heteroepitaxial InP Solar cells for Space Applications," Prog. in 

Photovoltaics: Research and Applications, Vol. I, 1993, pp. 43-45. 

2. C.J. Keavney, S.M. Vernon, V.E. Haven, S.J. Wojtczuk, and M.M. AI-Jassim, "Fabrication of 

n+fp InP Solar Cells on Silicon Substrates," Appl. Phys. Lett. 54 (12), 1989, pp. 1139-1141. 

3. C. Keavney, S. Vernon, and V. Haven, "Tunnel junctions for InP-on-Si Solar Cells," NASA Cont. 

Pub. 3121, 1991 Space Photo voltaic Research and Technology Cont. , pp. 1-1 to 1-7. 

4. R.K. Jain and D.J. Flood, "Design Modeling of High-Efficiency p+-n InP Solar Cells," IEEE 

Trans. on Electron Devices, Vol. 40, No. 1, 1993, pp. 224-227. 

5. M.M. AI-Jassim, R.K. Ahrenkiel, M.W. Wanlass, J.M. Olson, and S.M. Vernon, "The 

Heteroepitaxyof InP and GainP on Silicon for Solar Cell Applications," Mat. Res. Soc. Symp. 

Proc. Vol. 198, 1990, pp. 235-246. 

97 
6. M.W. Wanlass, T.J . Coutts, J.S. Ward, and K.A. Emery, "High-Efficiency Heteroepitaxial InP 

Solar Cells," Rec. at the 22nd IEEE Phatava/taic Specialists Cant., 1991 , pp. 159-165. 

7. G.P. Summers, E.A. Burke, P. Shapiro, S.R. Messenger, and R.J. Walters, "Damage 

Correlations in Semiconductors Exposed to Gamma, Electron and Proton Radiations," IEEE 

Trans. on Nuclear Science, Vol. 40, No. 6, 1993, pp.1372-1379. 

98 
SESSION II 

/t?C::> 

THIN-FILM AND HIGH EFFICIENCY 

CELL DEVELOPMENT 

PMIlOI04N8 PAGE BLANK NOT F!l O 
STATUS OF FLEXIBLE CIS RESEARCH AT ISET1 

~ 7'775.3 

N95·20512 

B.M. Basol, V.K. Kapur, A. Minnick, A. Halani, and C.R. Leidholm 

International Solar Electric Technology (ISET) 

Inglewood, California 

SUMMARY 

Polycrystalline thin film solar cells fabricated on light-weight, flexible substrates are 

very attractive for space applications. In this work CulnSe2 (CIS) based thin film devices were 

processed on metallic foil substrates using the selenization technique. CIS deposition method 

involved reaction of electron-beam evaporated Cu-In precursor layers with a selenizing 

atmosphere at around 400°C. Several metallic foils such as Mo, Ti, AI, Ni and Cu were 

evaluated as possible substrates for these devices. Solar ·cells with AM 1 .5 efficiencies of 

9 .0-9.34 % and good mechanical integrity were demonstrated on Mo and Ti foils. Monolithic 

integration of these devices was also demonstrated up to 4"x4" size. 

INTRODUCTION 

Great advances have been made in polycrystalline thin film terrestrial solar cell 

technologies since early 1980's when the first promising laboratory devices with high 

efficiencies were demonstrated. These cells were fabricated on polycrystalline CdTe and 

CulnSe2 (CIS) layers and they had AM1.5 conversion efficiencies of around 10%. During the 

last decade, the polycrystalline thin film solar cell efficiencies have improved to over 15% 

range and the stability data obtained from these devices has been very encouraging. 

As the efficiency and the stability of the polycrystalline thin film solar cells have 

improved through the years, these devices have become more and more attractive for space 

applications where a reliable power source with high specific power is needed (refs. 1 and 2). 

Cells and modules fabricated on foil substrates also appeal to some specific terrestrial markets 

where flexibility is either required or preferred. 

CIS and related compound thin film solar cells have already demonstrated terrestrial 

conversion efficiencies of over 16% (the highest efficiency reported is 16.4% by NREL for a 

Cu (In,Ga) Se2 device). Preliminary tests also indicated that the radiation tolerance of CIS thin 

film cells was superior to that of single crystalline devices under high energy electron and 

proton irradiation. 

Besides their radiation resistance and promise of high efficiency, CIS thin film devices 

also offer to the space power market a high specific power and low cost. If these devices 

could be fabricated on light-weight substrates and if they could be monolithically integrated 

to form modules, they would become very competitive with the existing single crystal 

technologies even if their beginning-of-life efficiencies were lower than those of the single 

crystalline cells (ref. 3). The typical substrate for a high efficiency terrestrial CIS solar cell is 

a 0.3 cm thick soda lime glass sheet. The main thrust of our effort in this program was the 

1 Work funded by NASA SBIR II contract No. NAS3-26615 

101 
fabrication of these devices on thin metal foil substrates .using the selenization method. 

EXPERIMENTAL 

A "substrate" device structure with Foil/Mo/CIS/CdS/ZnO configuration was employed 

in this work. The 1-2 micron thick Mo layer was sputter deposited on the 1-2 mil thick foil 

substrate. CIS and CIGS (Cu(ln,Ga)Se 2) layers were grown by the selenization technique (refs. 

4 and 5). In the first step of this process thin layers of Cu and In, and in some cases Ga, were 

deposited on the metallic foils bye-beam evaporation. Thicknesses of the Cu and In layers in 

these precursors were typically 0.2 microns and 0.47 microns, respectively. Ga content was 

varied from 0% to 20%. During the second step of the process the precursors were reacted 

in a selenizing atmosphere containing H2Se gas at 400 ° C to form the selenide compounds. 

The selenization profile and the selenization period were varied to optimize conditions so that 

films of good electrical and mechanical properties could be obtained. 

Devices were completed by CdS and ZnO depositions. CdS layer was obtained using 

the chemical dip method. This technique utilizes a meta-stable solution containing a Cd source 

such as Cd-acetate, a sulfur source such as thiourea, and a complexing agent that controls 

the rate of release of the Cd 2 + ions into the electrolyte. ZnO films were deposited by the 

MOCVD technique to a thickness of 1-1.8 microns. Further details of the processing steps can 

be found in our previous publications (refs. 4 and 5). 

RESULTS AND DISCUSSION 

Flexible Substrate Selection 

We have evaluated various metallic foils as possible substrates for the growth of CIS 

layers. Some of the factors that were taken into consideration in these evaluations are 

indicated in Table I and they will be reviewed here. 

The selenization technique for CIS film formation involves a reaction step during which 

the Cu-In precursor layer is annealed in a reactive atmosphere containing H2Se gas. It is, 

therefore, essential that the substrates selected for CIS cells do not participate in the reaction 

between the Cu-In layers and the H2Se gas and/or they do not themselves react extensively 

with the H2Se atmosphere at elevated temperatures. We found Mo and Ti foils to be the best 

in terms of chemical stability in the reactive atmosphere of our selenization chamber. 

However, we also determined that AI and Ni foils could be utilized provided that a Mo layer 

of good mechanical integrity was interposed between the foil surface and the growing CIS 

film. This thin Mo layer was found to act as an effective diffusion barrier between the foil 

surface and the CIS film and between the foil and the selenization atmosphere. Any pinholes 

present in the thin Mo inter-layer deposited on the highly reactive foils of AI and Ni, however, 

would allow an interaction between the CIS film and the foil substrate through these defects. 

Such an interaction would give rise to the formation of areas in the growing film which were 

associated with undesirable Cu-In-AI-Se, or Cu-In-Ni-Se compounds. The parts of the CIS film 

with a defect free Mo inter-layer, however, were highly uniform suggesting that the Mo layer 

deposited over the AI and Ni surfaces was an effective barrier to selenization at 400 °C. The 

defect density of the Mo layers deposited on Ti and Mo foils was not a critical factor in 

determining the stoichiometric uniformity of the CIS films grown on such substrates with 

102 

l 
-- --- -----------_._--

limited reactivity. The main requirement for the Mo inter-layer in these cases was "good 

adhesion to the foil surface". Cu foils were extremely reactive and they could not be utilized 

in our application even if they were covered on both of their surfaces by Mo layers. Mo/Cu/Mo 

structures tested under selenization conditions quickly inter-diffused and Cu reacted with H2Se 

forming copper selenides. More details of our studies on the reactivity of foil substrates can 

be found in references 4 and 5. 

Handling of the thin foil substrates during processing is a practical factor that needs 

to be considered. Our experience showed that 1 mil thick Mo, Ti and Ni foils could easily be 

handled and they kept their mechanical integrity throughout the device fabrication steps. AI 

foils, on the other hand, tended to crease easily. Specially drawn "annealed" AI foils were 

better in terms of handling during the precursor deposition, but these substrates lost their 

"springy" nature after the high temperature selenization step and they again became 

susceptible to creasing. 

In terms of specific power, AI and Ti are the two attractive choices as indicated in 

Table I. Both of these foils would contribute only 0.6-1 kg/kW to the overall specific power 

of CIS modules with 1 OW Iftl output. Thermal expansion coefficient match between the CIS 

film and the substrate is best for Ti and, to a certain degree, Mo foils. 

Based on these factors and the experimental results, we first adapted Mo foil as the 

substrate because of our familiarity with this material as the back contact to CIS devices. 

Later we initiated work on Ti foil substrates which are more attractive in terms of their light 

weight and near-ideal coefficient of thermal expansion. 

CIS Films and Solar Cells 

It is very important to control the nature of the Cu-In precursor layers in the first stage 

of our CIS deposition process. The thickness uniformity, the degree of alloying between the 

Cu and In layers and the morphology of the resulting Cu-lil precursor film are all factors that 

determine the quality of the CIS layer obtained after the selenization step. In films containing 

Ga, the place of this element in the precursor stack also affects the morphology of the 

resulting compound film. Adhesion is of utmost importance for CIS layers, especially for those 

deposited on flexible substrates. An important source of poor adhesion between a film 

prepared by the two-stage technique and its Mo coated substrate is the stresses generated 

in the CIS layer during the selenization process. We have eliminated this problem by carefully 

engineering the precursor layers (ref. 4) and have successfully deposited well adhering CIS 

layers on 6"x6" flexible Mo and Ti foil substrates. 

Figure 1 shows the I-V characteristics of two CIS cells fabricated on flexible Mo (fig. 

1 a) and Ti (fig. 1 b) foils. The area of these devices was 0.09-0.1 cm2 and their AM1.5 

efficiencies were 9% (active area), and 9.34% (total area) respectively. Witness cells 

fabricated on glass substrates utilizing the same Mo layers, the same Cu-In precursors and the 

same selenization procedures yielded efficiencies in the 11-12% range. Study of the flexible 

cell parameters indicated that these devices, on the average, gave 30-40 mV lower Voe values 

compared to the glass based cells. The Jse values were also lower but only by 1-2 mA/cm2 • 

However, the parameter that was consistently low in flexible solar cells was the fill factor. 

While the FF values of the glass based cells were typically in the 0.65-0.75 range, this range 

was only 0.5-0.6 for the flexible devices. 

103 
The SEM of figure 2 shows a cross sectional view of the CIS layer on a flexible Mo foil. 

There are certain characteristics of the flexible CIS films that we noted in studying such 

micrographs. The morphology of the CIS layers deposited on flexible foils are quite different 

than those grown on glass substrates. The dome-like features that are commonly observed 

in SEMs taken from the surfaces of flexible CIS layers originate from the dome-li ke pores that 

can be seen at t he Mo/CIS interface in figure 2. The crystalline quality of the f lexible films is 

also inferior t o the crystalline qualit y of the glass based layers. This, we believe, is due to the 

different surface qualit ies of the t wo substrates. Foil substrate surfaces provide a large 

num ber of nucleation sites f or grain growth and the resulting small cryst als are not w ell 

oriented. CIS films grown on foil substrates do not show the columnar grain structure often 

observed in layers deposited on glass substrates. We have initiated work to address this issue 

and increase the flexible cell efficiencies to the 12-13 % range. 

Module Integration Studies 

Monolithic integration of devices fabricated on metallic substrates requires deposition 

of an insulating layer over the metallic foil, and then a series of scribing steps to interconnect 

and isolate the adjacent cells (fig. 3). Glass based integration techniques which use 

mechanical scribers can not be utilized in the foil cell integration process because of the fragile 

nature of the thin insulator. Possible interaction of the insulating layer with the selenization 

environment is another factor that needs to be taken into account. We have carried out 

module integration work on 4 "x4 " size foils and monolithically integrated 16 cells on such a 

foil. Although these submodules demonstrated voltage addition the fill factor values are 

presently low and they limit the efficiency at this time to below 5%. Work is in progress to 

improve the efficiencies of these submodules to the 6-8% range by addressing the fill factor 

issue. It should be noted that the data reported in this paper represents the highest efficiency 

flexible CIS cells reported to date and the very first demonstration of monolithic integration 

of CIS cells on a flexible metal foil substrate. 

REFERENCES 

1. Landis G.A.; and Hepp A.F.: Appl ications of Thin-Film Photovoltaics For Space. 

26th Intersociety Energy Conversion Engineering Conf., 1991 , vol. 2, pp.256-261. 

2. Boswell J .; Anspaugh B.; and Mueller R.: Thin Film PV Development at Phillips 

Laboratory. Proc. 23rd IEEE PV Specialists Conf., 1993, pp. 1324-1329. 

3. Armstrong J .H.; Pistole C.O .; Misra M.S.; Kapur V.K.; and Ba~ol B.M .: Flexible CIS 

Films and Devices For Space Applications. Proc. Space PV Research and 

Technology Conf. , May 7-9 , 1991 , NASA Lewis Center, Cleveland , Ohio, p. 19-1. 

4. Basol B.M .; Kapur V.K. ; Halani A .; and Leidholm C. : CIS Thin Film Solar Cells 

Fabricated on Flexible Foil Substrates. Solar Energy Materials and Solar Cells , vol. 

29, 1993, pp. 163-173. 

5. Ba~ol B.M.; Kapur V.K.; Halani A.; Minnick A.; and Leidholm C.: Modules and 

flexible cells of CIS. Proc. 23rd IEEE Photovoltaic Specialists Conf., 1993, pp. 426-

430. 

104 
TABLE I. EVALUATION OF VARIOUS METALLIC FOILS AS A SUBSTRATE 

FOR FLEXIBLE CIS SOLAR CELLS 

Reactivity Ease of handling ' 

Foil Substrate with HzSe through process 

Mo low very good 

Ti ·moderate very good 

AI high poor (creases) 

Ni high good 

Cu very high very poor 

(reacts and 

becomes brittle) 

• CTE for CIS - 7-9 ppm/DC 

•• Assuming 10W/ftI modules on 1 mil thick foils 

I 

(mA) 2 

V(v) 

0.1 0.2 

v = 0.405 V 

1 nc = 37.5 mA/cm2 so 

FF = 0.597 

11 = 9.0 % 

Area = 0. 1 cm2 

(a) 

Coeff. of thermal Contribution to 

expansion 

(x 106

/

0

C) 

4.8 

8.6 

2.3 

13.4 

16.5 

I 2 

(mA) 

J 

. specific power 

(kg/kW) 

2.29 

1.01 

0.60 

2.00 

2.01 

v (V) 

0.1 

Voc = 0.4 V 

l oe = 38.9 mA/cm2 

FF. = 0.6 

11 = 9.34 % 

Area = 0.09 cm2 

(b) 

.. 

Figure 1. Illuminated I-V characteristics of two flexible cells fabricated on a) Mo foil, b) Ti 

foil. Measurements were made under AM1 .5 illumination. 

105 
Figure 2. 

Figure 3. 

Cross sectional SEM 

of a foil/Mo/CISI 

CdS/ZnO structure. 

Structure of a monolithically integrated 

flexible CIS submodule. 

106 

, 

CdS 

ZoO 

Mo 

I 
~c;'7?S-~ 

N95-20513 

DEVELOPMENT OF A LIGHTWEIGHT, LIGHT-TRAPPED, THIN GaAs SOLAR CELL 

FOR SPACECRAFT APPLICATIONS 1 

Margaret H. Hannon, Louis C. DiNetta, Michael W. Dashiell, 

John R. Cummings, and Allen M. Barnett 

AstroPower, Inc. 

Newark, Delaware 

ABSTRACT 

This paper describes ultra-lightweight, high performance, thin, light trapping GaAs solar cells for 

advanced space power systems. The device designs can achieve 24.5% efficiency at AMO and 1X 

conditions, corresponding to a power density of 330 W/m2 . A significant breakthrough lies in the potential 

for a specific power of 2906 W/kg because the entire device is less than 1.5 J.Jm thick. This represents a 

440% improvement over conventional 4-mil silicon solar cells. In addition to being lightweight, this thin 

device design can result in increased radiation tolerance. The attachment of the cover glass support to 

the front surface has been demonstrated by both silicone and electrostatic bonding techniques. Device 

parameters of 1.002 volts open-circuit voltage, 80% fill factor, and a short-circuit current of 24.3 mAlcm2 

have been obtained. This demonstrates a conversion efficiency of 14.4% resulting in a specific power of 

2240 W/kg . Additionally, this new technology offers an alternative approach for enabling multi-bandgap 

solar cells and high output space solar power devices. The thin device structure can be applied to any 

III-V based solar cell application, yielding both an increase in specific power and radiation tolerance. 

PERFORMANCE CAPABILITY 

III-V materials such as GaAs make excellent candidates for thin devices because they are direct 

bandgap materials. The high absorption coefficient of such materials for light of an energy greater than 

the bandgap makes it possible to fabricate cells in which the thickness of the active region is considerably 

less than in indirect materials such as silicon. Light of photon energy greater than the bandgap is 

absorbed within the first few microns of entering a direct bandgap semiconductor, so an ultrathin device 

design is both feasible and advantageous. Conventional high performance GaAs solar cells are usually 

comprised of epitaxial layers of GaAs and AlxGa1_xAs formed on a GaAs substrate. When GaAs devices 

are fabricated on a thick GaAs substrate, the substrate acts only as a support and does not contribute to 

the overall performance of the device. 

The advantages gained from fabricating thin solar cells include a high power-to-weight ratio 

(specific power) which is important for space applications. In addition, with a sufficiently thin device 

structure (base thickness on the order of a diffusion length) the free carrier absorption is minimized and a 

light trapped device becomes feasible. Light trapping increases the effective optical path length with the 

use of a reflector and/or a textured surface. Incorporating light trapping into the device increases the 

performance by increasing the short circuit current, while the reduced GaAs base thickness lowers the 

reverse saturation current. Both of these effects enhance the open circuit voltage [ref. 1]. 

1 This research was supported in part by the Ballistic Missile Defense Organizationllnnovative Science 

and Technology branch and the Department of the Air Force and managed by Phillips Laboratory, Space 

Power and Thermal Management Division 

107 
rBecause the device is thin, back surface recombination becomes an important issue. 

Recombination of the carriers at the back surface is reduced by adding an AlxGa1_xAs (x > 0.5) back 

surface passivation layer. This layer also reflects carriers back to the p-n junction due to the built-in 

electric field. Because the GaAs base is thin « 2 microns), the carriers can reach the junction before they 

recombine. The front surface is also passivated by an AlxGa1_xAs (x = 0.85) layer which has a large 

indirect bandgap. There is very little absorption in such a layer, allowing light into the underlying cell. 

Because of the good lattice match to the GaAs, this layer eliminates the surface states and other 

imperfections on the GaAs p-n junction surface that would ordinarily result in a high recombination velocity 

and decreased diffusion length. 

The use of liquid phase epitaxy (LPE) to fabricate thin devices offers significant advantages over 

other techniques such as MBE and MOCVD. This technique produces high quality material while 

maintaining low cost and simplicity. Inherent to the LPE technique is the fact that the dislocation density of 

the epitaxial films produced is generally lower than the starting substrate. Therefore, the material is 

superior in terms of diffusion length and lifetime. These benefits are partly attributed to the tendency of 

impurities to segregate to the liquid (solvent) as opposed to the solid (epitaxial film). The ability to grow 

multiple layers of controlled electrical conductivity is also useful in the proposed device design. 

Segregation coefficients are well known so that the proper conductivity type and carrier concentrations 

can be obtained in the epitaxial films. Phase equilibria for the AI-Ga-As system have been extensively 

studied, resulting in the ability to precisely control composition. 

The high efficiency and light weight of the cover glass supported GaAs solar cell can have a 

significant impact on space solar array technology. Fig. 1 shows the specific power (power to weight 

ratio) and power density of several candidate solar cells. AstroPower's GaAs solar cell design offers a 

440% increase in specific power over that of a 14.5% efficient silicon solar cell. The specific power is 

calculated assuming a 3-mil cover glass and a 1-mil silicone adhesive on the front surface of the solar cell. 

4000 400 

3200 

~ 2400 

~ 1600 

800 

o 

300 

N 

E 

:> > 200 

100 

o t--'- L..L...--L-j-.'--.......... ---I + - ....... -+- -== ........ ---' 

Silicon 

14.5% 

GaAs on 

Ge Substrate 

19% 

Thin Silicon Thin GaAs 

19% 24.5% 

W/m2 (AMO, 1X) 

W/kg (AMO, 1X) 

Fig. 1. Comparison of specific power and power dens;(y of candidate space solar cells. 

AstroPower's approach combines the technology for a thin, light trapped GaAs solar cell with the 

electrostatic bonding of GaAs to glass and a coplanar back contact technology. Rather than working with 

p-type front contacts that interfere with the bonding of the glass to the solar cell, contact to the p-type 

layers is made from the back of the device. This all back contact design also eliminates grid shading 

108 

_J 
which further enhances the performance of the GaAs solar cell. Since both contact grids are located on 

the back surface, cell degradation from grid flattening and metals migration, associated with electrostatic 

bonding of raised contact devices, is eliminated. Also eliminated are the bonding difficulties and the low 

quality bond yields that typify electrostatically bonded, raised contact solar cells. 

In order to obtain the highest efficiencies from thin (1 .65 micron) GaAs solar cells it is necessary 

to incorporate a high degree of light trapping in the cells. To generate the same level of current from a thin 

device as is possible from a conventionally thick GaAs cell, the optical path length of the light must be 

extended beyond the physical thickness of the device. The light must travel obliquely and be internally 

reflected many times, allowing more of the light to be absorbed at a given thickness. Chemical micromachining or random texturing can be achieved on the front or back surfaces to redirect the light at an 

oblique angle. Scattered light which is obliquely incident on the front surface at angles less than the 

critical angle (160 for GaAs), will be totally internally reflected. Such optical confinement leads to effective 

optical path lengths 2 to 10 times greater than the thickness of the active layers. The factor by which the 

optical path is increased due to light trapping is called the z-ratio. For example, when the z-ratio is equal 

to five, the optical absorption for confined light is equivalent to that of a solar cell with a thickness five 

times greater. 

The optimum reflector should have the maximum reflection over the appropriate wavelength 

range. This aspect of light trapping is important since many reflector options (e.g. quarter-wavelength 

dielectric films or distributed Bragg reflectors) are optimized for maximum reflectance at one wavelength 

and exhibit high reflectance over only a narrow bandwidth. A 1-micron thick base requires a back reflector 

which is effective over the wavelength range of 730 to 880 nm [ref. 1]. 

Using the LPE technique to grow a thin structure, bonding to a cover glass, and ultimately 

removing the GaAs substrate allows for access to the back of the active region of the device. Thus an 

optical reflector (such as Au or Ag) can be applied directly to the back surface. This offers significant 

advantages over other techniques such as the use of Bragg reflectors grown by MOCVD on GaAs 

substrates. The spectral width of Bragg reflectors is restricted, and to achieve a z-ratio higher than 2, 

multiple Bragg reflectors must be used. As pointed out by Tobin et aI. , [ref. 2], "the added complexity of 

multiple Bragg reflectors" does not make this a practical approach. The use of the appropriate metal 

reflector on the back surface provides reflection over a broad spectral range. This technology also 

removes the excess weight of the substrate thus Significantly increasing the specific power. 

Modeling the thin GaAs solar cell shows benefits similar to those achieved in light trapped silicon. 

For silicon devices, reducing the thickness of the device decreases the reverse saturation current, while 

trapping the light leads to an increase in the short circuit current. In GaAs, however, the current gains are 

smaller and most of the increased performance is realized from enhanced open circuit voltage. The three 

most important features which lead to improvements in the efficiency of GaAs thin-film solar cells are: 

increased optical absorption, improved collection effiCiency, and photon recycling [ref. 3]. Photon 

recycling is when photons generated by radiative recombination are optically confined so they can be reabsorbed to generate minority carriers again. The enhanced optical absorption and improved collection of 

minority carriers provide a modest increase in the short circuit current (lsd. Because the solar cell volume 

is reduced by thinning the device, the bulk recombination is reduced which reduces the dark current. The 

reduced dark current and improved Isc result in an increase in the open circuit voltage. Also contributing 

to an increase in the open circuit voltage is the fact that higher carrier concentrations can be used to 

further reduce the dark current since a low minority carrier diffusion length can be tolerated in a thin 

device. As light trapping increases for a given back surface recombination velocity, the solar cell 

efficiency increases. When light trapping is considered, the pIn structure is more efficient than an nIp 

structure since the long base diffusion length in the nIp structure is not as important when the solar cells 

are very thin. 

109 
RADIATION RESISTANCE 

Radiation damage is the primary mechanism of degradation of GaAs solar cells deployed in 

space. This gradual degradation in solar cell performance is due to a reduction in the minority carrier 

lifetime that results from cumulative damage to the crystal lattice. As high energy particles bombard the 

cell, the number of recombination centers is increased, resulting in a decrease in the minority carrier 

lifetimes. The fact that light is absorbed in a shallow depth indicates that GaAs should have a better 

radiation resistance than silicon. Since the minority carrier transport is over much smaller distances, the 

diffusion length can be reduced by irradiation to much smaller values before having a significant effect on 

the carrier collection at the junction [ref. 4] The thin light trapped GaAs solar cell design further enhances 

the radiation tolerance because damage created several micmns into the material by high energy particles 

has no effect on photo current collection. The recombination region is thinner, thus increasing the 

resistance to high-energy radiation. The cover glass can be specified to screen out low energy particles 

corresponding to the chosen orbit, which normally cause damage at the surface [ref. 5]. 

Optimized emitter thickness and absorber layer doping can also contribute to radiation tolerance. 

The emitter thickness is kept below 0.5 microns in order to reduce the distance which minority carriers 

generated near the surface must travel to be collected . Because the entire device is less than 2 microns 

thick, carriers generated deeper in the material can still reach the junction before recombining . This 

becomes important when radiation has decreased the minority carrier diffusion length. 

In conventional thick GaAs solar cells, the base layer carrier concentration is kept below 

3x1 017 cm-3 to improve the end-of-life (EOL) efficiency. In our thin device this effect is not as important 

because of the reduced dependence on diffusion length. Higher carrier concentrations are incorporated in 

order to reduce the dark current and thus enhance the open circuit voltage, while maintaining a radiation 

tolerant device. 

EXPERIMENTAL RESULTS 

A schematic cross-sectional representation of the AstroPower prototype thin GaAs solar cell 

design is shown in Fig. 2. 

Cover Glass 

Contact 

GaAs ~ Bonding agent 

cap layer ~='_., < Si,N. 

AIGaAs window layers < ------------ ~ GaAs emitter, 0.5 um 0.075 ).1m 

~iiiiii -ii~~II!. •• iIIliiiiiiiillliiiiiiiii iiiiiiiii~ ( G aAs base , 1.0 u m 

Si,N. ~ 

Fig. 2. Ultra-thin GaAs solar cell with light trapping. 

Back surface 

reflector/co ntact 

A key requirement for high performance ultra-thin GaAs solar cells is the incorporation of light 

trapping into the devices. Light trapping was demonstrated by growing a thin (1 .5 micron) GaAs base 

layer on an AIGaAs passivating layer using liquid phase epitaxy. The sample was then bonded to glass 

and the substrate was removed. 

110 

J 
Comparisons were made between samples with and without a metal reflector, and the 

effectiveness of the reflector in projecting photons back toward a junction was determined. Silver was 

used as the reflector in this case. Reflection + transmission (R + T) measurements were performed 

before and after the substrate removal. The absorption can be obtained from the reflection and 

transmission data(absorption = 1 - (R+ T)). These results are shown in Fig. 3. 

100 

90 

80 

70 

~ ----

.......... 60 - c 

0 

li 50 ~ 

0 

en 

.0 40 --GaAs substrate 

ro -D- Thinned 

30 - - - - Thinned, with reflector 

20 

10 

0 

350 450 550 650 750 

wavelength (nm) 

----, 

\ 

\ 

\ 

\ 

\ 

\ 

GaAs 

bandedge 

\ 

850 

\ 

\ 

\ 

\ Photons 

\ generating 

\heat 

\ 

\ 

I 

--------

950 

Fig. 3. Absorption results for: 1.5 11m GaAs LPE layer on a GaAs substrate; 1.5 11m glass bonded GaAs 

layer, and glass bonded 1.5 11m GaAs layer with a reflector. 

As expected, the amount of light absorbed decreases as the GaAs structure is thinned, 

particularly over the 550 nm to 850 nm range where there is a higher flux density and the photons are 

more weakly absorbed. Application of a silver reflector resulted in increased absorption which closely 

matched the measurements of the thin material on the GaAs substrate. As much as 70% of the light is 

absorbed over the 550 nm to 850 nm range. By incorporating a silver reflector, we have been able to 

successfully light trap a 1.5 micron thick structure of GaAs. This will enable high short circuit currents to 

be obtained on a thin, ultra-lightweight GaAs solar cell. 

Note that there is little absorption of sub-bandgap photons in the thinned material both with and 

without a reflector. This is advantageous for space solar cells because sub-bandgap photons that are 

absorbed generate heat in the device but do not contribute to the efficiency. As shown in Fig. 3, the 

structure on a GaAs substrate absorbs as much as 70% of the photons at 900 nm while the thinned 

structure with a reflector absorbs only 4% of the photons at 900 nm. Standard thick GaAs solar cells 

absorb these lower energy photons in the substrate. 

Device layers are grown by liquid phase epitaxy (LPE) [ref. 6]. The front and back surfaces are 

passivated by an AlxGa1_xAs window layer in order to reduce the surface recombination. The GaAs base 

layer is approximately 1 micron thick and the emitter layer is formed by diffusing the p-type dopant during 

the growth of the front passivating (window) layer. The junction depth is easily controlled by adjusting the 

window layer growth time. 

111 

l __ _ 
Fig. 4 shows the quantum efficiency of a free-standing 1.65 ~m , 1-cm2 device. The short circuit 

current, as corrected for grid shading and reflection losses, was 29.13 mNcm2

. 

100 

>- () 

80 

c 

Q) 

'u 60 ~ W;? 

E~ 

:::J 40 ..... 

c 

co 

:::J 

a 20 

o 

/ ,/ 

/ 

/ 

,/ 

,/ 

,/ 

,... / 

................ ... ... 

.... 

. . 

_ .... 

-

.. ..... 

.",.."...,----- ....... ...-...-------.".. 

Reflection 

External Quantum 

Efficiency 

--Internal Quantum 

Efficiency 

. ..................... ... "' ........... .. ........... 

350 400 450 500 550 600 650 700 750 800 850 900 

Wavelength 

(nm) 

Fig. 4. Quantum efficiency results for free-standing device F123148. 

The results of the current-voltage measurement (F12314B) is shown in Fig. 5. This thin solar cell 

demonstrated an efficiency of 14.4%, as measured. From the quantum efficiency curves, it can be seen 

that the antireflection coating is not properly optimized and results in a lower than optimal current 

generation from 350 to 600 nm. When corrected for reflection losses, the potential of this material would 

yield a 17.3% efficiency at AMO. 

Cell Data: F12314B 

1.001 v. 

24.3 mNcm2 

Fill Factor 80% 

Efficiency 14.4% 

, I I I I I I - - - - 1- ____ 1 ____ _ ~ ____ I- ____ ~ ____ .. ____ ...j. ___ _ 

I I I I ! 

, 

, tit I I I 

- - - _1 - ____ 1 __ _ __ L ____ L ____ L ____ ..L __ __ .J ___ _ 

I t I I I 

I I I 

I , 

- - - _1- ____ 1 ____ _ L ___ _ L _ _ _ _ _ ___ .1 ___ _ 

I I I 

I 

I - - - -,- - - - -,- - - - -

I , I 

-- - -.-- - - -,- - - - - ,- - - - -

, 

I I ---- 1"-- -- ,- ---

I 

I - ---1--- 1----

Fig 5. Current-voltage measurement results for free-standing device F123148 

112 
The best parameters obtained from conventional GaAs devices fabricated at AstroPower were an 

open circuit voltage of 1.020 V. a short circuit current of 35.6 mAlcm2

• and a fill factor of 82.7%. This 

demonstrates the potential for a solar cell efficiency of 22.2%. As can be seen. the measured open-circuit 

voltages and fill factors are close to the best devices fabricated on GaAs substrates. Optimization of the 

thickness. doping. and antireflection coatings will yield an increase in the performance of the thin GaAs 

solar cell 

The improved device design which utilizes electrostatic bonding and an all back contact 

technology is shown in Fig. 6. The p-type region is diffused form the back of the device to the emitter 

after thinning. The temperature required for this diffusion step necessitates a high temperature survivable 

electrostatic bond. This superior solar cell design solves many fabrication problems and enhances the 

manufacturability of the high performance GaAs solar cell. Development of this design is currently in 

progress. 

Antireflection'----t-_, 

coating 

Electrostatically Bonded 

Cover Glass 

AIGaAs window=---r_+ t----------------j ~ GaAs emitter, 0.5 ~m 

075~ ~ 

i ~ f ! i ~ Vapor diffused ~MM M.MMM~M __ : --------- --l r. ,---------;-___ p-region 

. , , , 

, 

, c-----'l .. GaAs base, 1.0 m 

Dielectric ------.·~II;;'~II~;t;;iIIliII;)r.iII~ 

'.ok ""'00 / reflector/n-type contact 

'-----n-type contact 

'------ p-type contact 

Fig. 6. Electrostatically bonded, all back contact, ultra-thin GaAs solar cell. 

CONCLUSION 

The results of this program have demonstrated the feasibility of the ultra-lightweight, high 

performance, thin, light trapping GaAs solar cell. This is a high payoff program and the resulting 

applications can have a dramatic positive effect on space solar power generation. Development of the 

thin light trapped GaAs solar cell will result in a new class of GaAs solar cell designs that can replace 

conventional GaAs solar cells because of their high specific power, radiation resistance, and durability. 

REFERENCES 

[1]. Honsberg and A.M. Barnett, "Light Trapping in Thin Film GaAs Solar Cells", Proc. 22nd IEEE 

Photovoltaic Specialists_Conference, Las Vegas, Nevada, (1991), pp. 262 - 267. 

[2]. TOBIN, S.M. VERNON, M.M. SANFACON, AND A. MASTROVITO, "Enhanced Light Absorption in GaAs 

Solar Cells With Internal Bragg Reflectors", Proc. 22nd IEEE Photovoltaic Specialists Conf , Las 

Vegas, Nevada, (1991), pp. 147 -152. 

113 
! -- --

[3]. LUSH AND M. LUNDSTROM, "Thin Film Approaches for High Efficiency III-V Cells", Solar Cells, 30 

(1991). 

[4]. MARKVART, "Radiation Damage in Solar Cells", Journal of Materials Science: Materials in Electronics, 

Vol. 1 , No.1, pp. 1-12, (1990). 

[5]. FLOOD AND H. BRANDHORST, "Space Solar Cells", Current Topics in Photovoltaics, T.J . Coutts and J.D. 

Meakin ed., Harcourt Brace Jovanovich, Publishers, (1987). 

[6]. NELSON, RCA Review, 24 (1963). 

114 
l 

N95·20514 

ADVANCES IN POLYCRYSTALLINE THIN-FILM PHOTOVOLTAICS 

FOR SPACE APPLICATIONS 1 

Bruce R. Lanning, Joseph H. Armstrong, and Mohan S. Misra 

Martin Marietta Corporation 

Denver, Colorado 

INTRODUCTION 

Polycrystalline, thin-film photovohaics represent one of the few (if not the only) renewable power sources 

which has the potential to satisfy the demanding technical requirements for future space applications. 

The demand in space is for deployable, flexible arrays with high power-to-weight ratios and long-term 

stability (15-20 years). In addition, there is also the demand that these arrays be produced by scalable, 

low-cost, high yield, processes. An approach to significantly reduce costs and increase reliability is to 

interconnect individual cells series via monolithic integration. 

Both CIS and CdTe semiconductor films are optimum absorber materials for thin-film n-p heterojunction 

solar cells, having band gaps between 0.9-1.5 ev and demonstrated small area efficiencies, with cadmium sulfide window layers, above 16.5% (Ref. 1,2,3). Both CIS and CdTe polycrystalline thin-film cells 

have been produced on a laboratory scale by a variety of physical and chemical deposition methods, 

including evaporation, sputtering, and electrodeposition. Translating laboratory processes which yield 

these high efficiency, small area cells into the design of a manufacturing process capable of producing 1-

ft2 modules however, requires a quantitative understanding of each individual step in the process and its 

(each step) effect on overall module performance. With a proper quantification and understanding of 

material transport and reactivity for each individual step, a manufacturing process can be designed that is 

not "reactor-specific" and can be controlled intelligently with the design parameters of the process. 

Development of a thin-film, manufacturing process depends not only on the scalability of the process but 

on the overall fixed and operating costs, the environmental compatibility (Le., material utilization/waste 

minimization with minimal heahh and safety risks), and the reproducibility/stability of the process. For this 

reason, the selection of deposition processes at MMC was influenced by: 1) cost; 2) environmental compatibility; and 3) reproducibility/stability of the process. In the development of CdTe and CIS devices at 

MMC therefore, CdTe films are being deposited by electrodeposition and CIS films are being deposited 

by DC, cylindrical magnetron sputtering. Both of these processes are scalable, low-cost processes with 

relatively minimal environmental impact and a discussion of the development of these processes is presented in this paper. 

The objective of this paper is to present an overview of the current efforts at MMC to develop large-scale 

manufacturing processes for both CIS and CdTe thin-film polycrystalline modules. CIS cells/modules are 

fabricated in a "substrate configuration" by physical vapor deposition techniques and CdTe cells/modules 

are fabricated in a "superstrate configuration" by wet chemical methods. Both laser and mechanical 

scribing operations are used to monolithically integrate (series interconnect) the individual cells into modules. Resuhs will be presented at the cell and module development levels with a brief description of the 

test methods used to qualify these devices for space applications. The approach and development 

efforts are directed towards large-scale manufacturability of established thin-film, polycrystalline processing methods for large area modules with less emphasis on maximizing small area efficiencies. 

1 This work is supported by Martin Marietta Independent Research and Development (I R&D) Project D-17R, 

"Photovoltaic Technologies" 

115 
CELL DEVELOPMENT 

A part of the development of a commercial scale process for thin-film polycrystalline modules is an understanding of each of the individual processing steps and its correlation to device performance and reliability. 

In this section, a summary is presented of each of the processing steps for both CIS and CdTe devices. 

Copper Indium Diselenide 

CIS heterojunction cells are deposited onto flexible or rigid glass substrates in the "substrate configuration"; substrate, molybdenum back contact, CIS absorber, CdS window layer, ZnO transparent conductive oxide, with a metallized top grid contact. Conventional CIS devices have been optimized on molybdenum-coated glass substrates and little has been reported on the performance of CIS films on flexible substrates. The issue of flexible substrate requirements is discussed along with the current approaches and 

results for producing thin-film CIS 

Substrate - For the rigid cells, borosilicate glass is an inexpensive substrate which matches well with the 

thermal characteristics of the semiconducting layers and is therefore the substrate of choice for a rigid 

module. Molybdenum can be uniformly sputtered onto glass substrates over 1 ft2 areas and depending 

on the characteristics of the sputtering chamber (i.e., planar or cylindrical cathodes, oxygen level, etc.), 

the molybdenum will have a surface roughness similar to the glass substrate with a well-defined texture. 

Surface profile and texture have been observed to have an effect on overall device performance and the 

selection of the substrate can have an effect on the roughness and texture of molybdenum back contact. 

Although flexible substrates offer inherent processing advantages in the development of a commercial 

scale process as well as meeting technical goals for space applications which are not obtainable with 

rigid substrates, substrate flexibility adds complexity to the fabrication process. Typical requirements for 

a flexible substrate in a CIS module would include such things as: 1) surface finish/profile; 2) high temperature stability (up to 550· C); 3) thermal compatibility (CTE); 4) insulating (dielectric breakdown voltage); 5) chemicaVvacuum stability; 6) cost and availability; and 7) flexibility versus strength. Each class 

of materials, whether it be metal, ceramic, or plastic, has certain drawbacks as a flexible substrate 

although the leading candidates for flexible substrates are polyimide-type plastics, "metal organic-based" 

flexible glasses, coated metallic foils, and mica sheets. 

Results from thermal gravimetric analysis have shown that nearly all the 'high-temperature' plastics tested out-gas to some degree above 400°C (this does not include evaporation of water above 100°C). This 

not only alters the properties of the material but can effect the adherence of the semiconductor films during deposition. CIS devices have been fabricated on polyimide films although efficiencies were not significant to report. 

Because of their mechanical toughness, high temperature stability, cost, and availability, coated and 

uncoated metallic foils have been used in the development of flexible CIS devices. Unlike the smooth 

glass substrates which have been used to produce the highest efficiency cells reported by others, metallic foil substrates have a rougher surface profile which can effect device performance. 

The differences between molybdenum coatings on the flexible foils and glass substrates were subtle. 

Peak to valley variations in the surface profile of the metallic foil were on the order of 200 nm whereas 

116 
I 

L.. _ 

the variation across a typical glass substrate is on the order of 10 nm. The molybdenum back contact 

films on metallic foils were highly textured with predominantly a (110) peak in the glancing incidence diffraction spectra. In comparison to single crystal molybdenum, the (110) peak of the sputtered molybdenum film on metallic foil broadened and shifted to higher d-spacings (i.e., residual tensile stresses after 

the molybdenum deposition and selanization processes). Molybdenum films on the glass substrate were 

also highly oriented although the (211) was the strongest peak and there was not any detectible broadening or shifting of the primary peaks. 

Surface texture of the substrate is certainly a factor in the deposition, nucleation, and growth behavior of 

semiconducting films. This as well as other subtle factors, such as the surface emissivity/absorptivity of 

the substrate during thermal processing, can effect the photoresponse of a polycrystalline CIS device. 

Molybdenym Back Contact - Depending on the substrate, a 200-1,000 nm thick coating of molybdenum 

is sputtered onto the surface. Substrate temperature, sputter rate and pressure are used to control the 

adhesion and stress state of the film; oxygen partial pressure is also used as a parameter to control the 

properties of the molybdenum film. To improve adhesion of the CIS absorber layer, a graded molybdenum structure can be deposited with copper to produce a pure molybdenum layer at the substrate/contact interface and a pure copper layer at the CIS/contact interface (as reported in the literature (Ref. 4), 

copper at the back surface of CIS would produce a "p+" structure in the CIS absorber layer and improve 

the ohmic contact). 

CIS Absorber Layer - Reproducibility and uniformity of the CIS absorber layer is one of the most important factors in the performance and manufacturability of large-area modules. Since copper, indium, and 

selenium are three elements with considerably different properties (i.e., melting points, vapor pressures, 

oxygen solubility, conductivity, etc.), the formation of a single copper indium diselenide phase over large 

areas, both kinetically and thermodynamically, will depend on the processing order or deposition 

sequence and the uniformity of the film deposit(s) over large areas. Lateral compositional and thermal 

gradients across a large area can result in the formation and microsegregation of secondary phases and 

since these gradientslheterogeneties are not significant in the "through-thickness dominated" growth 

behavior of a small area, the processing steps to produce high efficiency small area devices may therefore be limited over larger areas. 

The approach in the development of a large scale process has been to select processing sequences for 

CIS which minimize the number of reaction pathways and secondary phase formation while at the same 

time, select deposition processes for each sequence which are reproducible and scalable. Two of these 

approaches are: 1) Culln Bilayer approach where copper and indium are successively deposited on a 

molybdenum back contact and then reacted with either elemental selenium or hydrogen selenide vapor; 

and 2) Selenized Bilayer approach where first, indium and selenium are deposited/reacted to form indium 

selenide, followed by copper and selenium deposition to form copper selenide, and finally, the two selenized bi/ayers are reacted to completion in a selenium atmosphere. The Culln bilayer is a low temperature (-400°C) approach with a small number of easily controllable, processing steps, demonstrated largearea scalability, and excellent substrate adhesion. Since a number of reaction pathways are possible 

with this type of approach, stable binary phases (as well as other types of microsegregation) can form 

along with the copper indium diselenide phase and degrade the performance of a large-area module. 

The selenized bilayer approach on the other hand, reduces the total number of possible reaction pathways by the formation of essentially two stable intermediate binary phases which, when reacted to com117 
pletion, can only form a single CIS phase (refer to a CU(2_x)Se/ln2Se3 binary phase diagram). Although 

this type of approach minimizes formation of impurities by limiting the number of reaction pathways, the 

number of processing steps and substrate temperature are increased in comparison to the Culln bilayer 

approach. In addition, reported efficiencies for this type of approach have only been demonstrated on a 

small scale. 

To evaluate emerging CIS technology, Martin Marietta (MMC), in cooperation with the National 

Renewable Energy Laboratory, designed and fabricated a CIS flight experiment with cells based on 

NREL's selenized bilayer process which has demonstrated 16.4% efficiency in AM1 .5 insolation this 

year. In this case, gallium was added to the CIS to improve the bandgap as well as efficiency. The 

experimental array fabricated for the SAMMES flight experiment is scheduled for launch later this year. 

Active area efficiency for these devices calculated at 13.5% in AMO as measured by a pulsed solar simulator at Spectrolab during final assembly. It is anticipated that there will be at least one year of data in 

orbit from this experiment. 

CdS Window Layer - A thin, 40-60 nm - thick, CdS film is deposited onto the CIS absorber layer by the 

chemical bath deposition process. Cadmium to sulfur ratios in solution have been varied between 10 and 

50 to optimize both film properties (i.e., adherence and uniformity) and yield from this batch process. 

Actual deposition occurs at 85°C in a buffered solution between 9.0 and 10.0 pH and processing time is 

less than 4 minutes. Reproducibility and uniformity of the thin CdS layer have been demonstrated over 1 

ft2 areas. 

ZnO Top Contact - Since the approach for the molybdenum back contact and CIS absorber layers has 

been to utilize the potential scalability of a cylindrical, DC-magnetron sputtering process, a similar development strategy was used in the case of the ZnO top contact (ZnO can also be deposited by RF sputtering although the results presented in this paper are with a DC power source). ZnO films were sputtered 

from an 8" diameter, hot pressed target containing 98 wlo ZnO and 2 wlo AI20 3 and the transmittance 

response for a 902.5 nm-thick film is presented in Figure 1 (a). The transmittance response for an 

uncoated substrate is included for reference. 

With DC magnetron sputtering of ZnO, film properties are strongly dependent upon both the processing 

parameters and the target/substrate geometry. To determine the effect of substraje location with respect 

to the target source, glass witness coupons were positioned at increasing distances from the target 

source (normal to the target) and at positions away from the center of the target (parallel to the target surface). Results from this series of tests is shown in Figure 1 (b) in which the film resistivity is plotted as a 

function of location relative to the target center. Within the region defined by the "racetrack" of the target 

(Le., -10 cm), the ZnO film resistivities are less than 1 x 10-3 ohm-cm for all but the 10 cm, target-to-substrate distance. In the fabrication of 10 cm x 10 cm CIS device therefore, processing parameters can be 

varied to first deposit a thin, high resistivity ZnO film on top of the CdS, followed by the low resistance film 

referred to above. 

The "roll-off" in film resistivity for this 8" target occurs at around 15 cm relative to target center and there 

is a distinct increase in resistivity across the racetrack region. Degradation of film properties in the 

vicinity of the target racetrack is well documented in the literature and poor resistivities are said to result 

from bombardment by energetic neutral and negatively charged oxygen atoms. Such bombardment is 

believed to cause lower carrier mobilities and concentrations through a decrease in grain size, a mixed 

118 
-- ._- ------

crystalline orientation, and a higher defect density. An additional cause for the variations in resistivity 

may be due to the variations in the aluminum doping density. A plot of atomic ratio of AVZn as a function 

of the location relative to the target center is presented in Figure 2. Based on these results, the increase 

in resistivity across the racetrack zone correlates with the drop in AVZn ratio. By incorporating existing 

techniques to control neutral and negatively charged oxygen atoms and minimize racetrack effects, ZnO 

will then be uniformly deposited over a 30 cm diameter by DC magnetron sputtering. 

0.9 

0.8 

0.7 

£! ~ 0.6 

.~ 0.5 

ra 

.= 0.4 

0.3 

0.2 

0.1 

0 

a) 

Figure 1 

Location Near Target Center 

uncoated substrate 

-3 1.5(10 )O-em 

500 1000 1500 2000 2500 

Wavelength (nm) 

Film Resistivity (O-em) 10y---------------~ __ ~ ______ , 

.1 

.01 

.001 

-25 

b) 

* s 

Target to Substrate 

... Distance \ B • 

+ 

... B ~ . I x ... + 

a + 

+ B + 

XR oX +. • ... !I + 1 .T B + • 

~.l,x:i • XII s· x X .. +.1 ~ + 

~~ 

-15 -5 5 15 

Location Relative to Target Center (em) 

B 3.5cm 

+ 5em 

• 7cm 

I 10cm 

25 

a) Transmittance Spectra for 902.5 nm-Thick ZnO Coating on Glass; b) Film Resistivity as a 

Function of Distance From the Target Center and Away From the Target 

0 .07 

- ---:-- 0 .06 

E - 0 - ~ -

c 

N 0 .05 "-

< 

0 .04 

o 

)( 

---» target 

4-- + 

race- I!] 

G track 

.. III )C 

.. + Ii 

)( )( 

)C 

5 10 15 

Location Relative to Target Center (cm) 

20 

Source to Substrate 

Distance 

[i1 3.S cm 

+ S.Ocm 

)( 10.0 cm 

Figure 2 AluminumlZinc Ratio (Atomic Percent) as a Function of Distance Away From Target Center 

(Composition Measured by Wavelength Dispersive X-Ray Spectroscopy (WDS» 

119 
Cadmium Telluride 

Unlike CIS thin-film devices, CdTe devices are fabricated in the superstrate configuration with entirely 

wet chemical deposition methods and no vacuum; device structure consists of glass superstrate, transparent conductive oxide, CdS window layer, CdTe absorber, and metal back contact. Using entirely wet 

chemical processes, thin-film CdTe cells have been fabricated with 6-7% (0.08 cm2 active area) efficiencies on Sn02-coated glass. In this section, a brief overview is presented of the potential scale-up issues 

related to the CdS solution-growth and CdT e electrodeposition processes. 

CdS Window Layer - The CdS deposition process for CdTe thin-film cells is similar to the process for CIS 

cells in that CdS films are heterogeneously nucleated onto activated surfaces from an aqueous solution 

containing cadmium sahs and thiourea with ammonium as a complexing agent (reaction throttle). For CdTe 

however, the CdS films are grown on TCO-coated glass substrates with an average thickness between 250 

- 320 nm and at a deposition rate of -180 nmlhour. With a proper selection of the cadmium-to-sulfur ratio 

and ammonium, which acts as a buffer and a complexing agent for cadmium, film properties can be optimized. CdS quality, in terms of surface adhesion, structure, and yield from a batch process, is directly related to the solubility/precipitation of Cd(OHh and the concentration of unassociated [Cd2+] cations. 

CdS films are reproducibly and uniformly deposited onto 1 ft2 substrates from a batch, solution-growth 

process with a >90% process yield; i.e., minimal waste. For optimum n-type carrier density, proper control of the oxygen/sulfur ratio, and surface activation for the subsequent CdTe electrodeposition step, 

CdS films are used in the as-deposited condition and are not post heat treated. A typical transmission 

spectra for CdS on an Sn02-coated soda-lime glass is shown in Figure 3. Although there is a shift in the 

heat treated CdS films to higher wavelengths, the absorption edge for these films is around 520 nm. 

100 

- - Heat Treatment 

.••••• - No Heat Treatment 

80 

c: 

0 60 'iii 

III 

'E 

III 

c: 

~ 40 f-

~ : 0 : 

20 

0 

320 440 560 680 

Wavelength (nm) 

Figure 3 Percent Transmission of Heat Treated and As-Depoisted, Solution Grown CdS 

CdTe Absorber Layer - Optimally-doped CdTe films are produced by electrodepositing a 1.5 -2.0 Ilmthick CdTe layer from an acid bath containing cadmium salts and then heat treating these films at high 

temperature (400°C) in the presence of CdCI2. Although the electrodeposition process was selected 

120 
because of its low-cost and potential scalability to larger areas, a number of processing factors, such as 

cadmium to tellurium ratio in the electrolyte, transport/delivery of cadmium and tellurium to the electrode 

surface, electrolyte contamination, competing cathodic and anodic reactions, and film resistance of the 

as-deposited film, can all have a subtle effect on the final electrical quality of the CdTe films (not to mention what effect these factors will have on formation of the n-p junction). In contrast to CdTe films which 

are deposited by a vacuum, physical vapor deposition process, electrodeposited CdT e films are electrically coupled to the junction partner (i.e., CdS, which is also electrically coupled to a conductive oxide) as 

electrons are transported through the device during the deposition process. This inherent feature of electrodeposition provides an insitu monitor of CdTe film and device quality and can be used as active feedback control in a commercial scale process. 

Through proper control of bath chemistry (pH, cadmium/tellurium ratio, temperature, deposition potential, 

etc.), CdTe films can be reproducibly deposited over small areas and all of these factors should translate 

to larger areas, however; as the deposition area increases, film resistivity becomes more critical. Film 

resistance has a direct effect on the deposition potential which in turn, effects the cadmium to tellurium 

ratio in the film. With increasing film resistance, the equilibrium potential becomes more noble (positive) 

and more tellurium is deposited with respect to cadmium. Through-thickness resistivity in the CdS, i.e., 

from the conductive oxide through the CdS to the CdTe film, is negligible compared to the drop in lateral 

resistivity which can occur in the TCO over large areas. For example, the voltage drop across a typical 

20 ohms/square tin oxide was determined to be -25 mv/cm. Over a 10 cm area then, the shift in deposition potential would be -250 mv. A 250 mv shift would correspond to a change in the CdlTe ratio from 

1.0 to a ratio less than 0.8. Without reducing the resistivity of the TCO below 20 ohms/square, uniform 

films were deposited across 5 cm x 5 cm areas. Although a switch to more conductive TCO films can be 

made (at the expense of transmittance), lateral film resistivity will still be a significant factor as CdTe 

devices are fabricated over larger areas. 

MODULE DEVELOPMENT 

Other than the scaling issues presented above for the cell processing, the remaining issues limiting module and large area array development are the scribing operations which eliminate the hand touch labor of 

conventional series interconnects and the fabrication of reliable interconnects between modules. Results 

from the CIS scribing development efforts are presented in this section. A reliable method for bonding 

interconnects to a molybdenum film has been developed at MMC. 

Scribing of the various coating layers of the photovoltaic cells is crucial in the fabrication of monolithicallyintegrated minimodules. Flow of current occurs from the overlayer transparent conductive oxide (TCO) to 

the Mo back contact. To enable this current flow in an isolated manner requires the fabrication of scribes 

in the Mo back contact layer, in the CdS/CIS multilayers to the Mo back contact, and in the 

ZnO:AVCdSlCIS multi layers to the Mo back contact. Each of these materials exhibit different thermophysical properties (absorptance = f (wavelength (A)) and mechanical properties, which affects their 

removal by scribing processes. 

Candidate scribing processes that are being investigated at Martin Marietta include: 1) laser; 2) 

mechanical; and 3) chemical etching between photolithographically-deposited masks. 

121 

l 
For removal of material by laser incidence. critical material properties include absorptance and emittance 

as a function of wavelength. and melting temperature. For the coatings Mo and CdS/CIS. absorptance 

generally increases as the incident wavelength is reduced. This suggests the use of lower wavelength 

lasers, such as a frequency-doubled (or quadrupled) YAG (N=O.54 11m for frequency-doubled) or excimer 

lasers will more effectively couple with these coatings. resulting in more efficient coating removal. Care 

must be exercised, however, not to couple with the underlying substrate material, which may result in 

substrate damage. 

An example of a scribe in a 1 11m Mo coating on glass produced by a Q-switched YAG laser (A. = 1.06 

11m) is shown in Figure 4(a). In general, the scribe edges appear clean and straight with a small amount 

of Mo cracking or delamination adjacent to the scribe. Excimer processing also resulted in clean scribe 

edges, although more cracking and flaking of the Mo was observed. Scribes produced by a pulsed NdY AG exhibited some glass substrate cracking and larger berms of material adjacent to the scribe than for 

the other laser methods. 

Laser scribing of the combined CdS/CIS multilayers is more difficult. since the desire is to remove the 

CdS/CIS layers without damaging the Mo back contact. Of all the laser scribing methods. Excimer laser 

scribing shows the most promise for selective coating removal. Simple mechanical scribing using a synthetic diamond tool or a stainless steel blade has been found to be more effective for selective removal of 

CdS/CIS from the Mo back contact. Figure 4b shows an example of a scribe in CdS/CIS which shows 

removal of the multi layers. without extensive damage to the Mo back contact. 

Figure 4 a)SEM Micrograph of S ribe in Mo Coating on Glass Produced by Q-Switched YAG Laser 

Showing Relatively Clean Scribe Edges and Minor Cracking of Adjacent Mo; b) SEM 

Micrograph of Scribe in CdS/CIS Coating on Mo-Coated Glass Produced by Mechanical 

Diamond Scribe Showing Selective Removal of CdS/CIS 

122 
aUALIFICA TION TESTING 

In order to survive the demanding 15-20 year lifetimes in space, thin-film cells and modules need to be 

first tested in thermaVvacuum and simulated radiation tolerance tests (Le., solar, electron, proton, etc.). 

One distinct advantage of CdTe and CIS thin-film polycrystalline devices over silicon and gallium 

arsenide is the inherent radiation tolerance of the CdTe and CIS semiconducting layers. Long term radiation testing as well as thermallvacuum cycling of CdTe and CIS cells is in progress. 

Just as important as the environmental stability of these devices, however, is the mechanical stability of 

CdTe and CIS devices and particular, the effect of mechanical forces on the photoresponse of a flexible 

CIS module. Although preliminary data is available in the literature on the testing of rigid CdTe and CIS 

thin-film polycrystalline cells in simulated space environments and actual flight experiments, little is 

known of the effects of mechanical stresses and strains on the electrical behavior of a polycrystalline 

device. An effort is in progress to test the I-V and spectral response of flexible CIS cells before, during, 

and after the application of a cyclic bending force. 

SUMMARY 

An overview of the current efforts at MMC to develop large-scale manufacturing processes for both CIS 

and CdTe thin-film polycrystalline cells and modules is presented with an emphasis on those issues in 

each process that are critical to scalability/manufacturability. Except for the CdS window layer, all films in 

the CIS devices are being deposited by a DC magnetron sputtering system; large-area uniformity was 

also demonstrated for all the processing steps with the cylindrical magnetron sputtering system. CdTe 

cells were fabricated entirely by low-cost, wet chemical methods and small area efficiencies on Sn02-

coated soda-lime glass were on the order of 7%. Scalability issues were identified for the CdTe electrodeposition process. 

Since a major concern in the fabrication of monolithically integrated modules is the scribing operation, 

results from laser scribing the CIS back contact and mechanically scribing the CIS/CdS layers are presented. To evaluate emerging CIS technology, Martin Marietta, in cooperation with the National 

Renewable Energy Laboratory, designed and fabricated a CIS flight experiment with cells based on 

NREL's selenized bilayer process which has demonstrated 16.4% efficiency in AM1 .5 insolation this 

year. 

REFERENCES 

1. J. Hedstrom, H. Ohlens, M. Bodegard, A. Kylner, L. Stolt, D. Hariskos, M. Ruckh, H. Schock, 23rd 

IEEE PVSC, 364 (1993). 

2. M. Conteras, J. Tuttle, D. Du, Y. ai, A. Swartzlander, A., Tennant, and R. Noufi, Appl. Phys. Lett. 

63e(13) (1993). 

3. C. Ferekides, J. Britt, Y. Ma, and L. Killian, 23rd IEEE PVSC, 389 (1993). 

4. L. Chung Yang and A. Rockett, J. Appl. Phys., 75(2) 1185 (1994). 

123 
39?7IP a 

N95-20515 

. /c.)/,; 

AMORPHOUS SILICON THIN FILMS: THE ULTIMATE LIGHTWEIGHT SPACE SOLAR CELL 

G.J. Vendura, Jr., M.A. Kruer, H.H. Schurig, M.A. Bianchi, and J.A. Roth 

TRW Space and Technology 

Redondo Beach, California 

ABSTRACT 

Progress is reported with respect to the development of thin film amorphous (a-Si) terrestrial solar 

cells for space applications. Such devices promise to result in very lightweight, low cost, flexible arrays 

with superior end of life (EOL) performance. Each a-Si cell consists of a tandem arrangement of three very 

thin p-i-n junctions vapor deposited between film electrodes. The thickness of this entire stack is 

approximately 2.0I-lm, resulting in a device of negligible weight, but one that must be mechanically 

supported for handling and fabrication into arrays. The stack is therefore presently deposited onto a large 

area (12 by 13 in.), rigid, glass superstrate, 40 mil thick, and preliminary space qualification testing of 

modules so configured is underway. At the same time, a more advanced version is under development in 

which the thin film stack is transferred from the glass onto a thin (2.0 mil) polymer substrate to create large 

arrays that are truly flexible and significantly lighter than either the glassed a-Si version or present 

conventional crystalline technologies. In this paper the key processes for such effective transfer are 

described. In addition, both glassed (rigid) and unglassed (flexible) a-Si cells are studied when integrated 

with various advanced structures to form lightweight systems. EOL predictions are generated for the case 

of a 1000 W array in a standard, 10 year geosynchronous (GEO) orbit. Specific powers (W/kg), power 

densities (W/m2) and total array costs ($/ft2) are compared. 

INTRODUCTION 

During the next ten years, spacecraft power requirements will grow significantly over the presently 

typical 1 to 4 kW EOL systems. Also, more interest will be focused upon smaller and lighter systems in the 

0.1 to 1.5 kW range. Finally, the proliferation of small, less expensive launch vehicles will require lowmass, low cost, power sources. Current crystalline silicon technology using 8 mil thick devices is too 

heavy, costly and large to support higher power levels on satellites thrust into space by existing and 

planned vehicles. Thin, 13.5% efficient,-silicon cells and even higher efficiency gallium arsenide and 

indium phosphide cells reduce weight and area but increase cost. New generation lightweight 

photovoltaic devices are required to meet this challenge (ref. 1). These new devices, by nature, are 

expected to be both enhancing and enabling: enhancing by offering advantages in power, weight and 

cost compared to traditional crystalline solar cells in existing satellite designs for conventional orbits; 

enabling by extending array and mission capability beyond the present limitations of such space systems. 

For this reason, thin film solar cells are presently generating intense interest within the space 

community. Those technologies that have already enjoyed significant development for terrestrial 

applications are especially attractive. Both a-Si and copper indium diselenide (CIS) fall into this category, 

but of the two, a-Si is by far the more advanced (ref. 2-3). Very large area a-Si cells and integrated 

modules are already routinely manufactured for terrestrial applications with AMO efficiencies of 8 to 10%. 

Although this is considerably less than standard 13.5 and 18.5% Si and GaAs/Ge figures, the material's 

greater radiation resistance, ultra light weight, low cost, flexibility and the ability to be incorporated into 

existing, well-developed, lightweight, satellite array structures makes a-Si not only a viable but also a 

potentially superior alternative. Significantly, the cells can be interconnected in various series and parallel 

configurations by means of standard semiconductor monolithic integration techniques resulting in 

superior packing densities and the reduction in the yield and cost disadvantages associated with 

numerous discrete parts and corresponding handling operations. 

125 

~ PAGi BLANK NOT fiLMED 
1-

a-Si SPACE SOLAR CELL APPROACH 

The a-Si solar cell chosen for such space development, shown in Figure 1, is routinely fabricated 

by Solarex Thin Films and allows the maximum leverage of existing terrestrial technology (ref. 4). The 

active material consists of a stack of three individual u-Si p-i-n cells sandwiched between thin electrodes. 

The upper electrode of Sn02 is transparent to incident light, while the back electrode of silver is opaque. 

The three u-Si cells are not compositionally identical. Instead, uppermost and lowermost devices are 

carbon and germanium alloys respectively to allow for increased collection efficiency by utilizing a broader 

segment of the solar spectrum. The entire stack has a total cross section of only 2.0J..lm. However, the 

commercial product is deposited upon a 40 mil superstrate of either soda lime or borosilicate glass as a 

means of mechanical support during fabrication and handling. Although both single junction and double 

junction variations of this device are manufactured on superstrates as large as 4 square feet, the baseline 

space product considered in this study is limited to 12 by 13 in. 

The effort to develop this terrestrial commercial product into a device suitable for space can be 

divided into two major phases: 

I. Development and qualification testing of these glassed 12 by 13 in. terrestrial cells to 

create a usable, rigid space product. Although the superstrate contributes significantly to overall weight, 

results of early cost and power trades indicate advantages for certain missions. 

II. Development of materials and additional processes for transferring 3 by 3 in. areas from 

the glass superstrate onto a thin, polymer substrate to demonstrate a flexible space product. Earlier 

stages will focus primarily on mechanical issues, while later stages will address both mechanical and 

electrical stability. A later phase will concentrate on the scaling up of these processes to transfer 

12 by 13 in. and larger areas. 

The most important aspects of the Phase I effort center on radiation and temperature effects and 

long term stability. Studies by Woodyard and co-workers indicate that radiation damage, in large part, may 

be reversed by annealing (ref. 5). Also, attention must be focused on the degradation of u-Si output due 

to photons (the Staebler Wronski effect) and its reduction (ref. 6-7). Significant adjustments in the 

manufacturing sequence have already been made by Solarex and will continue to be considered to 

minimize this effect. 

ADVANCED PROCESSES 

Phase II, addressed simultaneously with Phase I to save time, focuses on the materials and 

process development of two key additional processes required for transfer: release and liftoff. These 

processes are illustrated in Figure 2 on the right, while the standard commercial sequence is shown on the 

left. The release process consists of sputter deposition of carefully controlled thin film layers directly onto 

the glass superstrate to partially isolate it mechanically and chemically from the commercial device that is 

deposited subsequently. This limits the adhesive strength which, in turn, facilitates eventual separation. 

The liftoff process, on the other hand, involves the attachment of polymer and other films to the back 

surface of the commercial stack by means of a thermo-compression technique. The number and 

orientation of these backing layers are carefully balanced to create the right relative mismatch in the 

various coefficients of thermal expansion. As a result of differential contraction upon cooling, separation 

at the release interface is accomplished and the solar cell is thereby transferred from the rigid, glass 

superstrate onto the flexible, laminated substrate. 

The release process is dependent upon very specialized, large-scale sputtering equipment. 

Either of two custom built sputtering machines can be used, depending upon solar cell size and quantity. 

Although cells of 12 by 13 in. are presently earmarked for the baseline process, individual device areas are 

expected to eventually increase to 4 and 8 ft2. It is advantageous to load large batches of such cells into a 

single machine for economy. 

126 
I 

I 

I 

\ 

I 

I 

I 

l 

The chamber of the first unit is 6.0 by 6.0 by 6.0 ft. and can be evacuated to the low 10-7 torr 

range by a 16 in. cryopump. The machine can operate in RF or DC modes and is equipped with three 5.0 

by 20.0 in. targets, capable of co-deposition onto three 20.0 by 20.0 in. substrates rotating via a planetary. 

Operation is computer driven and monitored to permit unattended deposition of multiple layers. Control 

devices include an in situ particle counter, a quartz crystal thickness monitor, a residual gas analyzer and 

an optical monitor to track reflective interference to a quarter of a wavelength. 

The second sputtering machine consists of a chamber with a floor area 20 by 12 ft. and a ceiling 

15 ft. high. Overnight evacuation to the low 10-7 torr is achieved by three 16 in. and one 48 in. cryopumps 

coupled to a Woods Root blower. The machine is fitted with three 5 by 40 in. planar cathodes that move in 

a raster pattern from 0.5 to 36 in. away from a substrate as large as 18 ft. long and 12 ft. high. 

Co-deposition is possible via two of the three targets. By means of another cathode assembly, 10 in. 

round, non-planar shapes can be coated. This machine can also operate in either RF or DC modes, is 

similarly computer controlled and monitored, and is fitted with a residual gas analyzer, a quartz crystal 

monitor and a particle counter. 

An earlier version of the overall release process involved the deposition in the smaller machine of 

three separate layers, shown in Figure 3a. After cleaning, the glass superstrate was loaded into the 

chamber which was then evacuated to 10-6 torr. A 400A layer of binder material was deposited. The 

purpose of this film was to promote adhesion between the glass surface and subsequent materials: a 

release layer of approximately 800A followed by a 1.5 ).Jffi cap of Si02. The purpose of the release layer is 

to provide a release interface (R.I.) - a plane of significantly weaker chemical and mechanical adhesion 

compared to all other interfaces - so that separation can eventually be achieved at this surface. 

After deposition of these three layers, the treated glass was shipped to Solarex, where the a.-Si 

solar cell components (Figure 1) were added. The device was then returned. Initial liftoff experiments, 

intended to separate the cell at the R.I., produced mixed results. In some cases the solar cell did not 

release at all; in others, release was uneven. EDAX and SEM investigations of suspect areas of the 

surface seemed to indicate atomic diffusion of the superstrate across the release layer resulting in 

pinning - localized areas of high adhesion - at the R.1. Since the commercial fabrication sequence involves 

Sn02 and contact annealing processes that approach the softening point of glass, a high temperature 

mechanism was suspected. 

To eliminate this pinning without affecting the solar cell manufacturing sequence, the release 

process was modified to include a 400A diffusion barrier as shown in Figure 3b. An additional 800A layer 

varying in composition from barrier to release layer materials was also necessary to ensure the R.1. 

remained the weakest link in the chain of interfaces in order to prevent separation at the barrier-release 

layer surface instead. 

Again the treated glass was shipped to Solarex and returned with solar cells attached. 

Experiments demonstrated significantly improved results, although the process continues to be 

developed further. 

The liftoff process involves the attachment of five plies of flexible material to the back surface of 

the cell in three stages. As shown in Figure 4, three layers of 1 .0 mil polymer are interspersed with two 

1.0 mil layers of fiberglass cloth. In the first stage, all but one polymer layer are aligned, placed in vacuum, 

degassed and subjected to a two step cure process. The second stage consists of surface preparation of 

the Ag contact on the back of the solar cell, followed by mechanical placement of the remaining polymer 

film. In the third stage, all parts are joined into a single unit by an additional vacuum, degas, and cure 

sequence. Under ideal conditions, upon cooling, the solar cell releases spontaneously and cleanly at the 

R.1. due to a differential in the coefficients of thermal expansion (CTE). 

The CTEs are balanced by careful selection of layer composition, thicknesses and orientations. 

For example, the two fiberglass cloth plies are aligned in different directions, one at 0,90 degrees and the 

other at ±45 degrees, as implied by the dissimilar slash patterns in Figure 4. Another key concern is the 

complete elimination of air bubbles during processing. Air bubbles result in voids - pOints of no adhesion 

between the flexible substrate and the solar cell. Thus, upon release of the bulk of the a.-Si, areas under 

127 
the voids can remain behind, creating pinholes in the surface of the cell which in turn may result in shorting 

and power degradation. 

Figure 4 is representative of one of several variations of the liftoff process still under 

development. Other variations use different quantities of layers or plies of different thickness. The 

objective, however, is to eliminate layers or to use thinner plies so that the total flexible substrate is 

approximately 2.0 mil. 

LIGHTWEIGHT STRUCTURES 

a-Si cells, both glassed (Phase I) and flexible (Phase II), are suitable for incorporation into 

conventional and low mass arrays. For the purposes of comparison, a 1000 W array was considered. In 

the first case a state-of-the-art 0.5 in. thick AI honeycomb with 5.0 mil graphite face sheets and a single 

layer of 2.0 mil Kapton to insulate the solar cells is assumed. In addition, two existing, well-developed 

lightweight structures were studied. The first is an adaptation of the Advanced Photovoltaic Solar Array 

(APSA) (ref. 8). The original APSA consisted of a 5.4 kW, 15.25 by 2.81 m, mast-deployed, 42 panel, 

prototype wing as shown in Figure 5 (ref. 9). This unit was populated by 2.2 mil thick crystalline silicon 

solar cells, 2.0 by 4.0 in. area, with 2.0 mil cover glasses. The efficiency of these cells was 13.5%. A key 

lightweight feature is the employment of a 2.0 mil carbon loaded Kapton substrate, accordion folded for 

stowage during launch. Despite this lightweight blanket, however, major contributions to mass resulted 

from the deployment mast, the frame and the stowage container. Since this study concerns an array less 

than 25% of the original APSA area, wherever possible, features such as this container size were scaled 

down accordingly. 

Another advanced lightweight structure involved using a TRW developed and tested framed 

membrane technology. Main features of such a system, highlighted in Figure 6, include a rigid membrane 

solar cell support consisting of a very thin laminate with a foam core and high modulus, graphite fiber 

reinforced plastic (GFRP) face sheets. Kapton is used to insulate the solar cells from the membrane 

surface. The frame tubes are transfer molded from a mixture of high modulus and high strength GFRP 

materials. To create a panel structure subassembly, the various GFRP components are joined together 

through a precision bonding process without the need for mechanical fasteners. Such a panel design is 

adaptable for use with cells of various types, sizes and thicknesses and can be readily scaled up or down 

as required. An advanced version of the system involves lighter frame and substrate elements. Indeed, 

an adaptation of the system, using a different rigid laminate and no frame whatsoever, was incorporated in 

the Earth Observing System (EOS) program. 

The ultimate lightweight array, the Ultra Light Film Array (ULFA), is presently limited to satellites 

$1000W. It includes a 2.0 mil flexible Kapton blanket, but not the relatively heavy components of either 

the APSA or framed membrane designs. In this case, the blanket is deployed and supported by 

lightweight strain energy hinges. Although development of such a structure is not as mature as APSA 

and membrane technologies, it is nonetheless included in this study for comparison purposes. 

RESULTS 

Results are in the form of EOl performance predictions of satellite systems incorporating various 

a-Si and conventional crystalline solar cells populating the four structures described. In all cases, a 

standard, 10 year GEO mission is assumed for the nominal 1 000 W array. Weight of stowage and 

deployment hardware is included. In the case of a-Si arrays, cell interspacing was set at 120 mil, while for 

crystalline devices it was 30 mils. Also, whenever possible, proven cell and system design factors were 

used. For example, empirical loss factors were applied to account not only for temperature and radiation 

degradation, but also for more obscure losses such as Staebler Wronski (SW), packing, wiring, installation, 

cycling, cover glass darkening, etc. All comparisons are thus at a system level of performance as opposed 

to often quoted device or cell level performance. It is noted, however, that all systems are not universally 

applicable to all cells, and in certain specific cases some overdesign and underdesign is inevitable. 

Therefore, the accuracy of results is estimated to be ± 10%. 

128 
Data are summarized in Table 1 for 13 different systems. Rows 1 through 4 present data for 

unannealed a-Si having 4 different cover glass thicknesses: 40 mils, 8 mils 2Jlm and 1 mil. Systems 5 

through 7 outline crystalline Si, while 8 and 9 highlight crystalline GaAs/Ge. Rows 10 through 13 examine 

the same a-Si systems as 1 through 4, but this time the systems are designed for self annealing, resulting 

in considerable radiation and Staebler Wronski loss recovery. Details such as cell type, size, device and 

cover thicknesses and BOL efficiency, 11 , are presented in the leftmost columns. For a-Si, BOL 11 was 

assumed to be a conservative 10.0% at AMO and 28°C. Crystalline 11, on the other hand, varied from 12.2 

to 18.2% as tabulated. Staebler Wronski degradation is assumed at 15% for unannealed a-Si and 5% for 

the same cells when annealed. In the table, honeycomb, membrane, APSA and ULFA data then follow in 

terms of three key parameters: specific power (W/kg). power density (W/m2) and areal density (lbs/ft2). 

In generating these data for the systems involving a-Si, radiation degradation behavior in 

response to orbital environment was calculated from a model using the standard approach of equating 

ionization and displacement damage with P/Po power reduction. This technique uses existing data for 

P/Po from 1 MeV proton f1uences and extends it to other proton energies. P/Po is defined as a function of 

1 MeV protons similar to crystalline technology, using 1 MeV electrons as the conversion parameter (ref. 

10). 

The a-Si comparison is presented graphically in Figure 7 in which the structural density (the sum 

of the system's areal density and the weight of peripheral hardware -hinges, booms, deployment 

hardware, etc.- spread over array area) is plotted as a function of specific power. As shown, even cells with 

40 mil covers generate respectable powers when compared with the -15 W/kg figure for a crystalline 

silicon system (not shown) using less than the state-of-the art honeycomb presented in this study. As 

expected, a-Si with 8 mil covers performs considerably better, especially in the case of APSA in which 

best results are 77.7 and 83.2 W/kg for unannealed and annealed cells respectively. Note that because 

of weight, only 2 ~ and 1 mil a-Si cells are appropriate for application to the ULFA structure. Here, results 

as high as 340.9 W/kg are indicated for the annealed 2 ~ cover system. Of note is the fact that 1 mil of 

cover glass and/or annealing makes a considerable difference over an unannealed 2 ~ a-Si array 

incorporated in any structure in the GEO environment. 

In Figure 8, less-than-optimal 8 mil covered a-Si is compared to the best of the crystalline Si and 

crystalline GaAs/Ge systems. The crystalline Si cell used was a 2.5 x 5.0 cm, 2.7 mil thick device with both 

a back surface field and reflector (BSFR) and a 2.0 mil cover. The GaAs/Ge device was 4.0 x 4.4 cm, 

5.5 mils thick, with a 3.0 mil cover. The curves demonstrate that both unannealed and annealed a-Si is 

superior at structural densities approaching APSA. At higher structural densities, however, a-Si and 

crystalline Si are comparable, while GaAs/Ge is superior. 

In Figure 9, the same crystalline systems are compared with those for the a-Si cell covered with 

1 mil of glass. Here the ULFA structure is inappropriate for all but the a-Si case. Best results are 266.5 and 

288.3 W/kg for unannealed and annealed devices respectively. 

Specific power alone, of course, is not the only major point of comparison. Depending on mission 

and program constraints, power density, and areal density can also be key considerations. Table 1 also 

offers these data for the 13 systems under study. 

Another essential factor is cost. System cost, typically in $/W, involves the sum of three separate 

figures: cell materials, structural materials and recurring fabrication labor. Only the first of these is 

presented in Table 1. The prices of the crystalline cells are well-established, while, admittedly, a-Si figures 

are rough estimates based upon current commercial terrestrial production prices. Depending on cover 

glass, these figures may be off by a factor as high as 5. Nevertheless, savings associated with a-Si 

compared to crystalline systems, depending on choice, readily approach an order of magnitude. 

129 

J 
SUMMARY 

An approach is described for developing very lightweight a-Si solar cells for space by leveraging 

progress of terrestrial devices. Glass and flexible versions are being addressed simultaneously. Critical 

release and liftoff processes for transferring such cells from a rigid 40 mil superstrate onto a flexible 2.0 mil 

substrate are presented. EOL performance predictions are generated based upon a 10 year GEO 

mission of a 1000W array incorporating different a-Si and crystalline cell configurations with four distinct 

structures. Results demonstrate that specific powers of 266.5 and 288.3 W/kg are achieved when a-Si 

cells with 1 mil covers, unannealed and annealed respectively, are combined with the ULFA structure. 

The specific power increases further, to 340.9 W/kg for annealed devices with a 2 flIT1 cover. These 

figures are two to five times better than the performance of conventional crystalline solar cells in similar 

systems for an identical mission. 

REFERENCES 

1) Vendura, Jr., G.J.; Malone, P.; and Crawford, L.: A Novel Lightweight Solar Array: Comparison 

with Conventional Systems. IEEE 23rd Photovoltaic Specialists Conference, Louisville, KY, 

May 1993, pp. 1381 - 1385. 

2) Arya, R.R.; Yang, L.; Bennett, M.; Newton, J.; Li, Y.M. ; Fieselmann, B.; Chen, L.F.; Rajan, K.; 

Wood, G.; Poplawski, C.; and Wilczynski, A.: Status, Progress and Challenges in High 

Performance Stable Amorphous Silicon Array Based Triple Junction Modules. IEEE 23rd 

Photovoltaic Specialists Conference, Louisville, KY, May 1993, pp. 790 - 794. 

3) Basol, B.M.; Kapur, V.K.; Halani, A.; Minnick, A.; and Leidholm, C.: Modules and Flexible Cells of 

CulnSe2. IEEE 23rd Photovoltaic Specialists Conference, Louisville, KY, May 1993, 

pp. 426 - 430 

4) Carlson, D.E.: Market, Manufacturing and Technical Progress in Amorphous Silicon Photovoltaics 

in the U.S. IEEE 22nd Photovoltaic Specialists Conference, Las Vegas, NV, Oct. 1991 , 

pp. 1207 - 1212 

5) Woodyard, J.R.; Landis, G.A.: Radiation Resistance of Thin-Film Solar Cells for Space 

Photovoltaic Power Solar Cells, 31 , 1991 , pp. 297 - 329 

6) Staebler, D.L.; and Wronski, C.R.: Appl. Phys. Lett. (1977) pp. 292 

7) Yamagishi, H.; Asaoka, K.; Nevin, W.A.; Yamaguchi, M. and Tawada, Y.: Light Induced Changes 

of Amorphous Silicon Solar Cells by Long Term Light Exposure. IEEE 22nd Photovoltaic 

Specialists Conference, Las Vegas, NV, Oct. 1991, pp. 1342 - 1346 

8) Kurland, R.M.; and Stella, P.M.: Demonstration of the Advanced Photovoltaic Solar Array. 

European Space Power Conference, Florence, Italy, Sept. 1991, pp. 675 - 680 

9) Stella, P.M.; and Kurland, R.M.: Operational Considerations of the Advanced Photovoltaic Solar 

Array. 27th Intersociety Energy Conversion Engineering Conference, San Diego, CA, 1992, 

pp. 6.29 - 6.34, 

10) Tada, H.Y.; Carter, Jr., J.R.; Anspaugh, B.E.; and Downing, R.E.: Solar Cell Radiation Handbook, 

Third Edition, National Aeronautics and Space Administration, Jet Propulsion Laboratory, 

Nov. 1982. 

l30 

I 

____ J 
No. 

I 

2 

3 

4 

5 

6 

7 

8 

9 

10 

11 

12 

13 

TABLE 1: PERFORMANCE COMPARISON: CELLS, STRUCTURES & SYSTEMS 

EOL, 10 Year, GeO Rig id Honeycomb 

0." 33Ib1ftJ -

CeU C e:W Ce:W Covef eOL ~ SW CeW Cost S po«. P . P. Of!nslt ""ui O. 

Ty pe: Annul SI ze Thick. hlckne s ('M Loss (SIWI IWlkgl IW/m21 IIbsll121 

,l ·5. N 12. 13 In. 2/l 40 mils 100 0.85 2 50 19.0 90.1 0.972 

('f-Si N 12x6.S in. 2/l 8 mils 10.0 085 8 50 35.3 97.2 0.565 

Cf-Si N 12x 13 in. 2/l 2/l 10.0 0.85 1500 3.9 88 0.465 

n-Si N 12xlJm. 2/l 1 mil 10.0 0.85 20.00 42.7 99.2 0.476 

S; BSR N 2 .SJC5.0 cm 8.0 mils 2 mils 12.2 1.00 24700 33.3 89.6 0.551 

5i BSFR N 2.5. 5.0 em 80mlls 2 mils 146 1 00 26900 32.3 87.0 0.551 

5i SSFR N 2.5x5 0 em 2.7 mils 2 mils 13 5 1.00 52600 37.7 92.7 0503 

GoAs N .. 0.4.4 em 5.S mils J m lls 18 2 1 00 135000 45.0 131 .2 0.597 

GaAs N 4 .Ox4.4 em 8 a mils "mils 18 2 100 90900 41.2 13".0 0.666 

fe-51 Y 12.13 In, 2/l 40 m ils 10.0 095 2 50 20.3 965 0.972 

(l. Si Y 12x6.5 in. 2/l aml/s 100 0.95 850 37.8 104 ,2 0.565 

n·Si Y 12x1 3 In . 2/l 2/l 100 0.95 1500 46.9 106.3 0.465 

11-Si V 12x l 31n 2" 1 mil 100 0 95 2000 "'5.7 1063 0.476 

• • cell OR module U • A real D en sl1y + Peripheral De nSity 

Glass 40 mIl 

Sn02 10.000,1. 

~ 

< i , sooA 

'ri 

p, - 200,1. T 

i n, - 200,1. 

T P2 - 200,1. 

a-SiC 

a-Si i2 4.000,1. 

~ 

i "" - 200,1. 

T P,- 200.1. 

i, 2.000,1. 

ZnO 800,1. 

i 11:3 - 200.1. T 

a-SiGe 

Ag 1.200,1. 

Figure 1: Cross Section of Triple Junction 

Solar Cell 

131 

FrAmed Membu lne 

0.300 Ibttt' -

Spoc. P. • Dens" AsuiD. Spec. P. 

[Wlkgl IW/m21 l'bsll121 IWlkgl 

21 .8 88.2 0.828 27.8 

45.4 95.2 0.429 71.7 

5.4 8.7 0.331 11.6 

58.2 97.2 0.342 121.2 

42.3 88.3 0.428 59.7 

4 1.1 85.8 0.428 57.6 

49,3 91.3 0.379 73.3 

56.3 127." 0.463 71.0 

SO.4 127.9 0520 65.9 

23.4 94.5 0.828 29.8 

48.7 102.0 0.429 83.2 

64.4 104.1 0.331 139.5 

62.3 10. 1 03"'2 129 8 

Glass 

Sn02 

a-SiC 

a-Si 

a-SiGe 

ZnO 

Ag 

1 

(a) 

Commerical Processes 

(RIGID Glass Superstrale) 

APSA 

0.'331"""'-

P . Dens,lt 

(W/m2 ) 

81.7 

88.2 

8.0 

90.0 

81.0 

78.2 

81.9 

117.7 

119.1 

87.5 

94.5 

96.4 

96.4 

, 

ULFA 

0.032Ib1ft' -

A'taiD. Spoc. P. P. Dens'" ArnID. 

IIbsll121 IWlkgl fW/m 2) IlbsJf1 21 

0.602 

nI_ nI_ nI_ 

0.233 "'- nI_ nJ_ 

0.142 28.4 8.8 0.063 

0.1 52 266.5 97.2 0.075 

0.278 nJ_ nJ_ "'- 0 278 "'- "'- nJ_ 

0.229 "'- nJa 

0.313 nJ_ nI_ nI_ "'- 0.370 "'- "'- nI_ 

0.602 "'- nJa nJa 

0.233 "'- nJa nJa 

0." 2 340.9 105.2 0.063 

0.152 283.3 1052 0075 

I RELEASE I 

UFTOFF 

(b) 

AOditionai Processes 

(FLEXIBLE Polymer Substrate) 

Figure 2: Total a-Si Processes 
CommeriC.al Glass CommeriC.al GlaSS 

t t Adhesiof1400A Adhesion 400A 

SatTier 400A 

Release 800A Graoed 800A 

Ag Contact 0' Solar Cell 

R.I. _ 

Cap 1.5" ReleaSe BOO}. 

_R.t. 

}- Stage 2 

~--------------------~ tage3 

CommoricaJ n·$i Solar Ce. Cap 1.5" 

com:J n·Si Solar Gel 

la) EaI1y Ib) Advanced 

Figure 3: Release Process (a) Early 

(b) Advanced 

Figure 4: Liftoff Process 

I 

2.76m 

(10B.6 Inl 

BLANKET 

2.Bl m WIDTH 

(110.6 inlll D 

AND PALLET 

LENGTH 

I 

t 

LID 

>__--------------- 15.25 m (600.6 inl FOR 39-CELL-COVERED __________ _ 

PANEL BLANKET PLUS 3 LEADER PANELS 

FLEXIBLE PRINTED ............ 

CIRCUIT HARNESS 

12 PLACES) 

Figure 5: Full Scale APSA Structure 

132 

MAST 

I 

_I 
. ~ 

350 

300 

250 

0; 

<:;. 

~ 200 

l; 

~ 

0 

Q. 

~ 150 

a. 

III 

100 

50 

T 

38mm 

(1 .50 in) 

25mm 19mm 

\",,(1 .0 in)>-\ (0.75 in) 

-~ ~r:z:.J 

M6OJ/TSOOH GFRP RESIN 

TRANSFER MOLDED FRAME 

PRE5S-CURED 

GFRP 

MEMBRANE 

90 IJIT1 (3.5 mil) 

XN70 GFRP 

FACESHEET 

(01±601 

~::;:;:;:#~ :{ 

:J:-O.9mm I=====~ _ (0.035 inl 

ROHACEL 

SEcnONA-A FOAM CORE 

Figure 6: Framed Membrane Structure 

". j' ,I ! ,I 11 11- Unannealed I . -. Annealed I 

~T -i--,I --! -1 i-- ~ ,....-----,---,-1---. . l-- I 

'. •• , I I I ,1 f---( 

r- I ..... \, ~ I 1_ I I ~ I ~""""""\'I'" ; I i ~ , I, @'+-L- , -l 

I ... , I I' ~ ~ f....';' I '.:. ! ~ , ~ I 

, I \~ . .L i ~ I I ~-I---I 

0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 

Strue1Ufal Density (Ib/sq tt) 

Figure 7: Comparison - a-Si VS. Structure 

133 

-----.. - ---
90 

I ~ ' I I ,, ~ 

I " ~ ~ 

--~: : II , u 80 - ' ---

I I' : ~ . 0 

~ 

l_~' . <I I -

70 ~ 

0; 

~ ~ ' ~~ ~ ...J I ~ ''\ ::J 

<> 

~ 

a 60 i Cl. 

(J 

~ I <> a. 

'" ___ 1- 50 I ------

0- a-Si (8 mil unannealed) 

40 ~ a-Si (8 mil annealed) 

GaAs Ge (5.5 mil) 

Si BSFR (2.7 mil) 

~ ~~~~~~~~~~~~~~~~~~~LW~~ 

o 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 

5truc1UTSI Density (IJ/sq ttl 

Figure 8: Comparison - Conservative a-Si with Best Crystalline vs_ Structure 

300 

280 

260 

240 

220 

0; 200 

~ 

~ 180 

<> 

~ 

a 180 Cl. 

,g 140 '13 

<> 

a. 

'" 120 

100 

80 

60 

40 

20 

0 

~_~~ ____ , w , I 1 z __ _ r---

I ~ , I ::;: ' I III 8 .---'-----"--::;: t -.-- ! ~ I z " ----fil -----r~ Cl. I I ::;: ; , 0 ._" _______ ~ ! I ~ 

i I IL 1 a: i 

- '--j 

f----.------. ---~-r---

0- a-Si (1 mil unannealed) 

f-------,-----I ~ a-Si (1 mil annealed) 

- GaAs/ Ge (5.5 mil) ----'-Si BSFR (2.7 mil) 

____ .__ _ __ . - __ ' __ I __ __ , ! __ · ___ J __ I ____ --- . . -L--r-----

0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5 

5 IruC1,,",' Denslly (1l>/sq ttl 

Figure 9: Comparison - Best Cases vs. Structure 

134 

----------
/9951 / J//OO 3 9?7~4/ 

N95-20516 

MODELING OF HIGH EFFICIENCY SOLAR CELLS UNDER LASER PULSE FOR 

POWER BEAMING APPLICATIONS 

Raj K. Jain 1 

NASA Lewis Research Center 

Cleveland, Ohio 

and 

Geoffrey A. Landis2 

NYMA, Inc. 

Brook Park, Ohio 

SUMMARY 

Solar cells may be used as receivers for laser power beaming. To understand the behavior 

of solar cells when illuminated by a pulsed laser, the time response of gallium arsenide and 

silicon solar cells to pulsed monochromatic input has been modeled using a finite element solar 

cell model. 

INTRODUCTION 

Solar cells have been used to convert sunlight to electrical energy for many years and also 

offer great potential for non-solar energy conversion applications. Their greatly improved 

performance under monochromatic light compared to sunlight, makes them suitable as 

photovoltaic (PV) receivers in laser power beaming applications. Laser beamed power to a PV 

array receiver could provide power to satellites, an orbital transfer vehicle, or a lunar base 

(ref. 1). Gallium arsenide (GaAs) and indium phosphide (InP) solar cells have calculated 

efficiencies of more than 50% under continuous illumination at the optimum wavelength 

(ref. 2). Currently high power free-electron lasers are being developed which operate in pulsed 

conditions. Understanding cell behavior under a laser pulse is important in the selection of the 

solar cell material and the laser. 

An experiment by NASA Lewis and JPL at the A VLIS laser facility in Livermore, CA 

presented experimental data on cell performance under pulsed laser illumination (refs. 3 and 

4). Reference 5 contains an overview of technical issues concerning the use of solar cells 

for laser power conversion, written before the experiments were performed. As the 

experimental results showed, the actual effects of pulsed operation are more complicated. 

Reference 6 discusses simulations of the output of GaAs concentrator solar cells under 

pulsed laser illumination. The present paper continues this work, and compares the output 

of Si and GaAs solar cells. 

1 Work funded by the National Research Council - NASA Research Associateship Programs 

and a NASA Research Grant (NAG3-1466) at the University of Toledo. 

2NYMA Inc. under a NASA Contract. 

135 
CELL SIMULATION RESULTS 

Figure 1 shows the cell designs and the laser pulse simulated. For simplicity, the laser 

pulse was assumed to be square. Most of the results have been calculated for a peak 

intensity of 50 W/cm2

, which corresponds to nearly 1000 suns concentration. The PC-1D 

computer code, a finite-element simulation of carrier transport in semiconductor devices 

(ref. 7), was used to analyze the cell current during and after the pulse for various 

conditions. 

The GaAs solar cell simulated was a concentrator p+n cell with an efficiency of 27.5% 

under AM1.5D, 1000 sun illumination. The current output was observed to be linear with 

laser intensity from 500 mW/cm2 to 500 W/cm2 • Figure 2 shows the cell short circuit 

current during and after the laser pulse, for a laser at three different wavelengths. The 511 

run wavelength corresponds to the available copper-vapor laser, 840 run corresponds to the 

optimum wavelength for GaAs and also proposed operating wavelength in NASA SELENE 

project, and 870 run is near the band edge of GaAs. The decay of the current can clearly 

be seent to have two distinct components: an initial rapid decay immediately following the 

laser pulse, followed by an exponential decay with much longer time constant. The amount 

of initial decay is greatest for the light with the strongest (hence, shallowest) absorption, 

511 run, and is least for the weakly absorbed light at 870 run. Figure 3 shows this initial 

decay on a shorter time scale. Here the parameter varied is the operating voltage of the 

cell. Further results of this simulation can be found in reference 5. 

Compared to GaAs, silicon solar cells have much longer minority carrier lifetimes and 

much weaker optical absorption, resulting in deeper absorption of the light and longer 

characteristic time constants. A typical silicon solar cell was modeled, with a diffused (erfc 

profile) n type junction. The efficiency is 17.2% under AMO (space) illumination, slightly 

better than cells used in space today, but well below the best efficiencies observed in the 

laboratory. Efficiency increases to 31.8% for monochromatic light at 900 run at an 

intensity of 50 W/cm2

. 

Figures 4 and 5 show the decay of short circuit current of the silicon cell compared 

with that of the GaAs cell. As expected, the silicon cell shows considerably slower 

response. 

As in the GaAs cell, the decay has a rapid initial decay followed by a slower 

exponential decay. Figure 6 shows the fit of an exponential to the portion of the decay 

between 150 and 250 nS after the pulse. The characteristic time constant for this portion of 

the decay is 360 nS, which is intermediate between the base lifetime of 20 f-LS and the 

emitter surface lifetime of 11 nS. 

Figure 7 shows the short circuit current at different wavelengths. Note that the currents 

have been normalized; the absolute response is best at 900 run (peak of 31 A). The 

response at 1.06 f-L is poor (peak 3.6 A). As with the GaAs cells, the most weakly absorbed 

light has the least rapid initial fall-off, and the most strongly absorbed light the most rapid 

initial fall-off. The response drops by a factor of e over a time scale on the order of 25 nS. 

136 
F or such a cell, then, we can expect that the silicon cell will tend to integrate the pulsed 

input into nearly CW output only if the time between pulses is short compared to 25 nS. 

The capacitance of a Si cell at zero bias is typically about 100 nF. The series 

resistance of this 1 cm2 cell was taken to be 4 mO. The RC time constant for the charge to 

be removed from the cell under short circuit is thus expected to be about 0.4 nS. This is 

much shorter than the time scale of the current decay. In actual operation, however, the 

cell would be connected to an external circuit with associated resistance, inductance, 

capacitance, and a battery-supplied bias voltage. This external circuit will considerably 

complicate the output characteristic (refs. 3 and 4). 

REFERENCES 

1. Landis, G.A.: Space Power by Ground-Based Laser Illumination. IEEE Aerospace and 

Electronics Systems Magazine, vol. 6, 1991, pp. 3-7. 

2. Jain, RK.: Calculated Performance ofIndium Phosphide Solar Cells Under Monochromatic 

Illumination. IEEE Trans. on Electron Devices, vol. 40, 1993, pp. 1893-1895. 

3. Anspaugh, B.; Mueller, R; Lowe, R ; and Landis, G.: Results of Illuminating Various Solar 

Cells with Pulsed Laser Beams. JPL Publication 92-25, November 1, 1992. 

4. Lowe, R.; Landis, G.A.; and Jenkins, P.: The Efficiency of Photovoltaic Cells Exposed to 

Pulsed Laser Light. 12th Space Photovoltaic Research and Technology Conference (SPRAT 

XII), NASA CP 3210, Oct. 1992, pp. 129-146. 

5. Landis, G.A.: Photovoltaic Receivers for Laser Beamed Power in Space. Journal of 

Propulsion and Power, vol. 9, 1993, pp. 105-112; IEEE 22nd Photovoltaic Specialists 

Conference, Oct. 1991 , vol. II, pp. 1494-1502. 

6. Landis, G.A.; and Jain, RK.: Approaches to Solar Cell Design for Pulsed Laser Power 

Receivers. 12th Space Photovoltaic Research and Technology Conference (SPRAT XII), 

NASA CP 3210, Oct. 1992, pp. 155-166; 1st Annual Wireless Power Transmission 

Conference, Feb. 1993, pp. 551-564. 

7. Basore, P.A.: PC-l D Version 3: Improved Speed and Convergence. IEEE 20th Photovoltaic 

Specialists Conference, Sept. 1988, pp. 462-468. 

l37 
L 

Peak: power = 

p 

f\ 

j+-116 IJ.S~I 3060 x average power 

1\ (1----11- · 

38n~ IT 

Wavelength = 511 run 

___ Peak of 50 W Icm2 

.--___ --, __ (except as noted) 

P 

- , .... ----25nS 

Figure IA Pulse format of copper-vapor laser (top) and pulse used in computer model (bottom) 

t 

n GaAs base 

doping 5 EI7 

lifetime 9.2 nS 

diffusion length 2.2 Il 

p GaAs emitter 

doping 2.5 EI8 

lifetime 0.85 nS 

diffusion length 1.75 Il 

junction 

depth 

0.7 Il 

pulsed monochromatic light 

25 nS square pulse 

Figure lB. GaAs solar cell model used for computer simulations. 

Diffused n+ emitter 

ERFC proftle 

surface doping n type 2 EI9 

surface lifetime I I nS 

Two layer 

pSi base 

doping 9 EI6 

lifetime 20 IlS 

pulsed monochromatic light 

25 nS square pulse 

FrontSRV 

10000 em/see 

+ 

t 

junction 

depth 

0.9 Il 

~~·T 

bm .. mmmmEm .. mm"Emmg~"Em .. mm"Em~~ Back Surface Field 3 11m p+ 1 E18 

Figure Ie: Silicon solar cell model used for computer simulations. 

138 
10000 

1000 

-< 100 E 

C 

.. 10 

~ U 1 

·S 

... . 1 .!:: 

U 

;: .01 0 

.<: 

<Il 

.001 

.0001 r 

0 

. 

10 20 30 

Time (nS) 

40 50 

70 om 40 nm 

Figure 2. Short circuit current of GaAs cell during laser pulse with incident 

wavelength as parameter. 

..,. -1.0 

~ > 

-" :'l E -2.0 

o 

~ 

R 

E -3.0 

'" .= 

'" g 

o -4.0 

Bias (V): 

' .. 1.22 

'~-...... - '-.. - .... ....... 1.15 

~-";";,:;; .... _. ;..;.;.;,;; ..""' .. ..... """,';:;: ' d 1.0 

-5.0 '--~---'-_~--'-~_-'--~---L_~-' 

0.0 0.2 0.4 0.6 0.8 1.0 

Time (nS) 

Figure 3. Current decay in first nanosecond of pulse as a function of bias (linear scale; 

current measured a decrease from illuminated value.) Wavelength 840 nm. 

~ 100 1 

'E 

.. 'u 

.. 

o 

.<: 

.1 

VJ .0 1 

50 100 150 200 250 

lime (nS) 

Figure 4: Comparison of Si and GaAs short-circuit current response to 25 nS pulse 

139 
5.0 II GaAs cell (840 nm pulse) 

4.0 

3.0 / 

/ \ 

2.0 I \ Si cell (1.06 micron pulse) 

f \,./ J 

i 

1.0 j 

i 

o 

'--

"'------_._-_ .. __ ._ .._ .. -

50 100 

time (oS) 

.... _ .... _ .... _ ..... ..., 

150 200 

Figure 5: Comparison of Si and GaAs short-circuit current response to 25 nS pulse 

1000 -·-·-7·--.-.-. ....... ........ ........ 

EXP [-t/(360ns)] 

Base J~etime 20 ).lS 

100 200 Emitter lifetime at surface 11 nS 

time (oS) 

Figure 6: Exponential fit to decay of short circuit current for Si cell 

1.00 

0.90 

0.80 

0.70 

0.60 

0.50 

0.40 

0.30 

0.20 

-. ' " -.. .. . 

. . ... . I O.OO "'---~----'--~--L...--~--'--~-----' 

0. 10 -:--~~~:-.'-.. - .......... - .. -........... . 

o 50 100 150 200 

time (oS) 

Figure 7: Decay of normalized short circuit current for Si cell at several wavelengths 

140 
3 9??~7 

N95-20517 

CALCULATION OF NEAR OPTIMUM DESIGN OF InP/lno.53Gao.47As MONOLITHIC 

TANDEM SOLAR CELLS 

P. Renaud, M.F. Vilela, A. Freundlich,1 N. Medelci, and A. Bensaoula 

University of Houston 

Houston, Texas 

SUMMARY 

An analysis of InP/lnGaAs tandem solar cell structures has been undertaken to allow for 

maximum AMO conversion efficiencies (space applications) while still taking into account both 

the theoretical and technological limitations. The dependence of intrinsic and extrinsic 

parameters such as diffusion lengths and generation-recombination (GR) lifetimes on NIP and 

PIN devices performances are clearly demonstrated. We also, report for the first time the 

improvement attainable through the use of a new pattemed tunnel junction as the inter cell 

ohmic interconnect. Such a design minimizes the light absorption in the interconnect region and 

leads to a noticeable increase in the cell efficiency. Our computations predict 27% AMO 

efficiency for NIP tandems with ideality factor y=2 (GR lifetimes:::: 1 jlS), and 36% for y=1 (GR 

lifetimes :::::100jlS). The method of optimization and the values of the physical and optical 

parameters are discussed. 

INTRODUCTION 

The use of monolithic InPllno 53GaO 47As tandem solar cells for space applications is still at the 

research stage. It has already been demonstrated that their band gap values are particularly 

adapted for maximum AMO performance levels (ref. 1) in addition to their particularly high 

resistance to radiation damage (ref. 2). Using chemical beam epitaxy (CBE) and new device 

processing techniques, we have already shown in our laboratory that high quality materials and 

devices can be obtained (ref. 3). Therefore, a detailed theoretical simulation is required in order 

to further optimize and refine the structures now possible with this new technology so that 

photovoltaic (PV) devices with maximum efficiencies can be achieved. 

Regarding both top and bottom cell emitter and base thicknesses and doping levels, we calculate 

the conditions for maximum photocurrent matching. The dark current is evaluated from the 

intrinsic limitation where the saturation current is dominant (giving ideality factor y=1), to the 

extrinsic limitation where GR current due to trap levels in the depleted region dominates (ideality 

factor y=2) . The trap level density, which is highly correlated to the epilayer quality, yield 

effective GR lifetimes in the microsecond range (ref. 4). Thus, their participation in the 

recombination processes of photo-excited carriers is negligible compared to band to band 

recombinations and does not affect minority carrier diffusion lengths. This is why most of actual 

photovoltaic devices show performances limited by the open circuit voltage and fill factor while 

still exhibiting an excellent photocurrent characteristic (ref. 5). 

The present study is undertaken to set a near optimum design for a maximum internal efficiency. 

The effect of grid shadowing and light reflection which are directly relevant on technological 

processing are not taken into account. 

IAlso at the LPSES-CNRS, Valbonne, France. 

141 
METHOD OF CALCULATION 

The values of the radiative recombination coefficient and minority carriers mobility versus doping 

levels in high quality InP and GalnAs epilayers were extracted from references 6 and 7. 

The expressions of the equations describing photovoltaic multi-junctions are well known (ref. 8) 

and will not be recalled here. Nevertheless, it is important to summarize those describing the 

bottom cell response and the dark current. 

Let us write Jt and Jb as the top and bottom cell photocurrent. Jb is a single InGaAs solar cell 

response J10GaAs (E) reduced by the absorption of the top cell and the tunnel diode. 

J b = f e -(cu(E)Xe + lll>(E)Xb + C!l(E)Xt)JlnGaA.s(E)¢(E)d(E) 1 

Where Xe, Xb, Xt are the emitter, base and tunnel diode thicknesses; ~(E), Cl.tJ(E) and Cl(E) are 

their absorption coefficients varying with doping levels, and eKE) is the solar spectrum. 

The top and bottom cell photocurrent matching is achieved by an appropriate choice of the 

individual layer thicknesses. 

The expression of the dark current related to such diodes, is set by considering the saturation 

current I. (ref. 8), the tunnel current Itun (ref. 10), and the GR current Igr (ref. 8) with Igr = qnjWltgr . 

Where q is the electron charge, nj the intrinsic carrier density, W is the depletion region width, 

and 'tgr is the effective GR lifetime. 

The I-V characteristic of a single cell is then expressed by (ref. 8), 

I = J - I (e Y / Ut + 1) - I (e Y / 2Ut + 1) - I - V / RM, cc. II' tun 2 

Where ut=O.026V at room temperature, V=Vo-R.1 (R.: series resistance), and Rsh is the shunt 

resistance. 

The output voltage of the tandem under illumination is the sum of that from the individual 

constituent cells reduced by the voltage drop in the tunnel junction. The tandem photocurrent 

however is controlled by the cell generating the lower photocurrent. 

RESULTS AND DISCUSSION 

Table 1 shows the parameters required to achieve a near optimum design of NIP and PIN 

tandem solar cells respectively. 

The calculated performances are: 

n=1 

Jcc= 33.3 rnA 

VOCjnP= 1.17V 

VOCjoGaAs=O.64 V 

,,= 36.24% 

for a NIP structure and 

n<2 ('tgr= 1 OOms) 

Jcc= 28 rnA 

VOClnP= 1.1 V 

VocloGaAs=O.62 V 

,,= 30.5% 

n=2 

Jcc= 33.3 mA 

VOClnP= 0.97V 

VOCjoGaAs=O.5 V 

,,= 27% 

n=2 ('tgr=1 ms) 

Jcc= 28 rnA 

VOCjnP= 0.9V 

VOCjoGaAs=O.5 V 

,,= 22°k 

142 

---

I 

I 

I 

I 

I 
for a PIN structure (because of a high series resistance we can not assume a idealiy factor 

equals to unity). 

OPTIMIZA TION 

PIN or NIP ? 

A priory PIN junctions may seem to be more suitable than NIP structures so as to take 

advantage of the absorption shift present in p doped compounds (ref. 9). N doped layers 

however have a lower sheet resistance and p type Ino.53Gao.47As compounds show much better 

minority carriers mobilities (ref. 7). Furthennore the relatively low absorption coefficients in these 

materials will necessitate the use of thicker junction layers (of 4-5j.1m) to minimize transparency 

losses. Thus a NIP structure is required for good carrier collection before recombination. 

In this work the surface recombination velocities (SRV) values used are those found in 

InGaAsllnP double heterostructures and InPmonolayer. 

Emitters 

The top cell emitter thickness is calculated to allow an optimum perfonnance through a 

compromise between a minimization of the surface recombination rate and a reduced sheet 

resistance (ref. 11). 

The bottom cell emitter thickness is derived to achieve the best carrier collection before 

recombination. There is no sheet resistance and the SRV is small due to the presence of a 

window layer. Figures 2 shows for comparison the perfonnances expected in the case of PIN and 

NIP tandems versus emitter width with optimum doping and base width shown in table 1. We can 

see that the different values of minority carrier mobilities versus doping level lead to a difference 

of 6% of absolute efficiency if a NIP rather than PIN structure is considered. Also p doped 

emitters need to be thicker due to high sheet resistance. Consequently, the base is required to 

be thin in order to allow the photocurrent matching. Therefore the optimum design of PIN tandem 

is a top cell inverted structure ( emitter thicker than the base). 

Bases 

The base participation in the tandem perfonnance can also be optimized by adjusting the width 

and the doping level. The high electron mobility in p doped InGaAs allows the use of a thicker 

bottom cell base without significant carrier recombination. The top cell base is therefore critical 

due to photocurrent matching constraints. 

Figures 3 and 4 show the variation of the photocurrent and efficiency as a function of the InP 

base thickness assuming a perfect case of y= 1 (tgr = 100 1-15). We can see that light absorption in 

the InP base leads to an increase of the top cell photocurrent and a decay of the bottom cell 

photocurrent. The condition of equal currents (top cell thickness: 0.3j.1m) corresponds precisely to 

the optimum conversion efficiency of the tandem shown in figure 4 (36% AMO). The high 

correlation between the photocurrent matching condition and the multi-junction efficiency is 

clearly demonstrated. Figure 4 shows also the case where a 100% intemal efficiency is assumed 

in order to emphasize the influence of SRVs, lifetimes and diffusion lengths on the photoresponse. These combined intrinsic and extrinsic limitations result in a 20% change in absolute 

efficiency between the two models. 

The calculated intemal spectral response of the near optimum design tandem solar cell is 

presented on figure 5. It emerges clearly that the device limitations are mainly due to the bottom 

cell characteristics. The limited values of diffusion lengths and low absorption do not pennit 

better expectations. 

Dark current 

As mentioned previously, the dark current is controlled by three different processes. First the 

saturation current Is which varies with the square of the intrinsic carrier concentration (ref. 4) 

leads to an intrinsic limitation. Second is the band to band tunneling current which becomes non 

negligible only in highly doped junctions. Third is the Generation-Recombination current which is 

143 
-~--

the main factor to be minimized by the growth technique. The density of trap levels in the 

bandgap is taken into account in the I-V characteristics as an effective GR lifetime (equation 2). 

To describe the effect of this latter parameter on the cell perfonnance, we have drawn in figure 6 

the variation of the conversion efficiency of a tandem over a wide range of GR lifetimes yielding 

ideality factors ranging from 2 to 1. There is a non negligible 8 % in absolute efficiencies 

enhancement when GR lifetimes vary from 1 ~ to 20 ~. 

Tunnel junction 

In the case of a monolithic tandem structure, the electrical interconnect between the two cells is 

achieved by an InGaAs tunnel junction (ref. 12). An ideal tunnel junction should modify neither 

the electrical nor the optical properties of the tandem. It has been shown previously (ref. 12) that 

very thin and extremely low resistivity InGaAs tunnel junctions can be fabricated with CBE. In 

addition, through patterning of the tunnel junction we can further minimize light absorption in this 

region (ref. 13). The influence of light absorption in the tunnel junction area on the tandem 

efficiency is shows in figure 7. 

CONCLUSION 

The characteristics of an InPllnGaAs monolithic solar cell have been evaluated using a large 

range of geometric considerations. Our calculations demonstrate the influence of the emitter, 

base, and tunnel junction layer thicknesses and have allowed us to set a near optimum design 

for a maximum tandem efficiency. The minority carrier mobilities, especially in InGaAs, seem to 

suggest the superiority of NIP over PIN structures. Finally it has been demonstrated that 

photocurrent matching is the principal condition for a near optimum design and that high 

efficiency tandem solar cells require a low trap level density. 

REFERENCES 

11 M. W. Wanlass, J. S. Ward, K. A. Emery, T. A. Gessert, C. R. Osterwald and T. J. 

Coutts; Solar Cells, 30 (1991) 363-371. 

21 M. W. Wanlass, J. S. Ward, K. A. Emery and T. J. Coutts; 

31 A. Bensaoula, N. Medelci, M. F. Vilela and A. Freundlich; 6th Int. Cont. on Indium Phosphide 

and Related Materials; (1994) Santa Barbara Ca. pp. 280. 

41 A. Zemel and M. Galland; J. Appl. Phys. 64 (11), pp. 6552 (1988). 

51 See as an example' A new structure of GaAs solar cell· M. F. Vilela, B. Beaumont, 

J. C. Guillaume, P. Gibart and C. Verie, 11th Eur. PV Energy Conf. (Montreux, 

Switzerland, 1992) 

61 Properties of Lattice-Matched and strained Indium Gallium Arsenide. 1993 INSPEC, 

(ed. london). 

7/Properties of Indium Phosphide 1991 INSPEC (ed.london and New York). 

81 H. J. Hovel. Solar cells in Semiconductors and semimetals. Vol. " 

(Ed. R. K. Willarson, A. C. Ber. Acad. Press, 1975 

91 H. Burkhard, H. W. Dinges, and E. Kuphal; J. Appl. Phys. 53 (1), 655 (1982). 

101 E. O. Kane, J. Appl. Phys. 32 (1). pp 83 (1961). 

144 

I ----' 
11/ C. Goradia, W. Thesling, I. Weinberg, 22nd IEEE PV. Specialists Cont.; Las Vegas 

(1991). 

121 M. F. Vilela, A. Freundlich, A. Bensaoula and N. Medelci; 5th Int. Conf. on InP and 

related materials, Paris, France, pp. 227 (1993). 

13/ A. Freundlich, M. F. Vilela, A. Bensaoula, N. Medelcl, Proc. 23rd IEEE-PV Conf. 

(1993), pp. 644. 

Metallization, Contact layer 

Emitter 

Base InP Top Cell 

Tunnel junction 

Window 

Emitter 

Base InO.53GaO.47As Bottom Cell 

Base wafer lattice matched to InP 

Figure 1: Schematic representation of the monolithic tandem solar cell. 

145 
Table 1. 

Top cell InP Bottom cell Ino 53GaO 47As 

NIP PIN NIP PIN 

Emitter 

SRV 105 cm2

s-1 105 cm2

s1 103 cm2s- 1 103 cm2

s-1 

Doping 3.1018 cm3 3.1018 cm-3 1018 em3 1018 cm3 

Thickness 0.02 Ilm 0.15 1lm 0.5 Ilm 1 Ilm 

Lifetime 1 ns 1 ns 0.3 ns 0.3 ns 

Diff length 0.45 jllTl 2.5 Illll 0.4 Ilm 2.3 Ilm 

Base 

SRV 103 cm2

s- 1 103 cm2s1 103 cm2

s1 103 cm2

s1 

Doping 5.1017 cm-3 5.1017 em3 1017 em-3 1017 em-3 

Thickness 0.25 jllTl 0.05 Illll 4.5 Ilm 3jllTl Lifetime 23ns 23ns 19 ns 19 ns 

Diff length 13, 8 jllTl 1.35 Illll 7 Ilm 1.5 Ilm 

Series Resistance 0.970 20 

Shunt resistance 104 0 104 n 

The near optimum design has been calculated through a systematic vanatlon of all the relevant 

cell parameters. 

0.>1 

go", 

.!! 

~o.ll 

~ 020 

~ 0." 

80.10 

0.06 

- '1M 

-WI> 

O~L;~~OLl~~O~~~~O~~~~OA~~OS 

Emitter width (micron) 

0" 

, 

, 

, 

, 

, 

, 

, 

, , , , 

-lnPtopcell 

- - - - InGaA. bonom cell 

o .,4'----'---------,'------"'--'----:''::------.J'-----' 

0.0 02 OA 0.1 0.1 1.D 12 

InP base Thickness (mlClOn$) 

Figure 2 : Conversion efficiency of NIP and PIN Figure 3: The effect of the top cell base 

tandem solar cell vs. top emitter width with thickness (emitter = 0.05Ilm) on the 

optimum doping and base width. photocurrent of a tandem cell with a pattemed 

tunnel junction. 

146 

- - - .--- - -------

. - • I 
., . . 

SESSION III 

RADIATION DAMAGE 

AND 

ENVIRONMENTAL TESTING 

Id1 

I /Y'?' 

I ~ PAGE BLANK NOT FILMED 
'3917~7 

N95-20518 

PERFORMANCE, DEFECT BEHAVIOR AND CARRIER ENHANCEMENT IN LOW ENERGY, 

PROTON IRRADIATED p+nn+ InP SOLAR CELLS 

/0/. 

I. Weinberg and G.C. Rybicki 

NASA Lewis Research Center 

Cleveland, Ohio 

C. Vargas-Aburto 

Kent State University 

Kent, Ohio 

R.K. Jain 

University of Toledo 

Toledo, Ohio 

and 

D. Scheiman 

NYMA, Inc. 

Brook Park, Ohio 

InP p+nn+ cells, processed by MOCVD, were irradiated by 0.2 MeV 

protons and their performance and defect behavior observed to a 

maximum fluence of 1013 cm- 2 . Their radiation induced degradation, 

over this fluence range, was considerably less than observed for 

similarly irradiated, diffused junction n+p InP cells. Significant 

degradation occured in both the cell's emitter and base regions the 

least degradation occuring in the depletion region. A significant 

increase in series resistance occurs at the highest fluence. Two 

majority carrier defect levels, E7 and EIO, are observed by DLTS 

with activation energies at (Ec -0.39)eV and (Ec -0.74)eV 

respectively. The relative concentration of these defects differs 

considerably from that observed after 1 MeV electron irradiation. An 

increased carrier concentration in the cell's n-region was observed 

at the highest proton fluence, the change in carrier concentration 

being insignificant at the lower fluences. In agreement with 

previous results, for 1 and 1.5 MeV electron irradiated InP p+n 

junctions, the defect level EIO is attributed to a complex between 

zinc, diffused into the n-region from the zinc doped emitter, and a 

radiation induced defect. The latter is assumed to be either a 

phosphorus vacancy or interstitial. The increased, or enhanced 

carrier concentration is attributed to this complex acting as a 

donor. 

INTRODUCTION 

The highest AMO efficiency (19.1%) InP solar cell consisted of an 

n+pp+ structure epitaxially grown on a p+ InP substrate [1]. 

However, the high cost and relative fragility of InP served as 

motivation for research efforts directed at heteroepitaxial growth 

of InP on more viable substrates [2,3]. The highest AMO efficiency 

(13.7%) for this type of cell was achieved using a GaAs substrate 

[3,4]. Considering only cost and fracture toughness, Si would be 

the preferred substrate. The fact that Si is a donor in InP 

introduces complexities which are necessary in order to avoid the 

formation of an efficiency limiting counterdiode [5]. One method 

used to overcome this problem, lies in employing an n+p+ 

149 

---- --

I 

_J 
tunnel junction in contact with the cell's p region. A simpler 

method consists of using an n+ substrate and processing the cell in 

the p+nn+ configuration. This eliminates the need for a tunnel 

junction. Unfortunately, the pin configuration has received 

relatively little attention the best cell with this geometry having 

achieved an efficiency of 17% [6]. Irradiation of these 

homoepitaxial cells, with 1 MeV electrons, showed that they were 

slightly more radiation resistant than diffused juntion nip cells 

[7]. Additional pin InP cells have been processed by closed ampoule 

diffusion [8]. Currently, there has been some activity aimed at 

producing heteroepitaxial p+nn+ InP cells using n+ Ge substrates 

[9]. Since, like Si, Ge is an n-dopant in InP, use of this 

configuration obviates the need for a tunnel junction. Obviously, 

before attempting to process heteroepitaxial cells, one must produce 

a reasonably good homoepitaxial cell. In the present case we focus 

our attention on homoepitaxial p+nn+ cells processed prior to 

producing the cells heteroepitaxially on an n+ Ge substrate [9]. 

EXPERIMENTAL DETAILS 

The cells were processed by MOCVD, at the Spire Corporation, under 

contract to NASA Lewis. Cell configuration, dopants and 

concentrations are shown in fig. 1. Processing details can be found 

in reference 9. Irradiations by 0.2 MeV protons, to a fluence of 

l013/cm2 were performed at the University of Michigan's ion 

implantation facility. Cell performance was determined at NASA 

Lewis using a Spectrolab Mark II, xenon arc solar simulator with 

flight calibrated InP standard cell. Spectral response and Isc-Voc 

measurements were also performed before irradiation and at each step 

in the irradiation process. Carrier concentrations in the cell's 

p-base, near the junction were determined by capacitance-voltage 

measurements. Defect behavior was monitored by DLTS measurements. 

RESULTS AND DISCUSSION 

Performance 

Pre-irradiation performance parameters are shown in table I. 

Considering the fact that theoretical modelling indicates possible 

efficiencies over 22%, the present efficiencies are excessively low 

[10]. This is attributable to the fact that the present cells were 

processed in an early stage of development. In fact, AMO 

efficiencies of 17% have subsequently been achieved at Spire [6]. 

Higher efficiencies can be anticipated with additional effort. 

The results of the 0.2 MeV proton irradiations are shown in fig.2. 

Comparision of normalized efficiencies with 0.2 MeV proton 

irradiated diffused junction nip cells is shown in fig.3. The nip 

cells had the same junction depth as the present cells with AMO 

efficiency=lS.l%, Voc=823 mV, Jsc=29.4 ma/cm2 , and FF=8S.6% [11]. 

Comparision of normalized efficiencies indicates considerably more 

radiation resistance for the present cells at the h~gher fluences. 

Also, comparing numerical efficiency values, the present cells 

outperform the nip cells at the higher fluences. 

150 
r

_________________

The external quantum efficiency, before irradiation and at a fluence 

of 1012 cm-2 , is shown in fig.4. The quantum efficiency at the 

highest fluence is lost in the system noise and is therefore not 

shown in the figure. Figure 4 indicates that considerable 

degradation occurs in both the emitter and base of the solar cell. 

A numerical estimate of the relative degradation is obtained using 

the relation 

Jsc = 

la 

Where SR(Aj) the spectral response, in rnA/mW, is obtained from the 

quantum efficiency using the relation 

lb 

where E(A') in mW cm- 2 micron-l is the solar spectral radiance at 

wavelengt~ Aj in microns, QE( Aj) is the external quantum 

efficiency at Aj, ~ Aj is an appropriate wavelength interval and 

the summation is over all wavelengths covered by the quantum 

efficiency in fig.4. The junction depth is approximated by the 

optical path length l/~j where C<j is the absorption coefficient at 

wavelength Aj. Using la and lb it is found that the degradation in 

short circuit current is approximately divided between the emitter 

and base. An estimate of the relative degradation in base and 

emitter is obtained from the Isc-Voc measurements obtained over a 

range of light intensities. The results before irradiation and at a 

specific fluence are shown in table II. Considering the reverse 

saturation currents J02 is attributed to recombination in the cell's 

depletion region while the major contribution to JOI arises from 

diffusion in the base of the cell. Hence, from the diffusion and 

recombination current densities in table II, it is concluded that 

the radiation induced degradation in the cell's base is much greater 

than that occuring in the depletion region. 

Defects 

The DLTS spectrum, at the highest fluence, is shown in fig.S while 

defect parameters are listed in table III. No defect levels were 

observed prior to irradiation. The defect concentrations obtained 

from fig.5, and shown in the table, have been corrected for band 

bending and its effect on space charge when crossing the Fermi level 

[12]. The majority carrier defect levels labelled E7 and EIO have 

been observed previously after irradiation by 1 and 1.5 MeV 

electrons [13]. The broad signal observed between E7 and EIO 

appears to be due due to the presence of one or more unresolvable 

defect levels. The present energy levels for E7 and EIO are in 

reasonable agreement with those previously reported for these 

defects [13]. However, the concentration ratio NT(EIO)/NT(E7) rv SO 

in the previously electron irradiated case [13] while in the present 

case the ratio is 0.56. Hence, although EIO could reasonably be 

assumed to be the major radiation induced defect, observed by DLTS, 

in the n region of electron irradiated p+n InP [13], the choice of 

major defect is not clear cut in the present case. tt is noted 

that, after electron irradiation, E7 was observed, but not EIO, in 

151 
the DLTS spectrum obtained using a Schottky barrier on n-type InP 

[13]. In fact, E10 was only seen when a p+n junction was used, the 

p+ region being heavily zinc doped. It was therefore initially 

concluded that E10 was a result of zinc diffusion into the n-region, 

the zinc complexing with an unidentified radiation induced defect 

[13]. Another possibility considered was the formation of a complex 

between a process induced and radiation induced defect [13]. 

Carrier concentrations in the n-region, measured over a range of 

fluences, indicated that at all but the highest fluence, the change 

in carrier concentration was negligible. However, at the highest 

fluence, the carrier concentration was significantly increased over 

the pre-irradiation value (table IV). This is in opposition to 

observations in the p-region of n+p InP where both proton and 

electron irradiations produced decreased carrier concentrations 

[14,15]. Although this appears to be the first reported observation 

of carrier enhancement in proton irradiated InP, the effect has been 

observed before after 1 MeV electron irradiation [16]. It is 

significant that, in both cases, carrier enhancement is observed in 

the n-region of an InP p+n diode where zinc is used as the p-dopant 

[16]. In agreement with the previous suggestion [13] it was argued 

that zinc diffuses into the n-region and complexes with a radiation 

induced defect [16]. In the latter case it was further argued that 

the defect was either a phosphorus interstitial or vacancy [16]. 

Furthermore, it was argued that the complex acts as a donor. In 

relation to the present solar cell parameters, the carrier 

enhancement does not appear to be a factor in improving cell 

performance or in decreasing series resistance. This is evident 

from fig. 1 and table V, the latter showing that cell series 

resistance, obtained from dark diode I-V data, increases 

significantly at the highest fluence. In any event, the present 

results indicate that, despite the anomolous increase in carrier 

concentration, the effects of radiation induced defects on transport 

properties, such as diffusion length are dominant in determining 

cell behavior under the present low energy proton irradiations. 

CONCLUSION 

Under irradiation by 0.2 MeV protons, it is concluded that; 

The radiation induced degradation is considerably lower in the 

present cell when compared to diffused junction n+p InP cells. 

Considerable radiation induced degradation is observed in both the 

base and emitter of the present cells, both degradations being 

considerably greater than that occurring in the depletion region. 

152 
The relative concentration of defects E7 and EIO, NT(EIO)/NT(E7), is 

considerably greater under 1 and 1.5 MeV electron irradiations 

than is the case for the present irradiations. 

A significant increase in carrier concentration (carrier 

enhancement) occurs in the cell's n-region after irradiation by 0.2 

MeV protons at a fluence of 10 13 cm-2 • At the lower fluences, the 

change in carrier concentration is insignificant. 

In concurrence with previous conclusions after electron irradiation, 

the defect level EIO is attributable to a donor complex formed 

between zinc and a radiation induced defect. The latter is assumed 

to be either a phosphorus interstitial or vacancy. 

In the present case, the radiation induced carrier enhancement 

appears to have little or no effect on cell performance. 

REFERENCES 

1. C.J.Keavney, V.E.Haven and S.M.Vernon, "Emitter Structures in 

MOCVD InP Solar Cells," 21st IEEE Photovoltaic Spec. Conf., 

pp141-147 (1990) 

2. C.J.Keavney, S.M.Vernon and V.E.Haven, "Tunnel Junctions for 

InP on Si Solar cells," Proceedings Space Photovoltaic Research and 

Technology Conf., ppl-l to 1-7, NASA Conf. Publication 3121 (1991) 

3. M.W.Wanlass, u.S. Patent No. 4,963,949 (1990) 

4 T.J.Coutts, "Progress in InP-Based Solar Cells," Proceedings 

3rd Int'l Conf. on InP and Related Materials, pp20-3l, IEEE-LEOS, 

Piscataway N.J. (1991) 

5. C.J.Keavney, S.M. Vernon, V.E.Haven, S.J.Wojtczuk and 

M.M.Al-Jassim, "Fabrication of n+p InP Solar Cells on Silicon 

Substrates," Appl. Phys. Lett.54, ppl139-l141 (1989) 

6. S.Wojtczuk, S.M. Vernon and E.A.Burke, "pin InP Solar Cells on 

Ge Wafers," This Conference (1994) 

7. I.Weinberg, C.K.Swartz and R.E.Hart Jr., "Radiation and 

Temperature Effects in Gallium Arsenide, Indium Phosphide and 

Silicon Solar Cells," Proceedings 19th Photovoltaic Spec. Conf., 

pp548-557, IEEE, Piscataway N.J. (1987) 

8. Mircea Faur, Maria Faur, C.Goradia, M.Goradia, D.J.Flood, 

D.J.Brinker, I.Weinberg and N.Fatemi, ·Progress in p+n InP Solar 

Cells Fabricated by Thermal Diffusion,· Proceedings, Space 

Photovoltaic Research and Technology Conf., pp23-32, NASA Conf. Pub. 

3210 (1992) 

153 
9. S.Wojtczuk and S.Vernon, "pin InP Solar Cells on Ge Wafers," 

Space Photovoltaic Research and Technology Conf. NASA Lewis Res. 

Center (1994) 

10. R.K.Jain, I.Weinberg and D.J.Flood, "Comparative Modelling of 

InP Solar Cell Structures," Proceedings Space Photovoltaic Rand T 

Conf., pp29-1 to 29-9, NASA Conf. Pub. 3121 (1991) 

11. T.Takamoto, H.Okazaki, H.Takamura, M.Ohmori, M.Ura and 

M.Yamaguchi, "Radiation Effects on n+p Homojunction Indium Phosphide 

Solar Cells," Proceedings 2nd Intll conf. on InP and Related 

Materials pp62-65, IEEE-LEOS, Piscataway N.J. (1990) 

12. Y.Zohta and M.O.Watanabe, "On The Determination of the Spatial 

Distribution of Deep Centers in Semiconducting Films From 

Capacitance Transient Spectroscopy," J. Appl. Phys.,53, pp1809-1811 

(1982) 

13. J.L.Benton, M.Levinson, A.T.Macrander, H.Temkin and 

L.C.Kimerling, "Recombination Enhanced Defect Annealing in n-InP," 

Appl. Phys. Lett.,45, pp566-568 (1984) 

14. I.Weinberg, R.K.Jain, C.K.Swartz, H.B.Curtis, D.J.Brinker, 

C.Vargos-Aburto and P.J.Drevinsky, "Proton Irradiated 

Heteroepitaxial InP Cells," Proceedings 23rd IEEE Photovoltaic Spec. 

Conf., pp1483-1486, IEEE, Piscataway N.J. (1993) 

15. I.Weinberg, C.K.Swartz and P.J.Drevinsky, "Carrier Removal and 

Defect Behavior in p-type InP," J. Appl. Phys.72, pp5509-5511 (1992) 

16. A.Sibi le, "Origin of the Main Deep Electron Trap in Electron 

Irradiated InP," Appl. Phys. Lett.,48, pp593-595 (1986) 

17. K.Rajkanen and J.Shewchun, "A Better Approach to the 

Evaluation of the Series Resistance of Solar Cells," Solid State 

Electronics 22, pp193-197 (1979) 

Table I: Cell Pre-Irradiation Performance Parameters 

Jsc Vac FF Eff. 

mA/cm2 mV % % 

23.6 851.4 84.7 12.4 

154 
Table II: Diffusion and Recombination Components 

of Reverse saturation Currents 

Ep = 0.2 MeV 

Fluence Al ~ J 01 J 02 

cm- 2 A/cm:1 A/cm2 

0 0.89 1.57 1.03 x 10-18 3.9 X 10-12 

1012 1.32 2.08 1.5 x 10-11 5.5 x 10- 8 

Table III: Characteristics of Majority Carrier Defect 

Levels Obtained by DLTS in p+n n+ lnP 

Ep = 0.2 MeV 

Fluence = 10 13 /cm2 

DEFECT ACTIVATION CAPTURE CONCENTRATION INTRODUCTION 

E7 

E10 

ENERGY CROSS SECTION RATE 

eV cm2 cm-1 cm- 1 

Ec-0.39 4.1 x 10-17 8.67 X 1015 867 

Ec-0.74 3.6 x 10-14 4.88 X 1015 488 

Table IV: Electron Concentration in n-Region of 0.2 MeV 

Proton Irradiated p+nn+ InP Cell 

FLUENCE ELECTRON CONC. CHANGE IN 

ELECTRON CONCA 

cm- 2 cm- 1 cm- 3 

0 3.18 X 1016 0 

1013 3.86 X 1016 +6.8 x 1015 

a Change in carrier conc. was negligible at the lower fluence. 

155 
Table V: Series Resistance in 0.2 MeV Proton Irradiated Cell 

FLUENCE 

(CM- 2 ) 

SERIES RESISTANCE 

OHM - cm1 

T - ([<J I CRONS) 

0.25 

1.5 

0.5 

0 2 X 1011 

0.49 0.36 

~ ~ 

P+-2E18 (ZN) 

N-3 .. 2E16 (SI) 

N+-5El8 (SI) 

1 -- -- N+-SEl8 (5) J 

I III allllZ! 

0.2 MEV PROTON RANGE-lo5 1·1ICRONS 

FIGURE 1. INP CELL DETAILS 

156 

1012 1013 

0.56 1.7 
1.0 

·0.8 

c::: 

UJ 

I--

!;tl 0.6 ~ <': 

0-

c:::> 

Ul 

N 

:J 

~ 0.4 

~ 0 

= 

0.2 

o 

1.0 

0.8 

>- LJ 

= UJ 

---. 0.6 L.) 

---. 

u.. 

u.. 

UJ 

Q 

W 

N 

:J 0.4 <': 

~ C> 

= 

0.2 

o 

o EFFICIENCY 

<> Vac 

o Isc 

t::" FILL FACTOR 

~ __ ~'~, ________ L-____________ ~ ______________ ~ 

o 

0.2 MEV PROTON FLUENCE (CM-2) 

FIGURE 2. NORMUZED INP .CELL PARAMETERS AFTER IRRADIAum; 

0 

-D- P+NN+ PRESENT DATA 

o N+P REF . 11 o 

~----~c~------~~------------~--------------~o 

o 

0.2 MEV PROTON FLUmCE (CM-2) 

FIGURE 3. COMPARISION OF p+rm+ AND N+P IRRADIAT@ INP SOLAR CELLS 

157 

I 

I 

J 
1.0 

0,8 

>- U 

:>!: 

UJ 

-U 

u.. 

u.. 0,6 w 

:E: 

I-= 

z: 

c:( 

= OJ 

...J 0,4 c:( 

z: 

ex: 

W 

S 

0,2 

o 

0.3 0,4 0,5 0,6 0.7 O,G 0,9 1.0 

WAVELENGTH (MICROMETERS) 

FIGURE 4, EXTERNAL QUANTU~ I EFFICIENCY I P+NN+ INP, Ep=O,2 MEV 

100 --_._------.- ~3 MS ,.... 

en 

I-

~ 80 = = 

~ c:( 

cr:: E7 I- 60 ~ 

~ c:( 

~ 

-l 

c:( ~o z: 

l!J 

~ 

en 

(/) ElO ~ 20 ~ 

0 __ I J 

200 250 300 350 400 450 

TEMPERATURE (K) 

FIGURE 5, DLTS OF P+NN+ I NP CELL, Ep=O, 2 NEV, FLUEtiCE= 1013 CM-2 

158 
N95- 20519 

ELECTRON-IRRADIATED TWO-TERMINAL, MONOLITHIC InP/Ga0.47Ino.53As TANDEM 

SOLAR CELLS AND ANNEALING OF RADIATION DAMAGE 

H.L. Cotal, R.J. Walters, and G.P. Summers 

U.S. Naval Research Laboratory 

Washington, DC 

and 

S.R. Messenger 

SFA, Inc. 

landover, Maryland 

ABSTRACT 

Radiation damage results from two-terminal monolithic InP/Gao 47Ino 5~s tandem solar cells subject to 1 

MeV electron irradiation are presented. Efficiencies greater than 22 % have been measured by the 

National Renewable Energy Laboratory from 2 x2 cm cells at 1 sun, AMO (25°C). The short circuit 

current density, open circuit voltage and fill factor are found to tolerate the same amount of radiation at 

low fluences. At high fluence levels, slight differences are observed. Decreasing the base dopant level of 

the Ga0.47Ino.5~s bottom cell improved the radiation resistance of Jsc dramaticaIlY'15Thi~}n tum, extended 

the serie;; current flow through the subcells substantially up to a fluence of 3x 10 cm compared to 3x 

10

14 

cm' , as observed previously. The degradation of the maximum power output from the tandem 

device is comparable to that from shallow homojunction (SHJ) InP solar cells, and the mechanisms 

responsible for such degradation is explained in terms of the radiation response of the component cells. 

Annealing studies revealed that the recovery of the tandem cell response is mostly dictated by the 

annealing characteristics exhibited by SHJ InP solar cells. 

INTRODUCTION 

In the past few years, tandem solar cells have been the center of much research as an alternate 

source of generating electrical power in space applications (refs. 1-4). However, for a solar cell to be 

suitable for this application, it must tolerate the harsh radiation environment of space. Such a cell is the 

two-terminal, monolithic InP/Ga0.47Ino ~s tandem solar cell grown by the National Renewable Energy 

Laboratory. In collaboration with the Naval Research Laboratory (NRL), the tandem cell program has 

been directed toward optimizing the radiation resistance of the tandem cells by improving the device 

str~cture. To date, this has lead to the fabrication of two-terminal, monolithic Rrototypes as large as 2x2 

cm with beginning of life (BOL) efficiencies greater than 22% (1 sun, AMO, 25°C). Several of these cells 

will be tested in the STRV 1 space experiment soon (ref. 5). 

An InP-based approach has been integrated into the tandem technology mainly because of the 

proven higher radiation resistance of shallow homojunction (SHJ) InP cells than other types such as Si 

and GaAs (ref. 6). A key issue in the development of the two-terminal tandem device is to design each 

subcell so that the series current flow through both junctions is matched end-of-life (EOl) after 

irradiation. Preliminary studies on InP/Gao.47Ino.53As tandem cells irradiated with successive f1uences of 1 

MeV electrons have already shown promising results (ref. 2). The subcell currents remained equal at 

relatively high fluences, and the degradation of the photovoltaic (PV) parameters occurred at a slow rate. 

By varying the base dopant level of the Gao 47Ino~s bottom cell, it was found that the rate of decay of 

the short-circuit current density (Joe> and the open-circuit voltage (VoJ is more pronounced for the 

heavily-doped case by the former parameter but less pronounced by the latter (ref. 7). As a first step in 

the optimization procedure for the tandem cell, the base dopant level of the bottom cell was reduced. 

The ability of a solar cell to recover from radiation damage is also an important aspect of the overall 

cell performance. The recovery of irradiated tandem cells due to thermal annealing in the dark. is 

presented, and the recovery is analyzed in terms of the annealing characteristics of the individual cells. 

The tandem cells were tested with successive fluences of 1 MeV electrons. Illuminated currentvoltage (I-V) measurements and annealing results are presented for InP/Ga0.47Ino.53As solar cells. The 

results are compared with previous studies on InP/Gao 47Ino~s cells and with each component cell, all 

with a similar structure. 

159 
CELL DESCRIPTION 

The tandem cells were grown by the atmospheric-pressure metal organic vapor phase epitaxy 

(APMOVPE) technique for which the details have been described elsewhere (ref. 8). The cell structure 

consists of an InP top junction, a Gao 47Ino ~s bottom junction and a Ga0.47Ino ~s tunnel junction to 

provide the electrical connection between the subcells, as shown schematically in Fig 1. The component 

cells are lattice-matched by adjusting the Ga and In compositions. The total area of the tandems in this 2 study were 1 x 1 cm . 

EXPERIMENTAL PROCEDURE 

The cells were irradiated with incremental fluences of 1 MeV electrons at NASA Goddard, and the 

fluence was measured with a Faraday cup. The irradiations were performed in ambient conditions at 

room temperature under open circuit. The beam current was usuallv kept in the nanoamp regime to • 13 16-2 avoid heating the device, and the electron fluence ranged from 1x10 to 2.4x10 cm . Measurements 

of the I-V characteristics were made within three hours after each irradiation. The results were obtained 

under one sun, AMO at 25°C using a 2500 W SpectroLab, Mark III solar simulator. The efficiency is 

measured using an InP reference cell calibrated by NASA lewis Research Center. Since the band gap of 

Gao.47'no 5~s is 0.75 eV, it should be pOinted out that the IR portion of the simulator in this energy region 

is somewhat stronger than what it should be. This is characteristic of the uncorrected Xe sources 

employed, and the overall effect this has on the I-V curve is to enhance the fill factor by a very small 

percentage (ref 9). Annealing of the radiation damage was carried out in the dark and in the air up to 

500 K. 

RESULTS 

The effects of radiation damage on the cell I-V curves are illustrated in Fig_ 2_ The main feature of 

importance from the I-V curve is the smooth kink which begins to develop at low voltages as a result of 

the current mismatch between the top and bottom cells. The kink appears at a fluence of about 3x10

15 

cm2

, and becomes more pronounced as the irradiation increases. This feature has been associated with 

the reverse-bias breakdown of the Ga0.47'no.5~s bottom cell, which has been observed before in 

InP/Ga0.47'nO~s devices (ref. 2). (Not related to radiation damage, however, a mismatch in the currents 

between the subcells at BOl has been seen due to the spectral content of different light sources (ref. 9).) 

This allows the larger generated current from the top subcell to pass through the tandem . The data also 

show excellent radiation resistance in these partially optimized cells despite the presence of a slight BOl 

current mismatch between the subcells. 

Figure 3 depicts the change of the PV parameters with electron fluence for a tandem ceiL At low 

f1uences, the degradation of the maximum power (P max) output of the cell was primarily due to the 

reduction of Jsc and Vex; ' The power loss at a fluence of 1x1015 

cm2 was about 20% relative to the BOl 

power. An additional plot of the degradation of the efficiency of a SHJ InP solar cell is included in the 

figure for comparison. The degradation of the tandem cell efficiency is comparable to that of the SHJ InP 

cell. It should be noted that for GaAs/Ge solar cells the FF has been reported to be affected by the 

infrared portion of the excitation source (ref. 9). Since the infrared content of the Xe source used in this 

study has not been entirely suppressed to simulate the sun's true IR spectrum, the FF considered here is 

that of an effective FF. Although this should boost the cell efficiency by a small percentage, the radiation 

response and the annealing behavior of the solar cells are the issue of importance and not the absolute 

values of the PV parameters. 

Illustrated in Fig. 4 are two I-V curves measured after irradiation with 1 MeV electrons at a fluence of 

3x1015 e-/cm2 for the tandem cell in this work, and a tandem cell from a previous study for comparison. 

The present cell is offset by 0.25 for clarity. At this fluence, P max from the early tandem is suppressed by 

the radiation-induced kink which appears as a result of the current mismatch. The present tandem cell 

does not show a pronounced current mismatch at this fluence, and in consequence, P max is considerably 

higher. Therefore, the devices studied here show a dramatic enhancement of the radiation resistance 

compared to the previous cell. This is known to be due to the reduced carrier concentration in the base of 

the bottom cell (ref. 2), and is discussed below. 

A desirable feature of space solar cells is the ability to thermally anneal the radiation damage while 

being subject particle irradiation. Annealing of InP/Ga0.47'no.53As cells exposed to 2.4x1 0

16 

cm-2 electron 

160 
--- ~ 

irradiation was carried out to examine the recovery of the cell response. Fig. 5 depicts a few I-V curves 

that were isochronally and isothermally annealed, both in 30-minute increments. As illustrated in the 

figure, the I-V curves gradually recover from the post-rad curve but fall short from recovering completely. 

The behavior of PV parameters as a function of annealing temperature is shown in Fig. 6. The preannealed value of each parameter is normalized with respect to the corresponding pre-rad value. Partial 

recovery of Jse is seen to begin at 350 K and begins to level off near 450 K and stays constant up to 500 

K. At this temperature, the sample was annealed isothermally to check whether more recovery was 

possible. All parameters except for Eff% showed no further appreciable recovery. At 500 K for 1.5 hours, 

P max has recovered by 16% from its pre-anneal value. 

DISCUSSION 

Recent studies on prototype, two-terminal, monolithic InP/Gao 47Ino ~s solar cells demonstrated that 

these devices perform very well when exposed to 1 MeV electrons. It has been shown that the 

degr~dation of the PV response is similar to that observed in SHJ InP cells up to a f1uence of 3x1 0

14 

efcm . The results were very encouraging considering that the previous tandem cells were not optimized 

for radiation resistance. The present study reports on the work of a first step in optimizing the tandem cell 

structure for radiation resistance. 

In an effort to optimize the InPfGa0.47Ino ~s tandem solar cells so that the subcell currents remain 

matched at EOl in a space environment, the PV response of the component cells must be adjusted. 

Particularly, Jse must be designed to be highly tolerant to radiation. Work dealing with the effects of base 

dopant level on Jse from Ga0.47Ino ~s and InP solar cells suggest that the radiation hardness of Jse is 

improved by decreasing the base dopant concentration (refs. 2, 7, 10). Similar results have been 

obtained from GaAs single junction solar cells (ref. 11). In view pf t~!S , the base dopant level of the 

bottom subcell was reduced by an order of magnitude (to about 101 

cm ) which caused the resistance of 

Jse to increase. In consequence, the onset of the current mismatch between the subcells was extended to 

substantially higher fluence levels. From previous tandem cells, the mismatch occurs at 3x1014 

e-/cm2 

whereas in this study it occurs at 3x1015 e-/cm-2

, which is a dramatic improvement. The improved current 

matching which is, in tum, due to the radiation-hardening of Jse, is a result of the decreased sensitivity of 

charge collection to the radiation-induced degradation of the minority carrier diffusion length (or its 

lifetime) in the bottom cell (ref. 10). 

It j~ i~teresting to note that the fill factor is only reduced by 8% relative to the BOl value at a f1uence 

of 10 e /cm2

• Using a technique developed by Handy (ref. 12), intensity-dependent measurements of 

the I-V curves for the tandem cells show that the series resistance is in the order of a few milliohms for 

the f1uence range studied. In addition to this, analysis of dark I-V fits suggest that the ~eries Iesistance 

may not even playa role in the degradation of Gao 47Ino~s devices at a fluence of 10

1 e/cm (ref. 13). 

This would seem to suggest that the overall series resistance has almost a negligible effect on FF, and 

thus the device performance. The low series resistance is probably due to the heavily doped base of 

each subcell thus decreasing the overall material resistivities. 

Partial recovery is seen in the irradiated tandem cells. The annealing characteristics exhibited by Jse 

and Vex; are similar to those reported elsewhere (ref. 14) on SHJ InP solar cells and they seem to be 

consistent with this type of cell. Particularly, in that study, no notable recovery of Jse, Vex; and P max is 

further seen at temperatures above 400 K. This appears to be consistent with other SHJ InP cells where 

a similar trend is observed (ref. 15, 16). The annealing behavior of the tandem cell can be described in 

terms of the minority carrier diffusion length of each component cell. The recovery of Jse under thermal 

annealing can be attributed to annealing of defects formed in the neutral regions of the device where Jse 

is mostly affected. This, in turn, results in an increase of the minority carrier diffusion length thus allowing 

the collection of charge carriers to be more efficient. A point worthy of mention is that according to 

unpublished results obtained at NRl, the recovery of Jse from the Gao.47Ino.53As bottom cell occurs at a 

faster rate than the InP top cell. This suggests that the recovery of Jse from the tandem cell is mainly 

characteristic of the recovery from the InP top cell. Likewise, Vex; from InP solar cells does not recover as 

fast as that from Ga0.47Ino.53As solar cells. This is an interesting result because it shows that unless the 

recovery rate of the PV response of the top subcell is the same as that of the bottom subcell, the tandem 

cell will not recovery effectively under thermal annealing. 

A final remark about the degradation of J and V can be made on the basis of their normalized pre- se 16 qc 2 

annealed values in Fig. 6. At a fluence of 2.4x 10 e fcm , Jse and Vex; have decreased to 62% and 78% 

from their pre-rad values, respectively. The modeling results from reference 10 combined with the results 

161 

~J 
of references 2 and 7 suggest that a more optimized cell design for the radiation resistance of Joe and 

V 0(;' the efficiency would be expected to increase by several percent at EOL. 

CONCLUSIONS 

High-performance, partly optimized InP/Gao "7Ino ~s tandem solar cells have been achieved, and 

have shown excellent radiation response. Efficiencies greater than 22% have already been obtained 

from 2x2 cm2 cells. After 1 MeV electron irradiation, the tandem cell*, with a BOL efficiency of over 20% 

have shown an efficiency of approximately 10% at a fluence of 10 6 e-/cm 2. This is higher than Si and 

GaAs, both with efficiencies of about 7.5% at the same fluence. Furthermore, the results described here 

show that the radiation resistance of Joe and V 0(; can be further optimized by fine-tuning the device 

structure, specifically, the base dopant level of the Ga0.47Ino s~s bottom cell. 

The annealing results revealed that the tandem cell annealing characteristics are mainly controlled by 

the InP top cell. This indicates that if annealing of tandem solar cells are viable while in earth orbit, then 

not only is the degradation rate of importance but also the annealing rate. That is, Joe in each subcell 

must recover at the same rate so that current matching can be maintained throughout the mission life. 

REFERENCES 

1. H.L. Cotal, G.P. Summers, and S.R Messenger, "Radiation Effects of Two-Terminal, Monolithic 

InP/Ga0.47Ino.s3As Tandem Solar Cells,"6th IEEE International Conference on Indium Phosphide and 

Related Materials, March 1994, pp. 496-499. 

2. RJ. Walters, S.R Messenger, M.W. Wanlass, and G.P. Summers, "1 MeV Electron Irradiation of 

Monolithic, Two-Terminal InP/Ga0.47Ino.s3As Solar Cells," 23rd IEEE Photovoltaic Specialists 

Conference, May 1993, pp. 1475-78. 

3. S.R Kurtz, J.M. Olson, KA Bertness, D.J. Friedman, A. Kibbler, B.T. Cavicchi, and D.O. Krut, 

"Radiation Hardness of Gaoslno.sP/GaAs Tandem Solar Cells," 10th Space Photovoltaic Research and 

Technology Conference, May 1991, pp.40-1 to 40-7. 

4. P.R Sharps, M.L. Timmons, R Vankatasubramanian, R Pickett, J.S. Hills, J. Hancock, J. Hutchby, 

P. lies, C.L. Chu, M.Wanlass, and J.S. Ward, "Development of 20% Efficient GainAsP Solar Cells," 

23rd IEEE Photovoltaic Specialists Conference, May 1993, pp. 633-638. 

5. C. Goodbody, and N. Monekoso, "The STRV-1 A & B Solar Cell Experiments,"23rd IEEE Photovoltaic 

Specialists Conference, May, 1993, pp. 1459-1464. 

6. RJ. Walters, S.R. Messenger, G.P. Summers, E.A. Burke, and C.J. Keavney, "Proton and Electron 

Irradiation of MOCVD InP Solar Cells: Experimental Results and Radiation Modeling," 22nd IEEE 

Photovoltaic Specialists Conference, Oct., 199. 

7. RJ. Walters, G.J. Shaw, S.R Messenger, and G.P. Summers, "The Effect of Base Dopant Level and 

Thickness on the Radiation Response of GalnAs Solar Cells," 5th IEEE International Conference on 

Indium Phosphide and Related Materials, April, 1993, pp. 602. 

8. M.W. Wanlass, J.S. Ward, K.A. Emery, and T.J. Coutts, "Monolithic, Series-Connected 

InP/Gao.47lnos3As Tandem Solar Cells," 223rd IEEE Photovoltaic Specialists Conference, May 1993, 

pp. 621-625. 

9. L.D. Partain, G.F. Virshup, and N.R. Kaminar, "Quantum Yield Spectra and I-V Properties of a GaAs 

Solar Cell Grown on a Ge Substrate," 20th IEEE Photovoltaic Specialists Conference, September 

1988, pp. 759-763. 

162 
9. L.D. Partain, G.F. Virshup, and N.R Kaminar, "Quantum Yield Spectra and I-V Properties of a GaAs 

Solar Cell Grown on a Ge Substrate," 20th IEEE Photovoltaic Specialists Conference, September 

1988, pp. 759-763. 

10. C.J. Keavney, RJ. Walters, and P.J. Drevinsky, "Optimizing the radiation resistance of InP Solar 

Cells: Effect of Dopant Density and Cell Thickness," J. Appl. Phys. 73, 1993, pp.60-70. 

11. M. Yamaguchi, and C. Amano, "Numerical Analysis for Radiation-Resistant GaAs Heteroface Solar 

Structures," J. Appl. Phys. 57, 1985, pp. 537-544. 

12. RJ. Handy, "Theoretical Analysis of the Series Resistance of a Solar Cell," Solid-State Electronics 

10,1967,765-775. 

13. RJ. Walters, S.R. Messenger, G.P. Summers, E.A. Burke, and C.J. Keavney, IEEE Trans. Nucl. 

Sci. N5-38, 1991, pp. 1153. 

14. N.M. Pearsal, N. Robson, H. Thomas, J.K. Luo, C.M. Hardingham, and T.A. Cross, "Annealing of 

Irradiation Damage in Epitaxial InP Homojunction Solar Cells," 23rd IEEE Photovoltaic Specialists 

Conference, May 1993, pp. 1442-1447. 

15. RJ. Walters, Ph.D. Dissertation (University of Maryland Baltimore County, Baltimore, MD, 1994). 

16. RJ. Walters, , G.P. Summe~ , M.L. Timmons, R Venkatasubramanian, J.A. Hancock, and J.S. 

Hills, "Annealing of Irradiated n p InP Buried Homojunctions," 13th Space Photovoltaic Research and 

Technology Conference, June 1994, this vol. 

top metal ZnS/MgF 2 AR Coating 

+ 

n -lnP emitter 

p-lnP base 

++ 

p -Gco.-47Ino.53As 

++ 

n -Gco .. UlnO.53As 

++ 

n -lnP + 

n -Gco . ..(7Ino.s3As emitter 

p-GaO.47Ino.S3As base 

p + -lnP BSr buffer layer 

p-lnP substrate 

top metal 

top cell 

tunnel 

junction 

bot.tom 

cell 

Figure 1: Schematic diagram of the InP/Gao47'no.53As tandem solar cell structure. 

163 

I 
~ 

N 

E 

o 

50 

" 40 f- 4: 

E 

Fluence 

Pre-red 

14 

1x1014 ....... - 3x 10 

-2 

(em ) 

15 ........ 1 x 10 

15 

-·_··3x10 16 

-··_·· 1x10 

~ 30 .. . ;:'."': .. ", ... -:."': .::"'-:. ... -: ... -; .. "': ... -: ... ": ... "': ... -:' ... '": .:: .. "': -~ -~:::: .::-": --= .. -:-: ... :.-..: - .... ............................................... -'. , (I) 

c 

CD 

o 

+- 20 

c 

CD 

~ 

~ 

:::J 

U 

CD 

U 

10 

-- , --"-.-.-.- " \ 

.. - .. -.. '-'-'-'-'-'-'-. '. _ .. - .. -.. - .. -.. - .. 

1 MeV Electrons 

1 Sun 

AMO 0 

T=25 C 

-.0_ .. _. '-

,..... . ...... \ 

, \ 00 . ... \ 

. . ., ~, 

\ :: I 

\ °

0 

~, 

'. \ '.: I 

'. \ : :', \ .. : '~, : \ : : I 

\ . .. !' 

\ : :' 

o 

0.0 0.5 1.0 

Cell Voltage (V) 

1.5 

Figure 2: The effects of 1 MeV electron irradiation on the InP/Gao.47Ino.53As tandem 

solar cell I-V curves. A kink is shown at low voltages. 

35 

30 

~ 

- 25 

-

w 

....--.-. - N 20 

E 

0 

" « 15 

E 

'-/ 

~10 ~ 

5 

o 

o 

V ---0-0 oc .. 

---0 :--:-:;;:: ---0 ~-~--- ---0 .. :--:-;;::: ~ ---~ ------ J sc ~ o~ 

-----0-0 ~~ 0 ---0 -0 

~ ~o~ ............... -L::,.____ --0 ~ 

... L::,.~ "" FF 

1.25 

1.00 

LL 0.75 LL 

----... L::,. 0 [ff% '--.. .. ~ ~ 

~ ... 

o 0.50 >0 

13 

3x10 

L::,. 

~ ... SHJ InP L::,. 

16 

10 

-2 

1 MeV Electron Fluenee (em ) 

0.25 

0.00 

10

17 

Figure 3: The gegradation of the cell response for the InP/Ga0.47In053As device shown 

in Fig. 2. The n p SHJ InP data was obtained from reference 15. 

164 
40 

30 

.......... 

N 

E 

0 

.......... 20 < 

E 

'-" 

"""") 

10 

o 

0.0 

1 MeV Electrons 

1 Sun, AMO 

15 -2 

3x10 cm 

Earlier 

Cell 

0.5 

Fluence 

Present 

1.0 

V (V) 

Cell 

1.5 2.0 

Figure 4: Comparison of the I-V curves for the tandem cell in this study with 

a tandem cell from a previous study (ref. 2). The present cell curve has been 

displaced by 0.25 V for clarity. 

50 

InP /GaO.47'no.S3As Tandem Cell 

1 MeV Electrons 

40 1 Sun, AMO 

v (V) 

Figure 5: Annealing of 1 MeV electron-irradiated I-V curves for the 

InP/Gao47'no.~s tandem solar cell shown in Fig. 2. 

165 
I 

L_ 

1.0 

LCD 

..-

CD 0.8 E 

0 

~ 

0 

Cl.. 0.6 

> 

Cl.. 

-0 0.4 CD 

N 

0 

E 

~ 

0.2 

0 

z 

0.0 

275 

, 

• ___ • __ • FF___ __ A __ A~. • __ __ .--.==1==.==.==* ~--~ 1 ".-".-".--".-". 

Voc __ e-e--'--e--e--e 

e __ e __ e--e--e--e , 

J , sc 

.-.--,-.-.-. .-- .-.-.--.--.-- , 

325 

Eft% , 

30 MIn Anneal , 16 - 2 I 500 K Anneal ~ =2.4x1 0 8 /em ,/ 

375 425 

Temperature (K) 

, 

475 0 30 60 90 

Time (min) 

Figure 6: The recover)' of the PV parameters of the InP/Gao47'no.s:f.s solar cell 

irradiated with 1 MeV electrons. Partial recover)' is seen up to 475 K. Only Voc 

continues to recover gradually above this temperature. 

166 
199 N95-20520 

RADIATION EFFECTS ON p+n InP JUNCTIONS GROWN BY MOCVD 

S.R. Messenger 

SFA, Inc. 

Landover, Maryland 

R.J. Walters 

U.S. Naval Research Laboratory 

Washington, DC 

M.J. Pan unto 

University of Maryland Baltimore County 

Baltimore, Maryland 

and 

G.P. Summers 

U.S. Naval Research Laboratory 

Washington, DC 

INTRODUCTION 

The superior radiation resistance of InP over other solar cell materials such as Si or GaAs has 

prompted the development of InP cells for space applications. The early research on radiation effects 

in InP was performed by Yamaguchi and co-workers who showed that, in diffused p-fnP junctions, 

radiation-induced defects were readily annealed both thermally and by injection, which was 

accompanied by significant cell recovery. 

More recent research efforts have been made using p-lnP grown by metalorganic chemical 

vapor deposition (MOCVD). While similar deep level transient spectroscopy (DL TS) results were found 

for radiation induced defects in these cells and in diffused junctions, significant differences existed in 

the annealing characteristics. After injection annealing at room temperature, Yamaguchi noticed an 

almost complete recovery of the photovoltaic parameters, while the MOCVD samples showed only 

minimal annealing. 

In searching for an explanation of the different annealing behavior of diffused junctions and 

those grown by MOCVD, several possibilities have been considered. One possibility is the difference in 

the emitter structure. The diffused junctions have S-doped graded emitters with widths cif -O.3lJm,while 

the MOCVD emitters are often doped with Si and have widths of -300A (0.03~). The difference in the 

emitter thickness can have important effects, -e.g. a larger fraction of the total photocurrent is generated 

in the n-type material for thicker emitters. Therefore the properties of the n-lnP material may explain 

the difference in the observed overall annealing behavior of the cells. 

EXPERIMENTAL 

In this study, n-lnP solar cells and mesa diodes were irradiated with either 1 MeV electrons or 3 

MeV protons, with both (current-voltage) IV and DL TS measurements being made. The IV 

measurements were performed both in the dark and under 1 sun AMO illuminations using an Oriel 

1000W portable Xe arc lamp simulator. A Kepco 50-2M bipolar amplifier, and two Keithley 617 

electrometers were employed in the IV measurements, which were computer controlled through a GPIB 

interface. DL TS and capacitance-voltage (CV) measurements were performed using the Bio-Rad 

DL4600 system which uses the Boonton 72-B capacitance meter. The cryostat used in the DL TS 

system has a temperature range of 8S-S00K which allows annealing experiments to be performed 

directly in the cryostat. The cryostat window is made of sapphire which allows for cell illumination as 

well. All cell measurements were thus performed in the cryostat. This unique system provides a 

simple and convenient means for collecting cell data. 

The samples were grown by Spire Corporation using Si as the n-type base dopant Figure 1 

gives a schematic description of the samples. The base dopant level was -3 x 1016 cm·

3 and that of 

the emitter was -2 x 1018 cm·

3

, both of which were determined from CV measurements. 

1 MeV electron irradiations were performed at a Van de Graaff accelerator at the NASA 

Goddard Space Flight Center in Greenbelt, MD. Incremental ftuences were performed up to _1018 

167 
e'Jcm2. The beam currents were law enaugh (-150 nA) to. avaid sample heating during the irradiatians. 

Dasimetry is knawn to. within 10-15%. The 3 MeV protan irradiatians were perfarmed at the Naval 

Surface Warfare Center in White Oak, MD an a Pelletron acceleratar. One irradiatian was perfarmed 

giving a tatal f1uence -3 x 10'2 p+Jcm2 (15-20% accurate). The pratan beam was rastered aver a 2.4 

cm2 area, and the current an the target was -14 nA/cm2

• 

RESULTS 

DL TS Measurements: 

Figure 2 shaws the DL TS spectrum measured on a p+ n InP mesa diade after 1 MeV electran 

irradiatian to. a fluence _10'8 cm'3. Majarity carrier traps EN1-4, and EB are farmed during the 

irradiatian. One minarity carrier trap, HN1, appears in small cancentratians and seems to be due to. a 

camposite defect The "EN" refers to. electran trapping in n-type InP. Similarly, "HN" refers to. hale 

trapping in n-lnP. 3 MeV pratan irradiatian praduces the same defect spectrum, except the relative 

peak heights are slightly different as shown in Figure 3. Trap EN1 is anly seen using rate windaws 

>1000/s. As can be seen in Figs. 2 and 3, EB is the daminant DLTS signal in bath cases. The 

activatian energy, capture cross sectian, and annealing characteristics af this peak are similar to. thase 

of the EB peak measured in irradiated p-type InP as shown by McKeever et al. in reference 1. No 

clear relationship between the remaining peaks and peaks measured in irradiated p-type InP seems to 

exist. The activatian energy of the EN2 defect is similar to that of the ED defect measured in irradiated 

p-type InP (-0.3 eV) as seen in reference 2, but the capture cross section of EN2 is about two orders 

of magnitude larger than that of ED, Also, the activation energy of HN1 is similar to that of the H4 

defect (0.34 eV at this dopant level) , but the capture cross sectian af HN1 is abaut an order of 

magnitude larger than that of H4. This may be due to. the fact that HN1 is most likely several 

overlapping peaks. Tables I and II characterize each defect according to its activation energy, capture 

cross sectian, defect cancentration, and introduction rate after 1 MeV electron and 3 MeV pratan 

irradiations, respectively. 

Intraduction rates were calculated for the 1 MeVelectran irradiatians and are included in Table 

I. The errors are -2-4%. Figure 4 shows DLTS data for the introduction of defects with 1 MeV electron 

irradiatio.n. The calculated introductian rate for EB of 0.15 cm" is much lower than the value of 0.7 

given for the dominant defect H4 in p-type InP by Sibille in reference 3 and also by Levinson et al. in 

reference 4 (who measured a value of 1.0). The total trap introduction rate is 0.2 cm" which is also 

much lower than that found for p-type InP as stated in references 3-5. The introductian rates for the 3 

MeV protons were obtained fram only one data point assuming a linear dependence. This is justified 

due to. the linear relatianships found in the 1 MeV electron irradiated samples. The total trap 

introduction rate here is 52,5 defects per incident proton per cm which compares nicely with the data 

shawn by Walters and Summers in reference 6. Overall, there are much fewer stable defects created 

in n-type InP than there are in p-type InP. This same result was also seen by Levinson et al. in 

references 4 and 7. 

Forward bias injection annealing at any temperature did not change the DL TS defect structure. 

This is in contrast to. the behavior of the H4 defect in irradiated p-type InP which injection anneals 

readily at room temperature. Isochronal thermal annealing was performed an the electron irradiated 

junctions and the results are shown on Figure 5. It is seen that the only defect which thermal anneals 

is EB. By a temperature of -4001<, EB is completely remaved, It should be mentioned that the 

annealing experiments were performed with the sample under a reverse bias of -2V. It was found that 

open circuit anneals led to an enhanced annealing rate. Experiments are currently under way dealing 

specifically with this effect. Thermal annealing both with and withaut an applied forward bias produces 

the same results implying that the annealing is not injection sensitive. There is therefare anly ane 

thermal annealing stage accurring -360-380K (far the case af V R =-2V) where the EB peak is 

campletely remaved. This happens to. be the same temperature where EB anneals in p-type InP as 

168 
discussed in reference 1. Isothermal annealing experiments on EB indicate a thermal activation energy 

to be 1.23 % 0.02 eV in a first order process. This is for the case of annealing with a bias voltage of -

2V. For zero bias, the thermal activation energy is 1.13 eV. These thermal energies are close to the 

value of 1.02 eV that Yamaguchi measured in p-type diffused InP1unctions for the H4 defect as shown 

in reference 8. Reference 9 contains isothermal annealing data on LEG n-lnP material. The 0.79 eV 

defect there has annealing behavior which is very similar to that of EB. First of all, it is insensitive to 

injection just as shown above. Also, its thermal activation energy is 0.98 eV which is also close to what 

was measured above. 

illuminated IV Measurements: 

Due to the similarity in the defect spectra under both electron and proton irradiation, the IV data 

were performed on the 3 MeV proton irradiated cells only. The first experiment measured the recovery 

of the celis due to illuminations below room temperature. The IV curves were taken at 86K because, in 

p-type InP, it has been found that illuminated IV measurements taken at 8SK do not cause injection 

annealing (ref. 10). The annealing experiment consisted of measuring the illuminated IV curve at 8SK 

before and after the 3 MeV proton irradiation. The cell was then warmed to 200K and illuminated, at 

short-circuit, for 35 minutes before being cooled to 86K where another illuminated IV measurement was 

made. This same procedure was performed for an illuminated anneal at 300K for 30 minutes. The 

measured IV curves are shown in Figure 6 and the photovoltaic (PV) parameters are given in Table III. 

The cell shows some recovery due to the illuminations. The largest increase is in the short circuit 

current (I..), but even this increase is small. The fill factor (FF) also increases slightly after the 200K 

illumination but not after the 300K illumination. The increase in these two parameters caused an 

increase in the maximum power (P moJ, but the recovery is inhibited because the open circuit voltage 

01 oJ shows no recovery. The next experiment was an anneal at room temperature. The same sample 

was then left in the dark at 300K for one week. As is seen in Figure 6, there is more recovery in Isc, 

but it is again only slight 

The same cell was then illuminated above room temperature. In these experiments, the 

illuminated IV curves could be measured at room temperature without altering the results. Figure 7 

includes the IV curves measured after 1 week in the dark to indicate the final state of the cell after the 

T s300K annealing stages. The corresponding PV data is given in Table IV. Illuminating the cell at 

350K for 4 hours caused significant recovery in all of the PV parameters. In particular, Voc is seen to 

recover for the first time. Subsequently illuminating the cell at 400K for 3 hours caused increases in all 

of the parameters. Two hours of illumination at 450K only caused an increase in V 00' but Isc was 

unaffected. 

The IV curve measured after the 450K illumination represents the maximum recovery of the 

cell. Significant recovery is evident, primarily due to the increase in Isc. However, while the recovery is 

substantial, it is far from complete. In particular, almost none of the radiation-induced degradation of 

the FF has been removed. Also, Voc shows only very little recovery. 

Dark IV Measurements: 

Measurements were also made of the dark IV characteristics. By fitting the measured dark IV 

data to the standard three-term diode equation, estimates of the diffusion term (101). the recombination 

term (loJ, and the shunt resistance term (R.tJ were determined. A commercially available curve fitting 

package, Peakfit, by Jandel Scientific was used. These fits were performed after each annealing step. 

The forward bias dark current characteristics of the same cell used above were measured at 

room temperature before and after irradiation, as well as after the successive annealing treatments 

169 
described above. The measured data were fit to the following equation : 

where 101' 102, and Rsn are the drtfusion, recombination, and shunt resistance fitting terms, respectively. 

In this equation, it was assumed that the energy level (EJ of the dominant recombination center is 

located at the intrinsic Fermi Level (E). Attempts were made to fit the data with Et as a free parameter, 

but the pre-irradiation recombination current in this particular cell was large enough that changes in ~ 

from E, had virtually no effect on the fit Some examples of the fits obtained are shown in Figure 8. 

The data for the fitting parameters are given in Table V for the annealing steps above. The fit to the 

pre-irradiation data does not seem to be very good. The reason for this is not known. Nevertheless, 

the fits to the post-irradiation and post-anneal data were excellent 

DISCUSSION 

These results with p+n InP grown by MOCVD will now be compared with those obtained using 

n+p InP fabricated both by thermal diffusion and by MOCVD. The last SPRAT conference contains the 

data for the diffused junctions from reference 11 . Figure 9 shows annealing data for all three of these 

cell types for the maximum power output. The scales are normalized to beginning-of-life (BOL). As 

seen on Fig. 9 the diffused junctions fully recover after a thermal anneal of 450K while both MOCVD 

cells do not recover even at 500K Moreover, the slow recovery of both MOCVD cell types appears to 

be the same indicating similar behavior for n- and p-type InP. This trend is also apparent in Figure 10 

where the same cell types are plotted this time for the normalized open circuit voltage. The fitted dark 

IV parameters 101 and 102 also follow this trend as can be seen on Figure 11 . Figure 11 plots the data 

from Table V on scales normalized to BOL the same slow recovery mentioned above occurs. It was 

also noticed at the last SPRAT conference from reference 12 that the radiation induced defects in p-lnP 

did not anneal until thermal anneals at 650K were performed. It might the case that the PV parameters 

also would anneal at such temperatures , but the experiments have not been performed successfully. 

CONCLUSION 

Defect introduction and annealing studies were performed on Spire-grown n+p InP solar cells 

and mesa diodes. The introduction rates of the defects formed were shown to be much lower than the 

p-type counterparts . The annealing characteristics were, however, the same. In comparing the 

MOCVD samples with the diffused junctions, the diffused junctions showed much better annealing 

characteristics . Full recovery of the PV parameters existed in the diffused junctions where only minimal 

recovery is seen in the MOCVD samples. Future studies will try to explain this important difference. 

170 
------- ----- ------

REFERENCES 

1. McKeever, S.W.S.; Walters, R.J.; Messenger, S.R.; and Summers, G.P.: Deep Level transient 

spectroscopy of irradiated p-type InP grown by metalorganic chemical vapor deposition. J. Appl. 

Phys., vol. 69, 1991, pp. 1435-1439. 

2. Walters, R.J.; Messenger, S.R.; Summers, G.P.; Burke, E.A.; and Keavney, C.J.: Proton and 

electron irradiation of MOCVD InP solar cells: Experimental results and radiation modelling. Proc. 

of the 22rd IEEE Photovoltaic Specialists Conference, October 7-11 , 1991, Las Vegas, NE, 

pp.1560-1565. 

3. Sibille, A. and Bourgoin, J.C.: Electron irradiation induced deep levels in p-lnP. Appl. Phys. Lett., 

vol. 41 , 1982, pp. 956-958. 

4. Levinson, M.; Temkin, H.; and Bonner, W.A.: Electron bombardment induced defect states in pInP. J. Electron. Mater., vol. 12, 1983, pp.423-432. 

5. Drevinsky, P.J.; Caefer, C.E.; and Keavney, C.J.: Processing- and radiation-produced defects in 

InP solar cells. Proc. of the 3rd IEEE International Conference on Indium Phosphide and Related 

Materials, Cardiff, Wales, April 8-11, 1991, pp. 56-59. 

6. Walters, R.J. and Summers, G.P.: Deep level transient spectroscopy study of proton irradiated ptype InP. J. Appl. Phys., vol. 69, 1991, pp. 6488-6494. 

7. Levinson, M.; Benton, J.L.; Temkin, H.; and Kimerling, L.C.: Defect states in electron bombarded 

n- InP. Appl. Phys. Lett., vol. 40, 1982, pp. 990-992. 

8. Yamaguchi, M. ; Ando, K; Yamamoto, A.; and Uemura, C.: Injection-enhanced annealing of InP 

solar-cell radiation damage. J. Appl Phys., vol. 58, 1985, pp. 568-574. 

9. Koyama, J.; Shirafuji, J.; and Inuishi, Y.: Annealing behaviour of gamma-ray-induced electron traps 

in LEC n-lnP. Electronics Letters, vol. 19, 1983, pp.609-10. 

10. Walters, R.J.:A study of the annealing of radiation-induced defects in InP solar cells. Doctoral 

dissertation, University of Maryland Baltimore County, May, 1994. 

11. Walters, R.J .; Summers, G. P.; and Bruening, J.: A detailed study of the photo-injection annealing 

of thermally diffused InP solar cells. Proc. of the 12th Space Photovoltaic Research and 

Technology Conference, NASA Lewis, Oct 20-22, 1992, pp. 1-7. 

12. Messenger, S.R.; Walters, R.J.; and Summers, G.P.: High temperature annealing of the minority 

carrier traps in irradiated MOCVD n+p InP solar cell junctions. Same as ref. 11, pp. 8-15. 

171 

j 
Table I DL TS defect parameters for 1 MeV electron irradiated n-lnP diodes. The concentrations 

shown are for a 1 MeV electron fluence of 1016 cm2 and a DLTS rate window of 2 S-1. 

Trap Activation Capture cross Concentration Introduction rate 

energy (eV) section (x 1 0-13 cm~ (x 1014 cm-~ (cm1) 

EN2 0.31 2.05 0.1 -

EN3 0.48 2.37 1.3 0.013 

EN4 0.55 0.287 1.8 0.018 

EB 0.76 15.6 14.5 0.22 

Table" DLTS defect parameters for 3 MeV proton irradiated n-InP. The f1uence was 3 x 1012 cm2

• 

The introduction rates were calculated assuming a linear dependence. The concentrations are for a 

rate window of 2/s. 

Trap Activation Capture cross Concentration Introduction rate 

energy (eV) section (x 10 13cm~ (x 1013cm3

) (cm-1) 

EN2 0.3 6.3 0.95 3.2 

EN3 0.43 9.5 3.2 11 

EN4 0.62 131 1. 4.7 

EB 0.74 5 8.6 29 

HN1 0.34 1.69 1.6 5.3 

172 
Table III Recovery of PV parameters of n-lnP solar cell under short circuit illuminations from 200K to 

400K IV curves measured at 86K Irradiation was 3 x 1012 3 MeV protons cm·

2

• 

Condition lac (rnA) Voc M Pmax (mW) FF Eff(%) 

Pre rad 21.88 1.270 29.90 0.9323 18.95 

Post rad 17.71 1.229 17.82 0.8186 13.03 

200K, 35 17.84 1.231 18.12 0.8257 13.26 

min. 

300K, 30 18.03 1.230 18.29 0.8245 13.38 

min. 

300K, 1 wk. 18.52 1.230 18.75 0.8230 13.71 

350K, 4 hr. 18.58 1.231 18.74 0.8192 13.71 

4001<, 1 hr. 18.76 1.230 19.14 0.8293 14.00 

400K, 3hr. 18.99 1.239 19.44 0.8259 14.22 

Table IV Cell recovery under illumination from 300K to 450K. IV curves measured at 298K. Irradiation 

was 3 x 1012 3 MeV protons cm·

2

• "Post rad" data was taken after 35 min. ilium. at 200K 

Condition Jsc (mA) Voc M Pmax (mW) FF Eff(%) 

Pre rad 25.13 0.8549 18.22 0.8480 13.33 

Post rad 19.55 0.7633 11 .21 0.7513 8.201 

300K, 30 19.58 0.7645 11.28 0.7537 8.251 

min. 

300K, 1 wk. 20.03 0.7654 11.54 0.7528 8.443 

350K, 15 20.57 0.7662 11.87 0.7533 8.682 

min. 

350K, 4 hr. 21.17 0.7731 12.40 0.7573 9.068 

400K, 1 hr. 21.43 0.7761 12.61 0.7583 9.223 

4001<, 3 hr. 21.79 0.7806 12.95 0.7614 9.475 

450K, 2 hr. 21.86 0.7875 13.10 0.7609 9.580 

173 

_I 
Table V Effect of illumination on the dark current parameters on n-lnP solar cells irradiated with 3 MeV 

protons to a fluence of 3 x 1012 cm2• Parameters are from fits of the dark IV data. 

Condition 101 (x 10-

16 Ncm'1 102 (x 10-9 Ncm'1 Rsh (x 105 Q) 

Pre rad 1.099 0.1717 4.390 

Post rad 9.451 3.090 3.665 

3501<, 15 min. 8.645 3.057 3.452 

350K, 1.75 hr. 8.507 2.736 3.149 

3501<, 4 hr. 7.229 2.766 2.678 

400K, 3 hr. 6.906 2.370 2.998 

450K, 2 hr. 5.603 2.205 3.385 

174 
Figure 1 

1 . 4 --.. 

I"') 

I E 1.2 

U 

o 1 . 0 

~ 0 .8 

o 

C 0 . 6 

CJ"l 

U) 

U) 0.4 

f--

-.J 

o 0 . 2 

-- ~- - - -

top metal (Ag/Pd) grid 

~ ----

I 

AR coating (ZnS/MgF2 ) I 

0.1 5)-l-m p·-lnP emitter (Zn-doped) 

3)-l-m n+-lnP base (Si-doped) 

n·-lnP BSF buffer loyer 

n+-lnP substrate (Si-doped) 

bottom metallization (Ag/Zn) 

A schematic drawing of the p+n InP/InP cells. 

No = 4.5 x 1 O'S cm- 3 

T = 2 S-1 

VR = -2 V 

VF=-0.3V 

PW = 500 mS 

1 MeVelectrons 

(IP = 10'6 e-/cm2 ) 

EN3 

EN2 0.0 L-~~ ________ ~~~~ __ ~~~ __ ~~~~~~ __ ~~~~ 

100 120 140 160 180 200 220 240 260 280 300 

Temperature (K) 

Figure 2 DLTS spectrum obtained after 1 MeV electron irradiation 

to a fluence of lE16 cm-2 • The dominant peak, EB, is identified 

with the EB peak found in irradiated p-type InP. 

175 
200 

...--.. 

ll... 

~100 o 

c 

Q'l 

U) 

p+n epitaxial InP 

3 MeV proton irrodiation 

1000/s 

EN3 

I 

/ 

?\ 2 s; ~ . 

.. . ; \ 

tE8\ 

: ~ . . . . ~ : . . ;; . . 

~ o~==~~~--~~----~------------~--~ ~ o 

-100 

Nd=4x1 016 C m-3 

if?=3x1 012 c m-2 

1 O/s, after thermal 

anneal of E8 

-200~~~~~~~~~~~~~~~~--~~~~ 100 120 140 160 180 200 220 240 260 280 300 

Temperature (K) 

Figure 3 DLTS spectra obtained after 3 MeV proton irradiation to 

a fluence of 3xl012 cm-2 • The HNl peak is most likely two 

overlapping peaks. 

1.4 

....--.... 

7 1.2 

E 

u 

1.0 

o 

X 0 .8 

"----"" 

o 0.6 

C 

O"'l 

(.f) 0.4 

(.f) 

r----.J 0.2 

o 

Introduction of Defects into p/n InP 

*1 MeV electron fluences 

160 

EN3 EN4 

180 200 220 240 260 

Temperature (K) 

280 300 

Figure 4 Introduction of defects resulting from 1 MeV electron 

irradiation. A linear dependence is found for all defects as 

shown in the inset. 

176 

_~J 
...-- 3 . 6 ,.,., 

I E 3.2 

U 

..,. 2 . 8 

2.4 

x 

-- 2 .0 

o 

C 1 . 6 

Q") 

U) 1 . 2 

U) 

~ 0 . 8 

---1 

o 0.4 

*Isochronal thermal anneals 

(20 minutes each at V R =-2V) 340K 

o~~~~~----~~~~~~~~~~--~~~~--~~ 

110 130 150 170 190 210 230 250 270 290 

Temperature (K) 

Figure 5 Isochronal annealing experiment. EB is the only defect 

whi ch anneals. Holding the sample at VR=-2V during the anneals, 

the annealing stage occurs -380K. 

24 

22 

20 Pre rod 

1 8 

-16 « 

E14 

--~ 1 2 

c 300K, 

300K, 

30 min 

1 week in the dark ~ OJ 1 0 \..... 200K, 35 min \..... 

::1 8 Post rod 

U 6 

4 p+n Epitaxial InP /lnP 1 sun, AMO 

2 3 MeV proton irradiation 86K c;t'=3x1012cm- 2 

o 

0 .0 0 . 1 0 . 2 0 .3 0 .4 0.5 0.6 0 . 7 0 .8 0.9 1 .0 1.1 1.2 1 .3 

Voltage (V) 

Figure 6 Photo-injection annealing of a 3 MeV proton irradiated 

n-InP cell. Illuminating the cell short circuited at 200K and 

then 300K caused a small increase in Isc and FF, but Voc did not 

change. The IV curves were taken at 86K. 

177 
30 r---~----~----~----~----~--~~--~-----r----~----' 

p+n Epitaxial InP /lnP 

3 MeV proton irradiation 

25 r-----------------------------------_ <p=3x1012cm-2 

-:;: 20 nFF=t===1==:f=====::;;;;::: E 

-+-' 1 5 

c: 

OJ 

l.-

I.-

~10 

U 

5 

450K, 2 hr (voe ) 

400K, 3 hr 

350K, 4 hr 

week in the dark 

Post rad 

1 sun, AMO 

298K 

o ~ __ ~ ____ ~ ____ -L ____ ~ ____ ~ __ ~~ __ ~ __ -u~ __ ~~ __ ~ 

0.0 0 . 1 0 . 2 0 .3 0.4 0 . 5 0 . 6 0.7 0.8 0 .9 1.C 

Voltage (V) 

Figure 7 Annealing of the same cell above 300K. The IV curves 

were measured at 298K. All parameters recovered under 

illumination at 350K and 400K, but only Voc showed more recovery 

at 450K. 

1 0- 2 

10- 3 

10-4 

...---.... 

<C 

'----'"' 10-5 -+-' 

c: 

Q) 10-6 L 

L 

:::J 

U 10-7 

10-8 

10-9 

p+n epitaxial InP /lnP 

3 MeV proton irradiation 

(<;0 = 3 x 1 012cm -2) 

o Po s t rod -----------.,f37 

v after 450K. 2 hr 

ill u m ination -------iJIi'fIi""' 

*symbols are measured data 

*Iines are fits to the data 

298K 

0.0 0 . 1 0.2 0.3 0.4 9.v. 5) 

Voltage \. 

0.6 0.7 0.8 

Figure 8 Dark IV data measured on the same cell. The 

irradiation increased both the recombination and generation 

currents, and the annealing caused only a small reduction. 

178 
, 

I 

\ 

\ 

, 

) 

, .0 

0 .9 

0.8 

~ :::::E 0 .7 

0... 

-0 0.6 

<l) 

N 0.5 

0 

E 0.4 

I0 0 .3 

z: 

0 .2 

0 .1 

0 .0 C1.> 1J) L-

• 

• • 

- - -- - - - - -

n+p OJ (1 MeV electrons, 1 E1 6 e-/cm2~ 

n+p SHJ (3 MeV protons, 1 E1 3 p+/cm2 

p+n SHJ (3 MeV protons, 3E1 2 p+ /cm2 

a.. o 

0.. 

275 325 375 425 475 

Anneal Temperature 

Figure 9 

junctions 

in t he DJ 

recove r ed 

, .0 

0.9 

0 .8 

g 0 .7 

> 

-0 0 .6 

<l) 

N 0 .5 

0 

E 0 .4 

I0 0 .3 

z: 

0 . 2 

0 .1 

0 .0 C1.> 

La.. 

Comparison of Pmax (normalized to BOL) between diffused 

(DJ), and both p-type and n-type MOCVD junctions. Pmax 

c ells recovered substantially while the MOCVD ones only 

minimally. 

-+-' 

(/) 

o 

a.. 

• 

• • 

n+p OJ (1 MeV electrons, 1 El 6 e-/ cm2~ 

n+p SHJ (3 MeV protons, 1 E1 3 p+ /cm2 

p+n SHJ (3 MeV protons, 3E1 2 p+ /cm2 

275 325 375 425 475 

Anneal Temperature 

Figure 10 Same as Fig. 9 for normalized Voc. The same trend 

exists for Voc as for Pmax above. 

179 

c-3 
---.J 20 

0 

en 18 

0 1 6 102 ....... 

""'0 

<l> 

N 

1 4 

0 1 2 *Dark current parameters determined 

E from fits of dark IV data 

~ 10 

0 

c: 8 

(f) 

101 ~ 

<l> 6 ....... 

<l> 

E 4 

0 

~ 2 1 sun, AMO illuminated anneals 

0 Annealing step D- O 

prerod postrod 350K 350K 350K 400K 450K 

15 min 105 min 4 hr 3 hr 2 hr 

Figure 11 Diffusion (101) and recombination (102) terms plotted 

normalized to the BOL to show relative changes. The increase in 

102 dominates, but both terms show a decrease under illumination. 

180 
1

~ 

I 199?/I¥/05 3 9 ???? 

N95-20521 

DESIGN OF HIGH-EFFICIENCY, RADIATION-HARD, GalnP/GaAs SOLAR CELLS1 

Sarah R. Kurtz, K.A. Bertness, A.E. Kibbler, C. Kramer, and J.M. Olson 

National Renewable Energy Laboratory 

Golden, Colorado 

SUMMARY 

Record air mass zero efficiency values are reported for GaO.5InO.5 P/GaAs devices before and 

after irradiation by 1015 cm2 1 MeV electrons. The two-terminal, two-junction devices are grown 

monolithically with a high-conductance, GaAs tunnel-junction interconnect and an area of 0.25 cm2. A 

device optimized for beginning-of-life (BOl) performance achieved BOl 25.7% (25.4%) efficiency, while 

devices optimized for end-of-life (EOL) performance achieved EOl efficiencies of 19.6% (19.8% and 

20.0%). (The efficiencies noted in parentheses were measured at NASA lewis) The effects of the 

thickness of the top cell and the doping level of the bottom cell were investigated in this study. A range of 

top-cell thicknesses and bottom-cell doping levels gave respectably high (greater than 18%) EOl 

efficiencies. 

INTRODUCTION 

In recent years, GaO.5Ino.5P/GaAs cells have drawn increased attention both because of their 

high efficiencies and because they are well suited for space applications. They can be grown and 

processed as two-junction devices with roughly twice the voltage and half the current of GaAs cells. They 

have low temperature coefficients, and have good potential for radiation hardness. We have previously 

reported the effects of electron irradiation on test cells which were not optimally designed for space. (ref. 

1) From those results we estimated that an optimally designed cell could achieve 20% after irradiation 

with 10 15 cm2 1 MeV electrons. Modeling studies predicted that slightly higher efficiencies may be 

achievable. (ref. 2) Record efficiencies for EOl performance of other types of cells are significantly lower. 

Even the best Si(ref. 3) and InP(ref. 4) cells have BOl efficiencies lower than the EOl efficiency we report 

here. Good GaAs cells have an EOl efficiency of 16%.(ref. 5) The InP/GaO.5InO.5As two-junction, twoterminal device has a BOl efficiency as high as 22.2% (private communication from M. Wanlass). but 

radiation results for these cells were limited. (ref. 6) 

In this study we use the previous modeling and irradiation results to design a set of 

Gao.5InO.5 P/GaAs cells that will demonstrate the importance of the design parameters and result in highefficiency devices. We report record AMO efficiencies: a BOl efficiency of 25.7% for a device optimized 

for BOl performance and two of different designs with EOl efficiencies of 19.6% (at 1015 cm-2 1 MeV 

electrons). We vary the bottom-cell base doping and the top-cell thickness to show the effects of these 

two important design parameters. We get an unexpected result indicating that the dopant added to the 

bottom-cell base also increases the degradation of the top cell. 

EXPERIMENTAL DETAilS 

A schematic of the device structure is shown in Fig. 1. The devices were grown from trimethyl 

gallium, trimethyl indium, trim ethyl aluminum, arsine, and phosphine in a hydrogen carrier gas. The 

dopant sources were diethyl zinc, hydrogen selenide, disilane, and carbon tetrachloride. Zinc and 

selenium were the p- and n-type dopant sources unless otherwise noted. The bottom-cell base doping 

levels were 1,3, and 8 X 1016 cm-3 , referred to hereafter as low, medium, and high, respectively. These 

doping levels could not be measured directly on the finished devices, but are estimated from calibration 

layers grown with similar diethyl zinc fluxes. Other details of the device structure and processing can be 

found elsewhere.(ref. 7, 8) 

1 Work funded by DOE contract No. DE-AC02-83CH 10093 

181 

______________________________________ J 
I 

I 

I 

I 

GRID 

~ ~ Si 

------"""" doping spike I 

0.5 GaAs n'" 6 x 1018 

0.025 AlinP n'" 4 x 1017 [Si] 

0.1 GainP n= 2x 1018 

0.5 GainP p= 1.5 x 1017 

(1.86 eV) 

0.05 GainP p", 3x10 18 

(1 .88 eV) 

0.011 GaAs P'" 8 x 1019 [C 1 

0.011 GaAs n'" 1 x 1019 

0.1 GainP n",, 1x1018 

0.1 GaAs n",,1x10 18 

3.5 GaAs p= 8X10 16 . 

0.07 GainP p= 3X10 17 

0.2 GaAs p",, 3X10 17 

substrate GaAs Zn-doped 

Thickness Doping 

(/lm) (cm -3 ) 

(window) 

TOP 

CaL 

(BSF) 

] TJ 

BOTIOM 

CELl. 

(BSF) 

Fig. 1. Device structure for the cell with high bottom-cell base doping, 0.65 J!m-thick top cell and a BOL 

efficiency of 25.7%. The total top-cell thickness of 0.65 I!m was varied by decreasing the thickness of the 

top-cell base layer (layer with 1.5 X 1017 cm3 doping). The BSF layers serve to passivate the back 

surface of each individual cell and the T J is a tunnel junction that makes an ohmic connection between 

the top and bottom cell. 

All of the devices were measured before and after irradiation on the fiber-optics, two-source 

simulator in B. Sopori's lab at NREL. The efficiencies were measured by adjusting the simulator to obtain 

the correct currents on two (top and bottom) reference cells. The top- and bottom-cell photocurrents were 

measured by shining a NIR or visible laser, respectively, on the device in addition to the simulator light. 

The spectral responses of the top and bottom cells were measured using red and blue bias lights, 

respectively. The record efficiencies reported in Table I were measured by K. Emery and coworkers at 

NREL, then sent to NASA Lewis for confirmation. In most cases the efficiencies agree within 2% 

(relative). A" of the cells are small: 5 mm X 5 mm. The cells are close to champion quality except that 

the anti-reflection (AR) coats were not well controlled and some variation was observed in the window 

layer of the top cell. EOl efficiencies of more than 20% would have been achieved if the AR coats had 

been optimal. The electron irradiation was done at JPl by Bruce Anspaugh and his staff. 

RESULTS 

A summary of the highest efficiency measurements is shown in Table I. A complete summary of 

all of the measured efficiencies, before and after irradiation, is shown in Fig. 2. Most of the data paints in 

182 

L-----------.------___________________________ J 
Table I. Summary of measurements on highest efficiency devices. All measurements were completed 

using the AMO spectrum and a cell temperature of 25°C. 

Cell design Irradiation Measurement 

Sase doping Top-cell thickness (eledcm2) place 

high 0.65 J.UT1 None NREL 

None NASA 

medium 0.45 J.UT1 1015 NREL 

1015 NASA 

ow 0.55~m 1015 NREL 

1015 NASA 

26~----~----~----~----~----~ 

24 

>-

g 20 

a.> 'u 

::= 18 w 

16 

base doping 

-+- low ___ medium 

.....- high 

.. - ............ -- " ... -- " -- EOL 

14~----~----~----~----~--~ 

0.2 0.3 0.4 0.5 0.6 0.7 

Top-cell thickness (11m) 

Voc Jsc FF Efficiency 

(V) (mAcm-2) (%) (%) 

2.393 16.55 88.7 25.7 

2.398 16.39 88.2 25.4 

2.221 14.53 82.9 19.6 

2.226 14.58 83.3 19.8 

2.198 14.72 83.0 19.6 

2.198 14.90 83.4 20.0 

Fig. 2. The AMO efficiencies of the devices before and after irradiation, as measured at NREL on the 

Sopori simulator. 

Figs. 2-4 represent the averaged values for four 5 mm X 5 mm devices. Data is not included for a few 

cells that were badly shunted or damaged. The bottom-cell base doping had little effect on the SOL 

efficiency. The top-cell thickness has a very large effect on the efficiency because the thinner top cells 

generate less photocurrent and the device is limited by this smaller photocurrent. After irradiation, the 

cells with lower bottom-cell base doping tend to show higher efficiencies. The optimal top-cell thickness 

decreases after irradiation. This is because the current of the thick bottom cell usually degrades more 

than the current of the thin top cell. Fig. 2 shows that respectably high efficiencies (greater than 18%) are 

obtained for top-cell thicknesses between 0.45 and 0.65 ~m when the bottom-cell base doping is not too 

high. 

The degradation of the photocurrents is shown in Figs. 3 and 4. The as-grown top-cell 

photocurrents show a very strong dependence on thickness, as expected. The device with top-cell 

thickness of 0.55 ~m has a lower photocurrent primarily because of a poor blue response, implying that 

some oxygen or carbon may have contaminated the window layer. After irradiation, the top-cell 

photocurrent shows a very significant dependence on bottom-cell base doping. This effect will be 

discussed below in more detail. The bottom-cell photocurrent decreases w~h top-cell thickness since a 

thinner top cell allows more light to penetrate to the bottom-cell junction. The larger decrease in 

photocurrent with higher base doping was expected from previous studies that showed the damage 

coefficient to increase w~h doping. The degradations of the Voc and the FF ranged from 6%-9% and 

2%-4%, respectively. The Voc showed a slightly greater degradation (8%-9%) for the cells with the low 

bottom-cell base doping compared with those with high doping (6%-7%). This difference is not great 

183 
enough to compensate for the opposite trend in bottom-cell photocurrent degradation (10%-11% 

degradation for the lightly doped and 21 %-23% for the highly doped cells). 

In order to better understand the degradation of the photocurrents we plot the spectral response 

of the bottom and top cells in Figs. 5 and 6. The primary loss mechanism in both cases is a decreased 

minority-carrier diffusion length. The increase in damage coefficient with base doping for the GaAs 

bottom cells was reported previously. (ref. 1, 9) Similar changes in the degradation of the spectral 

response curve as a function of base doping have also been reported for InP n-on-p cells.(ref. 4) The 

increased radiation hardness of the Gao.slno.sP top cell for low bottom-cell base doping is unexpected 

because no deliberate change was made in the top-cell doping. The similarities between the degradation 

of the top- and bottom-cell spectral responses imply that a memory or diffusion effect caused an 

unintentional change in the top-cell base doping. We are currently trying to confirm this hypothesis by 

using secondary mass ion spectroscopy to quantify the zinc levels in the top cells. 

18~--~----~----~-----r----~ 

(\.\--17 

E 

() 

< 16 

E 

--C 15 

Q) 

.... 

:; 14 

() 

~ 13 I 

a. 

~ 12 

base doping 

..... Iow 

.... medium 

-+- high 

11 ~ ____ L-____ ~ ____ ~ __ ~~ __ ~ 

0.2 0.3 0.4 0.5 0.6 

Top-cell thickness (Jl.m) 

0.7 

Fig. 3 Top-cell photocurrent before and after irradiation. 

18r---~~---'--~~-----r----~ 

(\.\ --

E17 

() 

~ 16 

--

~ 15 .... 

.... 

:::J ()14 

Q) 

<;'13 

E 

o 

~ 12 

m 

base doping 

• low 

• medium 

.. high 

.. ..... ..... ..... 

....... - .... .",."'- -- -- k- - EOL 

11 ~----~----~----~----~----~ 

0.2 0.3 0.4 0.5 0.6 

Top-cell thickness (Jl.m) 

Fig. 4 Bottom-ce" photocurrent before and after irradiation. 

184 

0.7 
--- - - --

The results of this study are very consistent with our previous study (ref. 1) in which we predicted 

an EOl efficiency of 20% for an EOl optimized device. However, our previous study differed in one 

significant way: the photocurrent for a 0.75 !-Jm-thick top cell degraded by only 2%, compared with 11 % for 

the 0.65 J.lm-thick top cell with low bottom-cell base doping in this study. The results of both studies 

together may imply that the base doping of the 0.75 J.lm top cell was lower than that used in this study. 

Thus, if we had used a lower top-cell base doping in this study, we should have seen improved radiation 

resistance of the top cell, and, hence, of the tandem cell. A decreased doping of the top cell might 

increase the degradation of the Voc . However, this is a small effect, and can be viewed as negligible 

since our previous study (with apparent low top-cell base doping) gave almost identical degradation of the 

Voc compared with this report. Thus, we conclude that the device can be even further optimized for EOl 

efficiency. 

1.0 

Q) 0.8 

en 

c 

0 

0.. 0.6 en 

Q) 

~ 

cu 0.4 ~ 

-<.> 

Q) 

0.. 

CJ) 0.2 

0.0 

1.4 1.6 1.8 

Bottom-cell 

base doping 

-low 

- - - medium - - high 

2.0 2.2 

Photon energy (eV) 

2.4 

Fig. 5. The bottom-cell spectral response of tandem cells with 0.65 J.lm-thick top cells before and after 

irradiation. 

1.0 

Q) 0.8 

en 

c 

0 

0.. 0.6 en 

Q) 

~ 

cu 0.4 ~ Bottom-cell t5 

Q) base doping 0.. 

CJ) 0.2 -low 

- - - medium 

- - high 

o. 0 E....----'-_-'--_'-----'-_~___''---~_...&...._I 

1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 3.4 

Photon energy (eV) 

Fig. 6. The top-cell spectral response of tandem cells with 0.65 J.lm-thick top cells before and after 

irradiation. The small variations in the blue response may be caused by contaminates in the window layer 

of the top cell. 

185 
L 

REFERENCES 

1. Kurtz, S.R., J.M. Olson, K.A. Bertness, D.J. Friedman, A. Kibbler, B.T. Cavicchi, and D.O. Krut: 

Radiation hardness of Gao.slno.s P/GaAs tandem solar cells. Space Photovoltaic Research and 

Technology, 1991 , p. 40-1-40-7. 

2. Cavicchi, B.T., D.O. Krut, D.R. Lillington, S.R. Kurtz, and J.M. Olson: The design and evaluation of 

dual-junction GalnP2/GaAs solar cells for space applications. 22nd IEEE Photovoltaic Specialists 

Conference, 1991 , p. 63-67. 

3. Garboushian, V., S. Yoon, and J. Turner: Radiation hardened high efficiency silicon space solar cell. 

23rd IEEE Photovoltaic Specialists Conference, 1993, p. 1358-1362. 

4. Keavney, C.J., R.J. Walters, and P.J. Drevinsky: Optimizing the Radiation Resistance of InP Solar 

Cells - Effect of Dopant Density and Cell Thickness. J Appl Phys, vol. 73, 1993, p. 60-70. 

5. Bertness, K.A., M.L. Ristow, M.E. Klausmeier-Brown, M. Grounner, M.S. Kuryla, and J.G. Werthen: 

16%-efficient GaAs solar cell after 101S cm-2, 1-MeV radiation. 21st IEEE Photovoltaic Specialists 

Conference, 1990, p. 1231-1234. 

6. Walters, R.J. , S.R. Messenger, M.W. Wanlass, and G.P. Summers: 1 MeV electron irradiation of 

monolithic, two-terminallnP/GaO.47Ino.S3As solar cells. 23rd IEEE Photovoltaic Specialists Conference, 

1993, p. 1475-1478. 

7. Olson, J.M., S.R. Kurtz, A.E. Kibbler, and P. Faine: A 27.3% efficient Gao.slno.sP/GaAs tandem solar 

cell. Appl Phys Lett, vol. 56, 1990, p. 623-625. 

8. Olson, J.M., S.R. Kurtz, A.E. Kibbler, K.A. Bertness, and D.J. Friedman: GalnP2/GaAs Tandem cells 

for space applications. Space Photovoltaic Research and Technology, 1991, p. 12-1-12-5. 

9. Bertness, K.A., B.T. Cavicchi, S.R. Kurtz, J.M. Olson, A.E. Kibbler, and C. Kramer: Effect of base 

doping on radiation damage in GaAs single-junction solar cells. 22nd IEEE Photovoltaic Specialists 

Conference, 1991, p. 1582-1587. 

186 

l 
N9S-20S22 

INVESTIGATION OF THE STABILITY AND 1.0 MeV PROTON RADIATION RESISTANCE 

OF COMMERCIALLY PRODUCED HYDROGENATED AMORPHOUS 

SILICON ALLOY SOLAR CELLS1 

Kenneth R. Lord II, Michael R. Walters, 

and James R. Woodyard 

Wayne State University 

Detroit, Michigan 

ABSTRACT 

The radiation resistance of commercial solar cells fabricated from hydrogenated amorphous silicon 

alloys is reported. A number of different device ~ctures were irradiated with 1.0 MeV protons. The cells were 

insensitive to proton f1uences below 1 E 12 cm The parameters of the irradiated cells were restored with 

annealing at 200°C. The annealing time was depend~t on proton f1uence. Annepling devices for one hour 

restores cell p~rameters for f1uences below 1E14 cm f1uences above 1E14 cm require longer annealing 

times. A parametric fitting model was used to characterize current mechanisms observed in dark I-V 

measurements. The current mechanisms were explored with irradiation f1uence, and voltage and light soaking 

times. The thermal generation current density and quality factor increased with proton fluence. Device 

simulation shows the degradation in cell characteristics may be explained by the reduction of the electric field 

in the intrinsic layer. 

INTRODUCTION 

The goal of our research is to develop a model to predict EOl performance of thin-film solar cells in 

space environments. In order to achieve the goal we have irradiated and characterized a number of different 

thin-film solar cell device structures fabricated from hydrogenated amorphous silicon alloys (refs. 1-5). The 

radiation resistance of single, d~1 and triple junction cells has been determined for 1.0 MeV proton f1uences 

in the 1 E11 through 1 E15 cm range. Both substrate and superstrate cell structures obtained from three 

companies producing commercial modules have been investigated. The J>-i-n layers in the cells are made up 

of hydrogenated amorphous silicon alloys deposited using plasma enhanced chemical vapor deposition; the 

details of cell structures have 

been reported in references 

one through five. Cells were 

characterized using light I-V, 

dark I-V and quantum efficiency measurements. 

Six cell structures have 

been investigated by our group. 

The results of 1.0 MeV proton 

irradiation on cell normalized 

power density are shown in 

Figure 1. Measurements on 

single and triple-junction cells 

with superstrate structures are 

shown. The triple-junction 

cells had two band gaps. The 

i-Iayers of the top two junctions 

were a-Si:H with the same 

band gap; the i-layer of the 

bottom junction was a-Si,Ge:H 

with a lower band gap than the 

top two junctions. Tandem and 

Effect of 1.0 MeV Proton Irradiation on Solar Cells 

1.00 ~~~I8iiti=:-::::""""-'-=:::==:-:~1 ~ Superstrate Cells 

~ Cl Single, a-SI:H 

~ 0.75 Triple, 2 Gap 

.. 

; 

0 

Q. 0.50 

't' Substrate Cells 

" '" Tandem, 2 Gap .t::! 

Ii 0.25 • Tandem,1 Gap 

e v Single, a-SI,Ge:H 

0 6- Single, a-SI:H Z 

0.00 '--....A...-................... ""'----'--'-' ................... ---'---'-.................... ~..4_...< ...................... oI..&.....:o....J 

1 E 11 1E12 1E13 1E14 1E15 

Proton Fluence (cm·2) 

Figure 1: Plot comparing the effect of 1.0 MeV proton irradiation on the 

normalized power density of a-Si:H alloy based solar cells. 

1 Portions of this work were supported by NASA, the Spacecraft Technology Division of TRW. and the 

Wayne State University Institute for Manufacturing Research. 

187 

) 

___ J 
single-junction cells with substrate structures are also shown in Figure 1; tandems with single and dual gaps 

were investigated as well as single-junction cells with a-Si:H and a-Si,Ge:H i-Iayers. The cell efficiencies ranged 

between seven and eight percent under AM1.5 global irradiance; the radiation resistance of the cells has been 

reported in prior papers published by our group (refs. 1-5). Figure 1 shows 1.0 MeV proton irradiation dewades 

the normalized power de~ of all the cells by less than a few percent for fluences less than 1 E12 em For 

fluences above 1 E15 cm the cell power degrades to I~~ than 10% of the initial power. Reductions in the 

normalized power density in the 1E12 through 1E14 cm proton fluence range depends on cell structure. 

Cells with triple junctions have the best radiation resistance, dual-junction cells are next and single-junction cells 

have the lowest radiation resistance. Single-junction cells 

with a-Si,Ge:H i-Iayers have 

better radiation resistance than 

cells with a-Si:H i-Iayers; it is 

not clear whether the effect is 

due to the role of germanium 

in these cells or the differences in the thicknesses of 

the i-Iayers. The i-layers of the 

a-Si,Ge:H single-junction cells 

are thinner than the i-Iayers of 

the a-Si:H cells. 

The effect of post-irra 

diation annealing at 200°C on 

the normalized power density 

of twenty-one solar cells fabricated from hydrogenated 

amorphous silicon alloys is 

illustrated In Figure 2. The 

data are for single-junction 

cells with a superstrate structure and 500 nm i-Iayer thickness. Three cells were irradiated at each f1uence; the 

power densroes following irradiation are shown by the open 

squares. The average power 

density of three irradiated cells 

at each f1uence is shown by 

the filled squares. The power 

densroes following a two-hour 

anneal at 200°C are shown by 

the open triangles with the averages represented by the filled triangles. Annealing the 

cells for one hour restores the 

normalized power density of 

cells irradiated with 1.0 MeV 

pro!~n nuences less than 1 E 14 

cm_2 f1uences above 1E14 

cm require longer annealing 

timas to restore the normalized 

power density. This is shown 

by the data in Figure 3. The 

average power density of the 

three single-junction cells irradiated with a f1uence of 1.5E15 

N 

-

Irradiation Data for Single-Junction a-Si:H Cells 

E u 8.00 

~ E 

- 6.00 

~ til 

c: 4.00 

~ '-; 

o 

D. 

1.022 MeV Protons 

Post- Irradiation 

2.00 

'Y • Averages 

o .00 L...---'-................................. _-'-...................... u..u.._..Io...-"-'-.......... .u..L. ....................... ~ 

1E12 1E13 1E14 1E15 1E16 

Proton Fluence (cm o2 ) 

Figure 2: Power densities of twenty-one cells a-Si:H single-junction cells 

measured after 1.022 MeV proton irradiation and following a postirradiation anneal at 200°C for two hours. 

N 

-

E 

u 

~ E 

-

~ 'i 

c: 

~ '-

; 

o 

D. 

8.00 

6.00 

4.00 

2.00 

Anneal Data for Single-Junction a-Si:H Cells 

1.022 MeV Protons 

• • • 

Ann .. llng conditions: 2hrs. at 200 oc 

• 

• 

Pre-Irradiation and following first anneal 

Post-lrradlatlon and following second anneal 

Post-Irradiation and following third anneal 

0.00 L....-............................................ _ ......................... .A.J...... ............................... ...........JL.-.......L-..................... ..uJ 

1E12 1E13 1E14 1E15 1E16 

Proton Fluence (cm-2) 

Figure 3: Average power densities following the first pre-irradiation 

anneal, and the second and third post-irradiation anneals of singlejunction cells irradiated with 1.022 MeV protons. 

188 
--- ---

cm2 recovers to 6.2 and 7.0 mW/cm2 with 2.0 and 4.0 hours annealing, respectively. The details of the time 

dependence of annealing have been reported in reference 6. In general, the higher the f1uence for MeV 

protons, the longer the annealing time to restore the normalized power density. -2 

The cross-over of the dual and triple-junction curves in Figure 1 at a f1uence of 1 E 14 cm is believed 

to be due the difference in the irradiances of the simulators used to measure the light I-V characteristics of the 

cells. The dual-junction cells were measured with a simulator which was optimized to match an AM1.S global 

spectrum. i he triple-junction cells were measured with our simulator which is an ORC model SS1000. The 

spectral irradiance of the ORC simulator contains xenon lines above 800 nm and is deficient in the red. The 

thicknesses of the i-Iayers in the triple-junction cells were matched to an AM1.5 global spectrum. The 

mismatch between the ORC simulator spectrum and the triple-junction cells probably resulted in the bottom 

junction limiting the cell current. We have modified the ORC simulator by adding a cold mirror and tungstenhalogen lamp (ref. 7). We plan to repeat the triple-junction measurements with the dual-source simulator to 

determine the reason for the cross-over of the dual-and triple-junction curves in Figure 1. 

The investigations summarized in Figure 1 must be extended to a range of proton and electron 

energies. It is our expectation that the measurements will provide the basic parameters for the development 

of a predictive model for determining the EOl performance of cells fabricated from hydrogenated amorphous 

silicon ba:.ed alloys in a variety of space environments. We plan to apply the techniques learned from our 

investigations with hydrogenated amorphous silicon based alloys to other thin-film solar cells of interest for 

space-power generation. 

PARAMETRIC FITTING MODEL 

The first step in developing a predictive model for EOl performance is the determination of parameters 

from measurements which can be related to basic material properties of solar cells. We have elected to 

develop a parametric fitting model to characterize current mechanisms in single-junction cells. Single-junction 

cells were chosen because triple-junction cells are far more complex in structure. Dark I-V characteristics were 

selected for the initial modelling investigations because they showed the largest changes in parameters with 

1.0 MeV proton irradiation. 

Determination of solar cell parameters from measured dark I-V characteristics requires curve fitting 

a parametric model to measured dark I-V characteristics. The objective of curve fitting is to minimize the 

differences between measured and calculated dark I-V characteristics. The parametric model which we 

developed includes current mechanisms which are characterized by a sum of analytical functions with 

parameters. The parameters in the analytical functions are referred to as fitting parameters; they are varied 

to fit a calculated I-V characteristic to a measured dark I-V characteristic. The model was used for analysis of 

dark I-V characteristics in the forward-bias region. The parametric model used for curve fitting is: 

where 

I = I [exp( q(V-IXR..»_1] + (V-lxR..) + a(V-lxR..)m 

o nkT Rill 

'0 = thermal generation current 

V = applied bias voltage 

q .. electronic charge 

n = quality factor 

T = temperature 

Rse = series resistance 

Rsh = shunt resistance 

a,m = constants 

k = Boltzmann constant 

The parameters in the equation may be related to physical mechanisms which are responsible for carrier 

transport in a solar cell. Each mechanism requires one or two fitting parameters. The four mechanisms used 

in the forward-bias region of the dark I-V curve-fitting investigations are: 

1. Injection current represented by the first term in the equation. The first term is the result of the simple diode 

equation. The model does not differentiate between injection and recombination current. It may be 

inappropriate to apply this expression to a p-i-n device, but it is useful for characterizing irradiation effects. 

The injection current parameters are 10 and n. 

2. Shunt current represented by the parameter Rsh in the second term . 

189 
3. Electric field and depletion effects in the intrinsic layer represented by the parameters a and m in the last 

term. 

4. Series resistance represented by Rse in all three terms. 

The results of curve fitting 

the dark I-V characteristic of a 

single junction cell in the 

forward-bias region is shown 

in Figure 4; the cell has a 500 

nm thick intrinsic layer. The 

dark I-V characteristic was 

measured follOwing the first 

two-hour anneal at 200°C 

following receipt of the cell 

from the fabricator. Curve 

fitting was carried out using 

Matlab software which employs a Nelder-Meade simplex 

search subroutine. The measured I-V values and calculated results are represented 

by filled squares and open 

triangles, respectively. The 

symbols for the calculated 

values are plotted over the 

measured values, and because the fit is good, the filled 

squares are barely discernParametric Fit to a Single-Junction Cell I-V Characteristic 

1E-2 

1E-4 

<" -

C 1E-6 

~ ~ 

o 1E-8 

1E-10 

10 = 1.5E-11 A 

n = 1.74 

Rsh = 800 KO 

0.10 

Rs = 270 

a. = 1.5 E-5 

m = 2.5 

Voltage (V) 

Measured 

Calculated 

1.00 

Figure 4: Typical curve fitting results for a single-junction cell showing the 

contributions of each of the four current mechanisms in a dark I-V 

characteristic. 

able. The contributions of each of the four current mechanisms are plotted and identified with labels in Figure 

4. The voltage ranges where the four mechanisms dominate in a major way are shown in Table I. 

Table I 

I Voltage Range Dominant Mecha- Corresponding Para meM nism ters 

0-0.2 shunt Rsh 

0.20-0.60 depletion a,m 

0.60-1 .00 injection 10 n 

near 1.00 series R~ 

A parametric study was carried out using seven single-junction cells with 500 nm i-Iayers. 

The cells were annealed for two hours at 200 "c following delivery. 10 was in the 1.5E-11 to 2.0 E-11 A range. 

The quality factor, n, ranged from 1.73 to 1-?! :2 rhe series resistance, Rse varied from 25 to 37 O. The 

parameter, a, ranged from 1.0 to 1.4 E-5 A/'V' m remained constant at 2.5. Agreement between the 

parametric model and measured I-V characteristic for the cells was <4% for the shunt and injection regions, 

and <8% in the depletion region. 

The dark I-V characteristics of a-Si:H single-junction cells have been investigated to determine the effects 

of proton irradiation on cell behavior; the cells had 500 nm thick intrinsic layers. The parametric model was 

used to quantify changes in dark I-V characteristics resulting from 1.00 MeV proton irradia_~on . The cells 

studied were irradiated with 1.00 MeV proton fluences between 1.46E 12 and 1 .46E 15 cm There were 

190 

1- _____ _ 
I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

I 

- -- 

twenty-one cells in the group, 

with three cells irradiated at 

each of the f1uences. One cell 

was chosen for the parametric 

study from each of the 

f1uences studied; the cell with 

lowest shunt current in the 

group of three cells was selected in order to determine 

fitting parameters with the best 

accuracy. Figure S shows the 

effect of 1.00 MeV proton f1uences on the measured dark IV characteristics for four of the 

seven f1uences investigated. 

The figure shows increasing 

proton f1uences result in lower 

currents in the 0.60 to 1.00 V 

range, the range associated 

Effect of 1.0 MeV Protons on Dark I-V Characteristics 

<" 1E-4 

-

-

c: 

~ ::J 

o 1E-8 

• 

• 

• 

1.46 E13 cm,2 

5.10 EU cm-l 

1.46 E14 cm,2 

1.46 E15 cm-l 

0.10 

Voltage (V) 

1.00 

with the injection current; it also Figure 5: Plot of measured dark I-V characteristics for single-junction 

shows the current in the 0.050 cells irradiated at various proton f1uences. 

to 0.20 V range is lower for 

higher f1uences; the current in 

this range is due primarily to the shunt current mechanism. 

Significant changes occurred in the thermal generation current density, Jo with irradiation and annealing. 

Figure 6 shows Jo following anneal 1, the first anneal after receiving the cells and pre-irradiation. Jo following 

anneal 1 is shown by open 

circles positioned on the graph 

at f1uences the cells were to 

be irradiated. Following irradiation Jo increased by a factor 

ot ~ tor a f1uence of 1.SE12 

cm and more than tour orders of magnitude !9r a f1uence ot 1.5E15 cm After 

annealing for two hours at 200 

OC, labeled anneal 2 on Figure 

6, Jo was restored to near pre1 E-3 

N - 1E-4 E 

(J 

~ 1E-5 

-

Effect of Proton Irradiation on J o 

o 

• 

c 

After anneal 1 

After Irradiation 

After anneal 2 

After anneal 3 

1.022 MeV protons 

irradiated values tor f1uences .,01 E-6 

less than 1.5E14 cm·2 • A third 

anneal further restored Jo for 

the two highly irradiated cells. 

Jo is much more sensitive to 

irradiation than power density. 

Figure 2 shows only a few 

percent change in power den-

~ at a fluence of 1.SE12 cmwhile Figure 6 shows a factor of two change in J o 

Changes in Quality factor, 

n, were also observed. Figure 

1 E-7 

1E-6 ~~~~~--~~~~~~~~--~~~~ 

1E12 1E13 1E14 1E15 1E16 

Proton Fluence (cm-2) 

Figure 6: Plot showing the effect of proton irradiation at various f1uences 

and subsequent annealing on reverse saturation current density, Jo 

obtained from parametric fitting. 

7 is a graph of fitted n values for pre-irradiation, post- irradiation and post-annealing conditions. For the same 

fluence range, n increased from 1.84 to 23.9 with irradiation. Subsequent anneals restored n to near preirradiated values as shown by the overlaid plots for these data; the open symbols essentially coincide for 

191 
anneals one, two and three. 

These results are consistent 

with the literature which shows 

Jo and n increase as the Quality of device material degrades; degradation in the material is believed to be due to 

an increase in defect density. 

Rse does not appear to be 

significantly influenced by irradiation. One exception was 

noted. Rse for one cell increased from 28 to 53 0 following irradiation; it was restored to 32 0 with annealing. 

The reasons for the large 

change in Rse for this cell are 

not understood. It is important 

to note that as the injection 

cu rrent decreases with increasing fluence, Rse becomes less of a factor in the IV characteristic and parametric fitting of Rse becomes more 

difficult. 

Effect of Proton Irradiation on Quality Factor 

25.00 r--"""--r-r-1"'TTT1,.,....--,--T'"T"T'T"nm----.-""T""T"""""''TTT''"---r-...-T'''T'TTTT1 

c: 20.00 

..: 

o 

1:) 15.00 

III 

&L. 

~ 10.00 

ii 

:::J 

a 5.00 

1.022 MeV protons 

o After .nn •• 11 

• 

(] After .nneal 2 

After .nneal 3 

0.00 '---'-..................... <.u.L_ ........... ......L..I .......... .u......--Io...--'-..................... '----'--"-'-............... 

1E12 1E13 1E14 1E15 1E16 

Proton Fluence (cm-2 ) 

Figure 7: Plot showing the effect of proton irradiation at various f1uences 

and subsequent annealing on quality factor, n, obtained from parametric 

fitting. 

The depletion current fitting parameters, a & m, behave differently with fluence. a is influenced by 

irradiation, while m does not a;wear to change wit~ ~radiation. a decreases from its annealed pre-irradiated 

val~5 approximately 1 E-5 Afr'" to about 7E-7 W ' with the largest fluence. It recovers to the 3E-6 to 5E-6 

A./'tr. range with annealing. We think it is significant that a does not recover to the annealed pre-irradiated 

value. m appears to remain constant at 2.5 with irradiation and annealing. 

Rst, of the cells appears to increase with irradiation and decrease with annealing. Pre-irradiation Rsh values 

range between about 300 and 800 kO, and increase to about 2000 kO with irradiation. Rsh decreases with 

annealing to values which range between 300 to 900 kO. Rsh exhibits a Switching behavior. Rsh of the twentyone irradiated cells was determined under three conditions: pre-irradiation , post-irradiation, and post-irradiation 

annealing. The behavior of Rst, from the analysis of the cells was similar to above results. Rsh was calculated 

using the current at a forward bias of 0.050 V; the procedure assumes Rsh dominates the current. The average 

Rst, prior to irradiation was 417 kn. Of the twenty-one cells, sixteen exhibited an increase in Rsh after 1.00 MeV 

proton irradiation. Cells irradiated with the same f1uences did not exhibit the same changes in Rsh and the 

reasons for this are not understood. The average Rsh increased to approximately 590 kO following irradiation. 

After annealing, seventeen of the twenty-one cells exhibited a decrease in Rsh The average Rsh following 

post-irradiation annealing was about 300 kn. The analysis of twenty-one irradiated cells confirms the trend that 

Rsh is increased by irradiation and decreases with post-irradiation annealing. 

We plan to pursue device simulation stUdies in an effort to explain the role of irradiation, in terms of 

fundamental material properties, on the parameters resulting from the fitting model. 

INSTABILITY STUDIES 

Triple-tandem cells cut from a module had fill-factors ranging from 0.37 to 0.68 under as-received or virgin 

conditions. Figure 8 shows light I-V characteristics for a cell under two conditions; the filled circles are for virgin 

conditions and result in a fill-factor of 0.37. The virgin cell is characteristic of one of the poorer as-received 

cells. Note the current has an unexpected structure between 1.50 and 2.00 V; the structure is concave down 

instead of concave up as would be expected. Following a two hour anneal at 200 °C, a light I-V measurement 

produced the open squares in Figure 8; the corresponding fill-factor is 0.66. After annealing the cell, the light 

I-V characteristic was similar to one of the better as-received cells. The fabricator of the triple tandem cells 

indicated the module was annealed for one hour in air at 100 °C following fabrication; the annealing procedure 

192 

------
~ ~ 

is employed by the fabricator 

as one of the steps in the fabrication of modules. 

It is difficult to investigate 

current mechanisms in triplejunction cells because of their 

complex structure. For this 

reason, single-junction cells 

were investigated in an effort to 

determine if the behavior observed in Figure 8 is also characteristic of single-junction 

cells. Figure 9 shows light I-V 

measurements for a singlejunction cell under light soaked 

and annealed conditions. The 

filled squares correspond to 

light I-V measurements for the 

cell annealed at 200 °C for two 

hours. I-V measurements following 32 hours of room temperature light soaking are rerr 

resented by open circles. Annealing the cell follOwing light 

soaking resulted in I-V measurements which are the 

same as before annealing. A 

comparison of Figures 8 and 9 

shows there is similar structure in the light I-V characteristics, suggesting the current 

mechanisms resulting in the 

initially poor performance of 

this partic~ar triple-junction 

cell may be elucidated with 

investigations of single-junction cells. 

The effect of light soaking 

on the forward-bias dark I-V 

characteristic for a single-junction cell is shown in Figure 10 

by open circles. The cell exhibits a switching effect in the 

region of the I-V characteristic 

where shunt current is the 

dominant current mechanism. 

Light I-V Characteristics for a Triple-Junction Cell 

4.0 .....---.---r---.--,---r----r-...,..--r-----r-----, 

3.0 • Virgin 

c 

- 2.0 

~ - 1.0 

Annealed two hours at 200 °c 

... 

c 

~ 0.0 

:J 

0 -1.0 

-2.0 

-3.0 

0.000 0.500 1.000 1.500 2.000 2.500 

Voltage (V) 

Figure 8: Light I-V characteristic for a triple-junction cell measured under 

virgin and annealed conditions. 

Light I-V Characteristics of a Single-Junction Cell 

30. 0 r--r---r-T"""-....--.....---r--.---r-~--.---r---.---.---r---.__, 

25.0 0 Light soaked 32 hours 

_ 20.0 • Annealed for 2 hours at 200 °C 

~ 15.0 

=10.0 

c 

~ 5.0 

:J 0.0 

o -5.0 

-10.0 

-1 5.0 1--L.-....'---"'---''---''---'---,----'----'----'---1---L---L---L---L--J 

-0.500 0.000 0.500 1.000 

Voltage (V) 

Figure 9: Light I-V characteristics of a single-junction cell under light 

soaked and annealed conditions. 

The switching effect has been observed in several cells; it occurs under both forward and reverse-bias 

conditions. An analysis of several cells shows the switching effect increases in frequency with light and forwardbias voltage soaking. It decreases with annealing and reverse-bias voltage soaking. The filled squares in 

Figure 10 represent a dark I-V characteristic measured following the application of a -2.00 V reverse-bias 

voltage for five minutes; this characteristic also corresponds to annealing at 200°C for two hours. The figure 

clearly demonstrates the role the history of the cell plays in I-V characteristics. 

Close inspection of dark I-V characteristics suggested the characteristics exhibited a time dependence. 

The role of time was investigated by including a delay time in the computer program. Measurements were 

taken with one and ten minute delays between the source-bias voltage steps. Time-dependent switching was 

observed in the reverse-bias voltage region of the dark I-V characteristics where the current switched from 

193 I 

I 

I 

J 
shunt dominated to a current 

mechanism with a slope between two and two and onehalf. Switching occurred at 

lower reverse-bias voltages as 

the delay time was increased, 

and the reverse-bias current 

was higher; the current mechanism is not understood. I-V 

characteristics exhibited another time dependence; as the 

voltage was scanned from the 

reverse-bias to the forwardbias region, the current decreased with increasing delay 

times. The observation suggests charge stored in a cell 

during reverse bias contributes 

to the current in the forwardbias region; however, the 

mechanism has not yet been 

fully characterized. 

Dark I-V Characteristics of a Single-Junction Cell 

10.1 

10.2 0 Light soaked 32 hours 

10-3 • Reverse biased at -2.00 V 

-

< 10-4 -

-

c 10'" 

~ ::l 10'-

0 10.7 

10'" 

10·' 

0.010 0.100 1.000 

Voltage (V) 

Figure 10: Dark I-V characteristics of a single-junction cell following light 

soaking and reverse-bias voltage soaking. 

DEVICE SIMULATION 

We have carried out device simulations with EPRI AMPS (ref. 8) in order to produce I-V characteristics. 

The simulated I-V characteristics may be compared with measured characteristics in an effort to determine 

fundamental material parameters. Changes were made in the material parameters for a simulation in order 

to produce an I-V characteristic similar to one obtained by measurement. Simulations were carried out and 

compared to I-V characteristics measured following irradiation of cells. The comparison provides information 

on the role of irradiation on fundamental material parameters. An understanding of the role of irradiation on 

the fundamental material parameters is necessary in order to develop a model for predicting EOl cell performance in a given space radiation environment. The work reported in this section represents our first efforts 

to use EPRI AMPS simulations to understand the effects of 1.0 MeV proton irradiation on hydrogenated 

amorphous silicon alloy based solar cells. 

EPRI AMPS was used to simulate a PIN device structure using various i-layer thicknesses and subband-gap density of state, DOS, functions. Devices were simulated with i-layer thicknesses of 200,500, and 

800 nm; 20 nm thick n+-Iayers and p+-Iayers were used. Gaussian and U-shaped DOS functions were used 

in the simulations. Characteristics of devices were simulated with midgap DOS values of 5E15, 5E16, 5E17, 

and 5E18 cm-3 eV' . The effective DOS at the conduction and valence band edges was 1 E19 cm3

• Devices 

were simulated under thermodynamic equilibrium conditions as well as light and voltage bias conditions. light 

and dark I-V's and electric field distributions have been generated in these preliminary simulations. 

The electric field is shown in Figure 11 for thermodynamic equilibrium conditions . The figure shows 

the spatial distribution of the electric field in a p+-i-n+ cell for a U-shaped DOS function with minimum DOS 

values of 5E15, 5E17, and 5E18 cm·

3 eV' . For reference purposes, the cell layers are shown at the bottom 

of Figure 11 . The n+-Iayer extends from 0 to 20 nm, the i-layer from 20 to 520 nm, and the p+ -layer from 520 

to 540 nm. The electric field decreases in the middle of the i-layer as the DOS is increased. The results 

suggest the reason for the decrease in cell power density with increasing 1.0 MeV proton f1uence is the 

generation of defects in the ~Iayer. Defects in the EPRI AMPS model are represented by electron energy states 

in the sub-band-gap region. The effect of the defects is to trap charge carriers which increases the spacecharge density. The increased space charge density in the i-layer reduces the electric field distribution which 

in tum reduces the collection of carriers. We plan to pursue the EPRI AMPS simulation in order to obtain the 

fundamental material parameters for the development of a predictive model. 

194 
I 

I 

I 

I 

I 

I 

[ 

I 

[ 

r 

I 

I 

I 

\ 

CONCLUSIONS 

The effect of 1.0 MeV 

proton irradiation is to degrade 

the normalized power density 

of cells with six different structures by less than a few percent for ~ences less than 

1 E 12 em _~or fluences 

above 1 E15 em cell power 

degrades to less than 10% of 

the initial power. Reductions in 

the normalized power den~ 

in the 1 E12 through 1 E14 cm 

proton fluence range depends 

on cell structure. Cells with 

triple junctions have the best 

radiation resistance, dual-junction cells are next, and singlejunction cells have the lowest 

radiation resistance. Singlejunction cells with a-Si,Ge:H ilayers have better radiation 

resistance than cells with aAMPS Simulated Electric Field for a Single-Junction Cell 

iE6 

";"E 1E3 u 

~ 

'0 i 1 

u.. 

p. -fayer 

~-----------I-fayer------------~ 

1E~~~~~~~~~~~~~~~~~~ 

o 100 200 300 400 500 

Distance (nm) 

Figure 11: AMPS simulated thermodynamic equilibrium electric field for a 

SOO nm thick I-layer PIN device with SE1S, SE17, SE18 cm·3 eV1 midgap 

DOS. 

Si:H i-layers. A parametric fitting model was employed to determine the effect on 1.0 MeV proto_~ irradiation 

on cell parameters. Following irradiation Jo increased by a f~por of two for a fluence of 1.SE12 cm and more 

than four orders of magnitude for a fluence of 1.SE1S cm After anneal~~ for two hours at 200°C, J o was 

restored to near pre-irradiated values for fluences greater than 1.SE14 cm A third anneal further restored 

J o for the two highly irradiated cells. Jo was found to be more sensitive to irradiation than power density. 

Changes in the quality factor, n, were also observed with irradiation. The quality factor increased from 1.84 to 

23.9 with irradiation; annealing at 200°C restored n to near pre-irradiated values. The investigations show 

there is similar structure in the light I-V characteristics of triple and single-junction cells suggesting the current 

mechanisms resulting in the initially poor performance of a triple-junction cell may be elucidated with 

investigations of single-junction cells. Investigations of dark I-V characteristics show there is a switching effect 

in the region of the I-V characteristic where shunt current is the dominant current mechanism. The switching 

effect has been observed in several cells; it occurs under both forward and reverse-bias conditions. An analysis 

of several cells shows the switching effect increases in frequency with light and forward-bias voltage soaking; 

it was observed to decrease with annealing and reverse-bias voltage soaking. Time-dependent switching was 

observed in the reverse-bias voltage region of the dark I-V characteristics where the current switched from 

shunt dominated to a current mechanism with a slope between two and two and one-half. Switching occurred 

at lower reverse-bias voltages as the delay time was increased, and the reverse-bias current was higher; the 

current mechanism is not understood. Device simulation studies were carried to determine the role of the subband-gap density of electron states on the electric field distribution in the p+-i-n+ ;Iayers. The electric field 

decreased in the middle of the i-Iayer as the density of states was increased. The results suggest the reason 

for the decrease in cell power density with increasing 1.0 MeV proton fluence is the generation of defects in the 

i-layer. The effect of the defects is to trap charge carriers which increases the space-charge density. The 

increased space charge density in the i-layer reduces the electric field distribution which in turn reduces the 

collection of carriers. 

195 
REFERENCES 

1. Geoffrey A. Landis and James R. Woodyard, Radiation Damage in Thin Film Solar Cells, Solar Cells, 

1991, Volume 31, page 297. 

2. James R. Woodyard, Dependence of 1.0 MeV Proton Radiation Resistance of a-Si:H Alloy Solar Cells on 

Cell Thickness, Amorphous Silicon Technologv-1992, Materials Research Society Symposium Proceedings, Edited by Malcolm J. Thompson, Yosihhiro Hamakawa Peter G. LeComber Arun Madan and Eric 

Schiff, Volume 258, 1992, page 1151 . 

3. Kenneth R. Lord II, Michael R. Walters and James R. Woodyard, Investigation of the Radiation Resistance 

ofTnple-Junction A-Si:H Alloy Solar Cells Irradiated with 1.00 MeV Protons, Space Photovoltaic Research 

and Technology 1992, NASA Conference Publication 3210, page 98. 

4. Kenneth R. Lord II, Michael R. Walters and James R. Woodyard, Investigation of Light and Dark 

Characteristics of a-Si:H Alloy Cells Irradiated with 1.0 MeV Protons, Twenty Third IEEE Photovoltaic 

Specialists Conference-1993, page 1448. 

5. Kenneth R. Lord II, Michael R. Walters and James R. Woodyard, Investigation of Shunt Resistances in 

Single-Junction a-Si:H Alloy Solar Cells, Amorphous Silicon Technology-1994, Materials Research Society 

Symposium Proceedings, Edited by Malcolm J. Thompson, Yosihhiro Hamakawa Peter G. LeComber Arun 

Madan and Eric Schiff, In Press. 

6. S. Abdulaziz and J. R. Woodyard, Annealing Characteristics of Amorphous Silicon Solar Cells Irradiated 

with 1.00 MeV Protons, Eleventh Space Photovoltaic Research and Technology Conference-1991, NASA 

Conference Publication 3121, 1991, page 43-1 . 

7. M. Bennettt and R. Podlesny, Two-Source Simulator for Improved Solar Simulation, Twenty-First IEEE 

Photovoltaic Specialists Conference-1990, page 1438. 

8. J. K. Arch, P. J. McElheny, S. J. Fonash, and W. H. Howland. AMPSI1D: A One-Dimensional Device 

Simulation Program for the Analysis of MicroElectronic and Photonic Structures. Penn State University. 

November 2,1990. 

196 

~--·1 

I 

i 

I 

I 

I 

I 

I 
----"~ ~-

39 ??fZ 

N95- 20523 

THE USE OF DISPLACEMENT DAMAGE DOSE TO CORRELATE DEGRADATION 

IN SOLAR CELLS EXPOSED TO DIFFERENT RADIATIONS 

Geoffrey P. Summers 

U.S. Naval Research Laboratory 

Washington, DC 

Edward A. Burke, Philip Shapiro, Richard Statler, 

and Scott R. Messenger 

SFA, Inc. 

Landover, Maryland 

and 

Robert J. Walters 

U.S. Naval Research Laboratory 

Washington, DC 

SUMMARY 

It has been found useful in the past to use the concept of "equivalent fluence" to compare the 

radiation response of different solar cell technologies. Results are usually given in terms of an equivalent 

1 MeV electron or an equivalent 10 MeV proton fluence. To specify cell response in a complex space· 

radiation environment in terms of an equivalent fluence, it is necessary to measure damage coefficients 

for a number of representative electron and proton energies. However, at the last Photovoltaic Specialist 

Conference (ref.1) we showed that nonionizing energy loss (NIEL) could be used to correlate damage 

coefficients for protons, using measurements for GaAs as an example (ref.2). This correlation means that 

damage coefficients for all proton energies except near threshold can be predicted from a measurement 

made at one particular energy. NIEL is the exact equivalent for displacement damage of linear energy 

transfer (LET) for ionization energy loss. The use of NIEL in this way leads naturally to the concept of 10 

MeV equivalent proton fluence. The situation for electron damage is more complex, however. In this 

paper it is shown that the concept of displacement damage dose gives a more general way of unifying 

damage coefficients. It follows that 1 MeV electron equivalent fluence is a special case of a more general 

quantity for unifying electron damage coeffcients which we call the effective 1 Me V electron equivalent 

dose. 

INTRODUCTION 

The most common way of specifying radiation environments for solar cells is in terms of their 

response to a fluence of 1 MeV electrons. The Solar Cell Radiation Handbook (ref.3) for example is full of 

such tables and figures showing the degradation of key photovoltaic parameters in a variety of space 

orbits. Although the effect of an electron or a proton fluence is the way displacement damage is generally 

determined, the absorbed dose is the parameter used to describe ionization effects in biological and 

microelectronic systems. Absorbed dose, which measures the energy deposited per unit mass as a result 

of ionization, was found to be so useful in comparing the effect of different radiations that a special unit 

was introduced to measure it. The original unit was the rad, but this has been superseded by the Gray (1 

J/kg). Presumably the reason for determining displacement damage in terms of fluence originated in the 

way dOSimetry is performed at particle accelerators. Conversely x-ray and y-ray dOSimetry is performed 

using techniques such as thermoluminescent emission, the magnitude of which is determined by the 

absorbed energy or dose. Unlike absorbed dose, fluence cannot be used to correlate the effect of 

different radiations. However, the product of fluence and NIEl gives the displacement damage 

197 
equivalent of absorbed dose, which we will show gives a sound physical basis for correlating both electron 

and proton displacement damage coefficients . The concept of equivalent fluence then follows in a 

straightforward way under certain conditions. In other cases ~ is necessary to introduce the concept of an 

effective 1 MeV electron equivalent dose in order to correlate electron damage coefficients. 

DAMAGE CORRELATIONS USING ABSORBED DOSE 

1. PROTONS 

Figure 1 shows experimentally measured curves of the normalized power degradation for pIn 

GaAs solar cells produced by increasing fluences of incident monoenergetic protons with energies of 0.5, 

1.0, 3.0, and 9.5 MeV. The data points plotted in figure 1 are taken from the original line figure given by 

Anspaugh in reference 2 and the lines are shown only to guide the eye. The displacement damage dose 

for each data point was then calculated from the product of the NIEL in GaAs S(E) and the respective 

fluence $(E) for the particular proton energy E, using the NIEL values given in reference 4. The data were 

replotted as shown in figure 2, where the abscissa is now the absorbed displacement dose given in units 

of MeV/g. As can be seen the data for all proton energies when plotted in this way collapse on to a single, 

universal line. This line, which represents the complete response of GaAs cells to protons of all energies, 

can be produced using protons of any single energy. Conversely, if degradation data exist for anyone 

energy the experimental line for another energy such as 10 MeV protons could be readily 0t?tained using 

the equation: 

(1) 

Equation (1) leads naturally to the concept of 10 MeV proton equivalent fluence, which is widely 

used to simulate the effect of a complex proton environment given in terms of a differential proton 

spectrum d<l>p(E)/dE. The 10 MeV proton equivalent fluence is calculated from the integral of the proton 

NIEL over the proton spectrum, divided by the NIEL for 10 MeV protons, Le., 

<l>p(10) = [1/Sp(1 O)J.iSp(E).[d$p(E)/dE) dE. (2) 

It is usual to take damage correlation further than Eq.(2) by specifying radiation effects in terms of 

the effect of a 1 MeV equivalent electron fluence. This requires first reducing the 10 MeV equivalent 

proton fluence to 1 MeV equivalent electron fluence and then adding the result to the 1 MeV electron 

fluence equivalent to the total electron environment present. However, because of complex~y in the way 

some semiconductors respond to electrons, such calculations require using the concept of equivalent 

damage dose in a modified way as we now show. 

2. ELECTRONS 

It has been found that a linear dependence of photovoltaic parameter change on absorbed dose 

as shown in figure 2 is always found for relatively high NIEL particles such as protons and helium ions. A 

linear dependence is also found for low NIEL particles such as electrons incident on n-type Si, GaAs and 

possibly other semiconductors. In these cases, Eq.(1) can be used directly to convert a 10 MeV proton 

fluence to a 1 MeV electron fluence from a ratio of the respective NIELs. Similarly, a 1 MeV equivalent 

electron fluence can be defined for an electron environment in the same way as discussed above for the 

10 MeV proton equivalence, Eq.(2), by simply substituting the appropriate symbols. 

However, for devices with p-type active regions electron-induced changes are often found to vary 

in a way which depends on the square of the NIEL of the electrons . This finding is analogous to the 

different response found for some biological systems to high and low LET ionizing radiations, which leads 

to the concept of the "quality factor, an of the radiation. The quality factor expresses the relative effect of 

a given radiation to the effect of x-rays, for which a ,. 1. In the displacement damage case we will show 

198 
below that we can define a quantity called the "effective 1 MeV electron equivalent dose", which is 

obtained by multiplying the absorbed dose for electrons of energy Ej, i.e., S(Ej).cp(Ej), by the ratio 

S(Ej)/S(1 .0), where 1.0 refers to 1 MeV. This ratio is the displacement equivalent of the quality factor. 

The reason for the choice of normalizing energy is that the response to 1 MeV electrons is the traditional 

way of comparing the behavior of different kinds of solar cells. 

As an example we consider the data of Yamaguchi and Amano (ref.5) for changes in minority 

carrier diffusion length in p-type GaAs irradiated with O.B, 1.0, 2.0, and 4.0 MeV electrons and with CoS) 

gamma rays. The data were derived from in-depth profiles of short-circuit current changes measured in 

the emitter of a pin GaAs cell using the EBIC method. The difference between the reciprocal square of 

the post- and pre-irradiation diffusion lengths, 4 and 4>, respectively is given by 

1/42 - 1/4>2 = K(E) .cp (3) 

where K(E) is the diffusion length damage coefficient for electrons of energy E. The NIELs for electrons 

in GaAs with energies of 0.8, 1.0,2.0, and 4.0 MeV are 21.4, 26.5, 44.2, and 63.2 eV.cm2/g, respectively. 

The data points from figure 4 of reference 5 has been replotted in figure 3. The line for 4.0 MeV electrons 

is a least squares fit to the data. The other lines are calculated from this reference line using the ratio of the 

squares of the NIELs for the respective energies . C060 gamma rays produce a spectrum of mostly 

Compton electrons and the average NIEL assuming a linear dependence on electron energy is 9.40 

eV.cm2/g. Assuming a quadratic dependence gives 155.0 (eV.cm2/g)2. Details of these calculations, 

which are somewhat complicated, have been discussed briefly in reference 3. As an example of the 

magnitude of the difference associated with the effect of a linear or a quadratic dependence on NIEL 

consider the data for 4.0 MeV electrons and C060 gamma rays. With a linear dependence the 4.0 data 

would be calculated to shift to the right by a factor of 63.219.40 = 6.72, which clearly would not coinCide. 

with the experimental data for C060. A quadratic dependence gives a shift of 45.2 and the agreement with 

the data can then be seen in figure 3 to be excellent. 

The quadratic dependence of NIEL means that Eq.(1) must be modified for electrons on p-type 

GaAs to give. 

(4) 

which can be rearranged to give 

(5) 

where the subscripts 1 and 2 for different types of particles has now been dropped because the 

discussion applies to electrons only. Eq.(5) can be written 

(6) 

Eq.(6) shows how an effective 1 MeV electron equivalent dose can be defined, i.e., 

Dose(1.0) = Dose(E2).[S(E2l/S(1 .0)) (7) 

Figure 4 shows the data in figure 3 replotted using Eq.(7) to calculate the effective 1 Mev electron 

equivalent dose for each point. As can be seen in figure 4, when plotted in this way, all the data col/apses 

on to a single line. This line represents the general response of GaAs solar cells to electrons of all 

energies . 

199 
DISCUSSION 

The results presented here show that the concept of displacement damage dose gives a more 

fundamental way of comparing the radiation response of solar cells to irradiation than the more commonly 

used particle f1uence. A question that comes immediately to mind in comparing the radiation response of 

cells, however, is the cause of the general linear dependence of damage coefficients on NIEL found for 

protons in contrast to the quadratic dependence found for electrons on p-type Si, GaAs and possibly 

other semiconductors. This question is more complicated than is at first apparent because clearly there is 

a point at which a plot of the coefficients for p-type cells versus NIEL for protons and electrons, when 

extrapolated, must coincide. At this "critical" point a linear dependence would presumably be found, 

assuming there is a particle that actually has the corresponding value of NIEL. The answer to the question 

must lie in the nature of the stable defects caused by different particles. Higher LET particles such as 

protons produce defect cascades that have a tree-like structure with dense defect concentrations at the 

end of branches containing isolated defects (ref.6). Lower LET particles such as low energy «-50 MeV) 

electrons produce mostly isolated point defects. It is the details of the formation mechanism of the point 

defects affecting the electrical properties of the solar cell that determine the dependence on NIEL. The 

"critical" value of the NIEL appears to correspond to the value at which the tree-like cascade structure 

becomes dominant. 

REFERENCES 

1. Summers, G.P., Burke, E.A. , Walters, R.J., Shaw, G.J., Shapiro, P., Statler, R.L., and Messenger, 

S.R.: A General Method for Predicting Radiation Damage to Solar Cells in the Space Environment. 

IEEE 23rd Photovottaic Specialists Conference, 1993, pp. 1426-1431 . 

2. Anspaugh, B.: Proton and Electron Damage Coefficients for GaAslGe Solar Cells. IEEE 22nd 

Photovottaic Specialists Conference, 1993, pp. 1593-1598. 

3. Tada, H.Y., Carter, J.A., Anspaugh, B.E. , and Downing, A.G.:Solar Cell Radiation Handbook, 3rd 

edition. JPL Publication 82-69, 1982. 

4. Summers, G.P., Burke, EA, Shapiro, P., Messenger, S.R., and Walters, R.J.: Damage Correlations 

in Semiconductors Exposed to Gamma, Electron and Proton Radiations. IEEE Trans. in Nucl. Sci. , 

vol., 1993, pp. 1372-1379. 

5 Yamaguchi, M. and Amano, C.: CoSO Gamma Ray and Electron Irradiation Damage of GaAs Single 

Crystals and Solar Cells. J.Appl. Phys., vol. 54, 1983, pp. 5021-5029. 

6. Wood, S., Doyle, N.J., Spitznagel, J.A., Choyke, W.J., More, A.M., McGruer, J.N., and Erwin, R.B.: 

Simulation of Radiation of Damage in Solids. IEEE Trans. in Nucl. Sci., vo1.28, 1981, pp. 4107-4112. 

200 

- --- - - - - --- -- - -- - - - --- - - - - - - - - - -
f';:; 

-- OG t:=:::: "- ~ 

01 

"" ~ 

~ '" '\ " .... 

':l 0 .1 

... r 

0.5 .8.5 w.v 

*3.0 "'.V 0.5 . 1.0 W.V 

.0.5 "'V 

0.' 

1E . OQ 

'\ '\ 

1E+10 lE+l1 lE+12 lE+ll 

Proton Auence (pIcm 2, 

Figure 1-- Power loss of GaAslGe solar cells 

versus proton lIuence from reference 2. 

":i 

Co 

<l 

.~--------~----------------------, 10 ........ _ 

x. .. ...., __ ......... A ..... .., ........ 

10'Y~ 

, ..... . • 0 

10.,L, ~...L...J.LL---ML---~ .. ,--...L-~ .. ~---7. .. :------=-' .. .0 10 ~0F\lenco ~_.) .0 '0 

Figure 3-- Diffusion length damage coefficients 

versus electron fluence from reference 5. The 

line through the 4.0 MeV data is a least squares 

fit. The other lines are calculated from the 4.0 

MeV line assuming a quadratic dependence on 

NIEL. 

• ... 

o. g '. 0.1 

.. 0.5 ... 8.5 MeV 

• 3.0 u.v 0.5 • 1.0 M.V 

1 O.S u.V 

0.' 

t E+07 lE+04!I lE+09 l E +t O iE.,1 

Figure 2-- Power loss of GaAslGe solar cells 

versus displacement damage absorbed dose 

using data from figure 1 . 

lE.12 

• r----------r------------------------~ 10 ........ _ 

x 

........ ....... -

_ 

Ao ..... V .. __ 

.Y~ . . . . ....... ... .... .. . 

10 1----------' 

, .. .. .... ..... .. .. ... .. . 

10 

1 ~~ .. 1-----,-

0

.'-. ----,-=0.'-. ---'-'-0'~' ----,-0-= ..-----, 0-:'; .. -----' o~ .. ;-' 

, MeV E1ectToft EqutvaJent 00 •• ('V/gI 

Figure 4- Diffusion length damage coefficients 

versus 1 MeV electron equivalent dose using the 

data from figure 3. 

201 
-1 

I 

· .. I 

I 
397?f5 

N95-20524 

ANNEALING OF IRRADIATED n+p InP BURIED HOMOJUNCTIONS 

R.J. Walters and G.P. Summers 

U.S. Naval Research Laboratory 

Washington, DC 

and 

M.L. Timmons, R. Venkatasubramanian, J.A. Hancock, and J.S. Hills 

Research Triangle Institute 

Research Triangle Park, North Carolina 

INTRODUCTION 

At the last SPRAT conference, the Naval Research Laboratory (NRL) presented results from two 

experiments. One studied n+p diffused junction (OJ) InP solar cells (ref. 1), and the other studied n+p 

shallow homojunction (SHJ) InP mesa diodes grown by metal organic chemical vapor deposition 

(MOCVD) (ref. 2). The former work showed that a DJ solar cell in which the maximum power (P....J 

had been degraded by nearly 80% under irradiation recovered completely under short circuit illumination 

at 450K (fig. 1). The recovery was accompanied by the removal of all but one of the radiation-induced 

defect levels (fig. 2). The latter work, on the other hand, showed that the radiation-induced defects in 

the SRJ diodes did not anneal until the temperature reached 650K (fig 3). These results suggest that an 

irradiated DJ solar cell, under illumination, will anneal at a temperature 200K lower than an irradiated 

SRJ cell. This is an unexpected result considering the similarity of the devices. The goal of the present 

research is to explain this different behavior. 

This paper investigates two points which arose from the previous studies. The first point is that 

the DJ cells were annealed under illumination while the SRJ diodes were annealed without bias. Given 

the known sensitivity of radiation-induced defects in InP to minority carrier injection-annealing (ref. 3), 

it is possible that the illumination is the cause of the lower annealing temperature. To test this, irradiated 

SHJ InP solar cells of the same structure as the diodes of ref. 2 have been annealed under illumination 

at temperatures ranging from 300-500K. The results of these experiments are presented here. The 

conclusion is that the illumination had no effect on the annealing, and the annealing follows that which 

was expected from the diode study of ref. 2 - i.e. the recovery is limited by a lack of defect annealing 

in this temperature range. 

The second point investigated here is that the emitters of the DJ and SRJ devices were 

significantly different. The emitter of the DJ cells of ref. 1 were thick (O.3J.Lm) and grown by diffusion 

which forms a graded carrier concentration profile. However, the emitter of the SHJ diodes of ref. 2 and 

cells studied here were thin (0.03J.Lm) with a uniform carrier concentration profile. The thicker emitter 

means the DJ emitter current is much larger than that of the SRJ cells. Also, the different concentration 

profiles results in different spacial distributions of the junction electric field. These differences may 

impact the solar cell annealing characteristics. To investigate this, the Research Triangle Institute (RTI), 

under contract to NRL, used MOCVD to grow epitaxial cells with the structure of the DJ cells. The cells 

have a O.25J.Lm thick emitter with a graded dopant profile (fig. 4). The cells are referred to as deep 

homojunctions (DHJ). The growth of these cells is the subject of a companion paper given by RTI at 

this conference. The present paper reports the annealing characteristics of the DRJ cells following 

irradiation. Results of annealing both in the dark and under illumination in the temperature range of 

175K-450K are presented. In general, the annealing characteristics of the DRJ cells was similar to that 

of the SRJ cells especially in terms of the defect spectrum. However, significantly more recovery of the 

short circuit current (lac) was seen in the DRJ cells than in the SRJ cells. Therefore, the thicker, graded 

203 

PR~ ,- Ot ' PAGE LANK NOT FILMED 

- - - _. - --~--~-- -~- - -------- --~- -~-~----~~-------~-~ 
emitter seems to enhance the recovery of I.e, but the overall cell recovery is incomplete due to a lack of 

defect annealing just as in the case of the SHJ cells. 

10 c=~~----~--~--------~------, 

9 t:::==:::::=:""----_=__ 

8 

'< E 6 

'0.../ 

C 5 ., 

~ 4 u 3 

2 

300K. 45min------'~ 

1 MeV electron irrodiated DJ InP solor cell 

~.O 0.2 0.4 

1 sun , AMO 

86 K 

OJ InP solar cell 

1 5 1 MeV elec ton irrod. 

o 5 

c 

. S2' 0 j....-..-.----- (f) 

(f) 

f-- 5 

--.l 

Ql0 

-"-"'- ---~''''-'''' ''' ... "", ... -.... -,,-

....... ........ .... 

EC 

ED 

H5 

After 

irradiation 

Illuminated anneals 

.............. ·450K. 1 hr 

·--·-"-"·--500K. 1 hr 

100 1 40 180 220 260 300 

Temperature (K) 

Figure 1: Photo-injection annealing of an irradiated Figure 2: The defect annealing measured in the 

DJ solar cell. Illumination at 275K causes DJ cell of fig. 1. The illumination at 450K 

substantial recovery, and full recovery is seen at removed all but the ED defect. 

450K. 

EXPERIMENT AL DETAILS 

1.0 

n+p SHJ InP meso diode 

1 MeV elec tron irrad. HS 

Data measured on two n+p SHJ solar cells 

are presented. One cell was annealed in the dark, L2 0.5 

N =2xl0'· cm-J ) 

T':'=50s-' H 3 "-

and the other was annealed under illumination. The ~ 

data for the cell annealed in the dark has already ~ 0 .0 

been published in ref. 4, but the data is reproduced ~-0.5 . 

here to allow for direct comparisons. This cell was is 

grown using MOCVD by Spire Corporation under -1 .0 E C 

contract to NRL and has the structure of fig. 5. The -1 .5 

--~- ~ 

' - ' , 

Thermal Annealing 

500K 

586K 

- ---.-.. --.-.. - --650K 

cell was part of run number 5414-7-1. Si was the 

emitter dopant, and Zn was the base dopant. The 

total cell area was O.25cm2 and BOL PV parameters 

were: 1..,=8.78mA, Voc=O.883V, P max=6.48mW, 

FF=O.836, Eff= 18.95% (total area). The cell was 

-2.0 L.--'----'_~_'_ __ "'-_'__'--_'_---'_~__l 

BO 100 120 140 160 1 BO 200 220 240 260 2BO 300 

Temperature (K) 

10 MeV proton irradiated up to a fluence of 1012 cm' 

2, Dosimetr~ was achie~ed ~rough a faraday cu.p Figure 3: Annealing of an irradiated n+p SHJ 

and current. I~tegrator CirCUit, an~ the fluence IS diode. Only the H3 and H4 defects anneal below 

known to wlthm about 15%: CapaCItance vs voltage 500K. Not until 650K do the remaining defects 

(CV) measurements on a dIOde grown on the same anneal which is 200K higher than in the DJ cells 

wafer as the solar cell and irradiated simultaneously (ref. 2). 

204 

L ___________________________________________ _ _ 

-1 
I~ 

T9 Metal (Ag/Pd) ~d 

~ns.N~F.' ~Rcoahog ~ 

- >- - 0l~m n + emi Her - Se doped ~- ----------------------------- 3~m p base - Zn doped 

D· buf! er aver 

SUBSTRATE. p-type. Zn-doped 

t Back metal (Ag/Zn) contact 

MOCVD 

layers 

Figure 4: A schematic drawing of the DHJ cells 

grown by RTI. These are epitaxial cells grown to 

simulate the DJ cell structure, so the emitter is thick 

with a linearly graded dopant profile. 

Top Metal Grid 

311m thick p-type base 

InP SUBSTRATE 

Back metal contact 

Figure 5: A schematic diagram of the n +p SHJ 

cells. The emitter of these cells is thin with a 

uniform carrier concentration profile. This is the 

same structure as the mesa diodes of ref. 2. 

with the cell showed the base dopant concentration to be a uniform 1016 cm3 before and after irradiation. 

The second n+p SHJ solar cell studied here (the cell which was annealed under illumination) was 

also grown using MOCVD by Spire Corporation. This cell came from run number 2353-1-l. The cell 

structure was also that of fig. 5. Se was the emitter dopant, and Zn was the base dopant. The total cell 

area was 0.25cm2 and BOL PV parameters were: I.c=7.32mA, Voc=0.850V, P max=4.927mW, 

FF=0.792, Eff= 14.4% (total area). The cell was irradiated with alpha (a) particles from an Am-241 

source. A special acknowledgement goes to Dr. Pascale Gouker of Spire for performing the irradiations 

and supplying NRL with this cell. The cell was irradiated up to a fluence of 8 x 1011 a-particles/cm2 

which is equivalent to 5 x 1016 1 MeV electrons/cm2 (the calculation of the 1 MeV electron equivalent 

fluence was done by E.A. Burke formerly of Spire). No pre-irradiation CV measurements of this cell 

were possible, but the target base carrier concentration was a uniform 5 X 1016 cm-3 • CV measurements 

were performed on the cell after irradiation. The irradiation caused carrier removal near the junction 

(fig. 6). 

Two of the RTI DRJ cells were studied. These cells came from RTI MOCVD run numbers 

1955-a1 and 1955-a4. The total cell area was 0.25cm2. Due to metal contact shadowing on top of the 

cells, the active area was O.16cm2. BOL PV parameters were: cell 1955-al- I.c=4.82mA, Voc=0.833V, 

Pmax=3.32mW, FF=0.828, Eff=15.2% (active area) and cell 1955-a4 - I",=4.85mA, Voc=0.841V, 

P max=3.40mW, FF=0.834, Eff= 15.6% (active area). The cells were irradiated with 3 MeV protons 

at the Naval Surface Warfare Center (NSWC) in White Oak, MD using a Pelletron Accelerator. The 

sample.." were mounted on 1I grol1nciE'.d plMe) :mcl the flu~nce was determined by measuring the target plate 

current. Unfortunately, due to incomplete charge collection and an obscuring aperture in the beam line 

which was not accounted for, the dosimetry was poor. The best guess at the fluence is 6.12 X 1011 cm-2 , 

but this could be off by as much as 40%. The cells were mounted and irradiated separately. The beam 

205 
L_ 

1017 r---~----~----~--~----~----' 

Am-241 (X - irrad n+p SHJ cell 

8 9" = 8x l 0 " c m - 2 

7 

c 6 o 

:.;:; 5 

E' C 4 OJ 

u 5 J U 

'- 2.5 Q) 

.c:: 

'- 2 o 

U 

1.5 

.. 8 __ After irradiation 

·······0 ··· .. After annealing 

6 10 ~--~----~----7---~~--~--~ 

o 2 3 4 5 

Width of Depletion Region (J.Lm) 

Figure 6: The carrier concentration profile of Spire 

cell 2353-1-1 after irradiation and after annealing. 

Carrier removal occurred near the junction, and the 

annealing had no effect. 

1019 

~ 

I' 

E1018 ~ 

j:;-

' iii 

c 

OJ 

0 

:u 

'E1017 

0 

U 

3 

10 16

0 

Cell 1955-a4 

2 3 4 5 6 7 8 9 10 

Voltage (V) 

Figure 7: Carrier concentration profiles of the 

two DRJ cells studied here. The irradiation had 

no affect on the concentration. The inset scale 

indicates the depth relative to the junction. 

current was ;:5; 22 nA/cm2 , and the irradiations were completed in a matter of seconds. The irradiations 

were done in vacuum, at open circuit, and in the dark. CV measurements on diodes grown adjacent to 

the cells before irradiation showed the base dopant concentration of cell 1955-a1 to be about 3 x 1018 cm3 

and of cell 1955-a4 to be about 1017 cm-\ but both profiles were somewhat graded (fig. 7). These diodes 

were irradiated simultaneously with the cells with no effect on the carrier concentration. 

The solar cells were mounted in a continuous flow liquid nitrogen cryostat which is equipped with 

a sapphire window to allow for illumination. The temperature range of the cryostat is 86 - 500K. All 

of the experiments were performed using the cryostat including the annealing experiments. The chamber 

was under constant vacuum. An Oriel l000W Xe arc lamp solar simulator with AMO filtering was used. 

The lamp intensity was adjusted to one sun using an InP reference cell calibrated by Keith Emery at 

NREL. Illuminated IV curves were measured using two Keithly 617 electrometers and a Kepco 50-2M 

bipolar amplifier. The radiation-induced defect spectrum was characterized through deep level transient 

spectroscopy (DLTS). A Bio-Rad DL4600 spectrometer with a Boonton 72-B capacitance meter was 

used. The reverse bias for every DLTS measurement was -2V. A saturation fill pulse was consistently 

used to ensure complete trap filling. The same capacitance meter was used for the CV measurements. 

EXPERIMENT AL RESULTS 

Thermal Annealing of a SRJ cell in the Dark 

One of the goals of this report is to compare the annealing characteristics of SRJ InP solar cells 

which have been annealed in the dark with those which have been annealed under illumination. NRL and 

Spire have published thermal annealing in the dark data in ref. 4, and that data is now reproduced for 

206 
I ----- - - ----

comparison purposes. It should be noted that a similar experiment has been performed at NRL on a 1 

MeV electron irradiated Spire MOCVD grown n+p SHJ loP solar cell which showed similar results. 

Following irradiation with 10 MeV protons up to a fluence of 1012 cm-2

, cell 5414-7-1 was 

thermally annealed. at 415 and then 500K. The cryostat was completely dark and no connections were 

made to the cell during the annealing. At 415K, the cell was annealed in incremental time steps 

beginning with 15 minutes and ending after 3 hours. After 3 hours at 415K, no further changes were 

observed, so the temperature was raised to 500K. The cell was again annealed in incremental time steps 

beginning with 15 minutes and ending after 1 hour and 15 minutes. After each time increment, the 

illuminated IV curve was measured at 298K (fig. 8). Also, both the majority and minority carrier trap 

DLTS spectra were measured after each time step (fig. 9). With the present DLTS system, a positive 

DLTS signal indicates majority carrier capture while a negative signal indicates minority carrier capture. 

Considering the annealing at 415K, the only recovery of the PV parameters occurred after the 

first 15 minutes. The recovery consisted mainly of an increase in I.e (about a 3.5% increase). The 

recovery of Voc and the FF was 1 % or less. The first 15 minutes at 415K also caused a small decrease 

in the concentration of all of the hole traps (fig. 9). The ED peak just become visible due to the decrease 

in the obscuring H3 peak. The continued annealing at 415K caused the complete removal of the H3 and 

H4 defects. Increasing the temperature to 500K caused another increase in I.e, a very small one in V"',, 

but none in FF. No changes in the DLTS spectra were seen. In total, the recovery was minimal- Pmax, 

which was degraded by 25%, only recovered by 9.4%, and only the H3 and H4 defects showed 

annealing. 

Thermal Annealing of a SHJ cell Under Illumination 

One of the missing pieces in the current discussion is a study of the annealing of an irradiated 

SHJ cell under illumination. This data is now presented. Following the a-particle irradiation, Spire n+p 

SHJ loP solar cell number 2353-1-1 was illuminated short circuit at 350, then 400, and then 500K. As 

in the previous experiment, the annealing was carried out for progressively longer times at each 

temperature until no more changes were apparent. The temperature was then increased. and the annealing 

repeated. After each time step, an illuminated IV curve was measured at 298K along with the full DLTS 

spectrum. The recovery of the IV curves is shown in fig. 10. The DLTS spectra are shown in fig. 11. 

The illuminations at 350K had no effect on the IV curve or the DLTS spectrum. At 400K, all 

of PV parameters recovered a small amount except the FF which showed no change. Raising the 

temperature of illumination to 500K caused another moderate rise in Voc, but lac and the FF did not show 

an increase. Overall, the IV curves showed only minimal recovery. In particular, the complete lack of 

increase in the FF significantly suppresses P max' These results are similar in nature to the DLTS results. 

Illumination at 350K had no effect on the defect spectrum. At 400K, the H3 and H4 defects were 

completely removed, and at 500K, the H3' defect emerged (ref. 2). No other defect reactions were 

observed. The recovery of the IV curves is minimal and only two of the 6 radiation-induced defects show 

annealing stages. 

Annealing of a DHJ Under Illumination 

The second part of this study investigated the effect of the emitter structure on the annealing 

properties of irradiated loP solar cells. The first data set to be presented is the annealing of the RTI 

grown DHJ cell #1955-a4 under short circuit illumination. Since irradiated DJ cells showed substantial 

recovery due to short circuit illuminations below room temperature (ref. 1 and figs. 1 and 2), the first 

experiments were below room temperature illuminations. Initially, it was established through DLTS 

207 
~ 

E 

~ 

C 

~ :::J 

U 

10 MeV proton irrad n+p SHJ cell 

9 

8 

7 

6 

5 

4 

3 

2 

----------------. 500K 1 5 min 

.-._-._.-_ ... _-- 500K 75 min 

Annealings in the dark 

Pre-rod 

1 sun 

AMO 

298K 

<6.0 0 .1 0.2 0.3 0.4 0.5 0 .6 0 .7 0 .8 0.9 1.0 

Voltoge (V) 

Figure 8: Thermal annealing in the dark of Spire 

cell 5414-1-7_ Even at 500K, only limited recovery 

is seen especially in Voc and the FF. 

Am-241 oc-irrad n+p SHJ ",,=8x1 0" cm-2 

r---------~------~ __ 7 Pre-ra 

2 

1 sun, AMO 298 K 

<6.0 0.1 0.2 0 .3 0.4 0.5 0.6 

Voltoge (V) 

0.8 0.9 

Figure 10: Illuminated thermal annealing of the airradiated SHJ cell. Only moderate annealing 

occurs even after 5 hrs at 500K. The lack of 

recovery in FF suppresses P maX' 

208 

10 10 MeV proton irrad n+p SHJ cell 

",,=10'2 c m- 2 

H3 H4 

-10 

--- 0 

·················15 415 K 

'- '- '- '- '- '- " 1 60 

------------- 15 

._._ ._". __ . 75 500 K 

Temperature (K) 

Figure 9: Defect annealing in the SHJ cell of fig. 

8. Only the H3 and H4 defects anneal. Increase 

in negative signal is due to the removal of the 

positive signal and not an increase in defect 

concentration. 

150 Am- 41 ",-orrod SHJ cell H3' 

9'=8x1 0" em-' H3" H4 

'G:' 50 a. 

~ 

o 

c 

~-50 

til 

:; 

o 

-150 

-250 

/ / --Pre-anneol 

i I .............. ·350 K 2hr 20min ii' ......... _.-........ ' ... ,: ... ,,\,./ fr,~ _._._._._._-, 5 min 400 K 

EC ", i ;1 ·_-·- ·- .. _·5 hr 

" \ ji \ /\ \ Ii .------- 500K 3 hr 

v' '. \ ii 1 sun, AMO illuminotions 

EA \ \,/ I short eire uit \ ! 

\/ ED 

-350 ~-10 20 40 22 24 60 300 

Temperature (K) 

Figure 11: The defect annealing in the cell of fig. 

10. As seen during thermal annealing in the dark, 

only the H3 and H4 defects show any annealing 

stages. 

_I 
1---,--- 

------- ~ . 

measurements that illuminated IV measurements at 86K did not cause annealing. Then the cell was held 

at 175K and illuminated short circuit for 1 hour. A subsequent illuminated IV measurement at 86K 

showed no change (fig. 12). A subsequent DLTS measurement also showed no change (fig. 13) which 

is a very unexpected result because independent of PV recovery, the H4 defect has always shown 

annealing at temperatures of 175K and above (1-5). 

Following the 175K experiment, cell 1955-

a4 was illuminated short circuit at 300K and above. 

As shown in fig. 13, illumination at 300K did cause 

the familiar reduction of the H4 peak height. The 

H5 peak is seen to grow concurrently which is 

characteristic of the SHJ cells, but not the DJ cells 

3 MeV proton irrod. OHJ cell 

cp=6x1 0" c m- 2 

before irradiation 

(refs. 1,2, and 5). A very small recovery stage of 1 3 

the PV parameters may be apparent (fig 14 -note ~ c 

that since the annealing temperature is 300K or t 2 

larger, the IV measurements were taken at 298K). J 

lllumination at 375K caused significant recovery in 

the N curves, and illumination at 450K caused even 

more recovery which lead to complete recovery of 

Ioc. However, the recovery of Voc is quite limited. 

The changes in the DL TS spectra are shown in fig. 

13. The illuminations only affected the H3 and H4 

defects which is very similar to the SHJ cell results. 

3 MeV proton irrod. DHJ cell . 

6 c;o=6x1 0" cm-2 .' 

5 T-'=20s-' 

4 

l2 3 

-.9; 2 

H5 

o 

c 

~ 0 

~ -1 jt, / .' is -2 . r~J", ,;;,:.:-.. / /'''';' 1 sun illuminations '-Ij ~ ;;, ~ ,,"',- ... ..... 1 75K 1 hr 

3 " ~,: r , I - EC •..• / '\ f .. -... -.-..... 300K 2hr 

-4 ED 1EAi ---.. -------. 375K 2hr '. I -5 \j ... -.. -... -.. -. 450K 5 .25hr 

-6 ~-lO~O-1 ~14~O-'~6~0-'~8-0~~O~ 20 ~4 ~60~28~O~300 

Temp~roture (K) 

FIgure 13: Illumination ofRTI DHJ cell 1955-a4. 

At 175K, annealing of the H4 defect was expected 

but not seen. Above 300K, only the H3 and H4 

defects anneal while the H5 peak grows. 

1 sun AMO 86 K 

'b.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 O.B 0 .9 1.0 1.1 1.2 1.3 

Voltage (V) 

FIgure 12: Low temperature illumination of RTI 

DHJ cell 1955-a4. One hour of short circuit 

illumination at 175K causes no recovery. IV curves 

measured at 86K to avoid annealing during the 

measurement. 

.3 MeV proton irrod . DHJ cell 

SI'=6xl 0" c m- 2 

S ______ .. _. _____________ .. __ ~ __ :.::__> before 7diotion 

after irradiation ~ •• ~ . 

C .3 

t 

:::J 

U 2 

illuminated anneals 

........ 300K 2hr 

.............. .375K 2hr 

.--.. ---- 450K S.2Shr 

1 sun, AMO. 298K 

. \\ .\\ 

~ ~\ 

:'i 

\\ 

~\ 

:i 

:i 

H 

:= ., 

~\ 

III O 0.1 0.2 0.3 0.4 0.5 0.6 0.7" 0.8 0.9 1.0 

Voltage (V) 

FIgure 14: Continued illumination of RTI DHJ cell 

1955-a4 at T ~ 300K. Some recovery seems evident 

at 300K. lac completely recovers at 450K. Voc 

limits the cell recovery. 

209 

---------- ~ 
Thermal Annealing of a DR] in the Dark 

The final data sets to be presented are the thermal annealing of an irradiated DR] cell in the dark. 

The DR] cell #1955-a l was annealed in the dark (and unbiased) at 375 and 450K. Initially, the cell was 

illuminated at 300K to ensure all room temperature annealing stages were complete. The recovery of the 

cell is shown in fig. 15. All of the PV parameters recovered with lac recovering the most. The effect 

on the DLTS spectrum is shown in fig 16. The annealing removed R3 and R4 but did not affect the 

remaining defects. These results follow the same trends as the illuminated annealing data. 

I-~~~~==~====~~~~ RTI cell 01 

-." ------.-------.- --.- ----.. -- ~~ :~ ;;-.- ------.--. -. , ',-' .... 

. «E. 43 3 M.V :'~:~; ;;'~~%D,HJ ,." >':::~\\ _ --Pre-rod ", \\ 

c ············· ·Past-rad and 30m, 300K illum. \\\ 

t .. ·······--375K 1 hr dark ":\ 

:J 2 -··_··----450K: 1 hr dark ':\ 

u D 

1 sun AMO 298K 

~.O 0.1 0.2 0.3 

~ : ~ 

~ l \ 

~ :\ 

; ~! ::i 

,\\ 

0 .4 0.5 0.6 0.7 0.8 0.9 

Voltage (V) 

Figure 15: The effect of thermal annealing in the 

dark on RTI DRJ cell 1955-al. The recovery 

stages are similar to those seen during the 

illuminated annealing of cell 1955-a4. 

'-~~--T-~~--~~--r-~~--, 

4 3 MeV proton irrad. DHJ cell 

SO=6xl O"C m-2 

:3 T- 1 =205- 1 

RTI cell 01 

2 

/r""'·"'-·~\/~""'"\.vr-..... .-1 

.. ', "'\. , ! ofter injec tian ........ ,..,r 

\"/\ ED i at 300K 

-6 

' If \! anneal in dark 

\! 375K. 1 hr 't.,t "" 450K, 1 hr .. -··-··· __ ·-

-5 "EC 

-78 --:-! 1 """ -::- 0 "",""' 40:--' ,...,.-: -::- 0 ""' 20:--4'": 26"""-: -::- 0 ~:300 

Temperature (K) 

Figure 16: The effect thermal annealing in the 

dark on RTI DRJ ceIl 1955-al. Basically the same 

defect reactions are seen as in the illuminated 

annealing of cell 1955-a4. 

DISCUSSION 

The first question under investigation was - would illuminating an irradiated SR] cell above room 

temperature enhance the cell annealing characteristics? The answer is no. Essentially the same amount 

of annealing of the PV parameters was seen following annealing under illumination and in the dark, and 

the recovery is far from complete. Furthermore, the residual defect spectra measured at the end of both 

annealing experiments are essentially the same. As was predicted by the results of ref. 2, very little solar 

cell recovery is seen in the SRJ cells after annealing at temperatures below 500K, regardless of the 

illumination. 

These results provide a better understanding of the cause of the different annealing behavior in 

DJ and SRJ solar cells. In a DJ cell, illumination at 450K removed all of the radiation-induced defects 

from the cell junction except for the electron trap, ED, which only partially annealed after 2 hours at 

450K. In a SRJ cell, on the other hand, below 5OOK, none of the defects show annealing stages except 

the R3 and R4 defects. These two defects are removed in the range of 375-450K, but all others remain. 

210 
Therefore, the enhanced annealing evident in the DJ solar cells can be attributed, at least in part, to the 

enhanced defect annealing in the cell junction. Considering these results with those of ref. 2, it is 

expected that the defects in the junction of a SHJ cell will anneal if the temperature is raised to 6S0K. 

Therefore, it may be concluded that a SHJ solar cell will show complete recovery after annealing at 

6S0K. If this is the case, then it can also be concluded that, irradiated n+p InP solar cells grown by S 

diffusion will completely recover at a temperature which is 200K lower than an irradiated n+p InP SHJ 

cell grown by MOCVD. This is a strong possibility given the present results, but since data from an 

irradiated n +p SHJ cell annealed above SooK do not yet exist, no definite conclusions can be drawn. 

The second question under investigation was if the difference in emitter structure between the SHJ 

and DJ cells was responsible for the different cell annealing behavior. The enhanced annealing of the 

DJ cells was above shown to be, in part, due to enhanced defect annealing. Since the radiation-induced 

defects in SHJ InP have been shown to be sensitive to an electric field (ref. 6), this enhancement may 

be an electric field related effect. The DHJ cells were grown to simulate the DJ cells, so the junction 

electric field is expected to be similar in both cell types. However, the residual defect spectrum measured 

in both DHJ cells (Le. in both the cell annealed in the dark and the cell annealed under illumination) was 

essentially the same as that of the SHJ cells. Therefore, the difference in the junction electric field of 

the DJ and SHJ cells most likely is not the cause of the enhanced defect annealing. However, it should 

be noted that the exact structure of DHJ and the DJ cells has not yet been analyzed. Therefore, these 

conclusion can orJy be tentative. A more precise determination of the structure of these devices is the 

next step in the NRLIRTI research project. 

The annealing of the DHJ cells is also similar to that of the SHJ cells in terms of the recovery 

of Voc. Voc shows almost no recovery in both of these cell types. The DJ cells, on the other hand, have 

shown complete recovery in Voc. This is almost certainly due to the fact that most of the radiationinduced defects in the base of the DHJ cells are not removed by the present annealing experiment while 

in the DJ cells, all but the ED defect are removed. The persistence of the radiation-induced defects in 

the epitaxial cells following these annealing experiments inhibits the recovery of Voc while the annealing 

of the defects in the diffused junctions induces full recovery of Voc in the DJ cells. 

While the annealing of the defect spectrum in the DHJ cells is similar to that of the SHJ cells, 

the recovery of the PV parameters shows a significant difference. Annealing a DHJ cell for 5.25 hrs 

under illumination at 4S0K caused complete recovery of I"" and annealing a DHJ for 1 hour in the dark 

at 450K showed substantial I.., recovery. The SHJ cells did not show this much recovery, so it does seem 

that the emitter structure of the DHJ cells has enhanced the annealing characteristics of I",. The DLTS 

results suggest that this is not due to enhanced defect annealing in the base (DLTS samples only the base 

region), so the enhanced recovery it most likely due to an annealing mechanism in the n-type emitter. 

This result is consistent with work presented at this conference by Messenger et al. on irradiated p+n InP 

solar cells. Those results show recovery in lac due to short circuit illumination at room temperature and 

below. It seems, now, that a possible explanation has emerged for why I.., in the DJ cells recovers under 

illumination at T < 300K and J", in the SHJ cells does not. It may be that the I.., recovery is caused by 

an increased carrier collection efficiency in the n-type InP due to the illumination and only the DJ cells 

have enough n-type material for the effect to be seen. 

CONCLUSIONS 

The present research has shown that illuminating an irradiated n +p SHJ InP solar cell during 

thermal annealing above 300K does not enhance the recovery of the PV parameters or the defect 

annealing rate. The results strongly suggest that irradiated n +p DJ solar cells will anneal at a temperature 

211 

-- ~ 
200K lower than an irradiated n+p SHJ cell with or without illumination. The measured full recovery 

of I.e in an irradiated DHJ cell has indicated that illumination may be only affecting the current collection 

in the n-type InP material, so only n+p cells with thick emitters, Le. the DJ and DHJ cells, show 

recovery of I.e under illumination. However, the fact that the lac of the DHJ cells did not recover under 

illumination below room temperature shows that the present DHJ cell structure still lacks qualities of the 

DJ cell structure which are essential for optimizing this effect. Furthermore, the lack of recovery of Vex: 

and the lack of defect annealing in the DHJ cells indicates that there is still a major difference between 

the epitaxial and diffused cell structures which inhibits the epitaxial cells from obtaining full recovery of 

radiation-induced damage. 

These points clearly indicate that while the present investigation of the effects of the cell structure 

on the cell annealing characteristics has indicated the importance of the n-type emitter, more research into 

the device structure needs to be done to fully understand the mechanism for the enhanced annealing 

properties of the DJ solar cells. 

REFERENCES 

1. Walters, R.I., G.P. Summers, and J. Bruening, "A Detailed Study of Thermally Diffused InP 

Solar Cells", XII Space Photovoltaic Research and Technology Conference, p.1, Oct 20-22, 1992 

2. Messenger, S.R., R. J. Walters, and G.P. Summers, "High Temperature Annealing of Minority 

Carrier Traps in Irradiated MOCVD n+p InP Solar Cell Junctions", XII Space Photovoltaic 

Research and Technology Conference, p.8, Oct 20-22, 1992 

3. Walters, R.J. and G.P. Summers, "Deep Level Transient Spectroscopy Study of Proton Irradiated 

p-type InP", J. Appl. Phys. 69 (9), 6488, 1 May 1991 and references therein 

4. C.I. Keavney, R.J. Walters, and P.J. Drevinsky, "Optimizing the Radiation Resistance of InP 

Solar Cells: The Effect of Dopant Density and Cell Thickness", J. Appl. Phys., 73 (1),1 Jan. 

1993 p 60-70 

5. Walters, R.I., "A Study of the Annealing of Radiation-Induced Defects in loP Solar Cells", 

Ph.D. Dissertation, University of Maryland Baltimore County, May 1994 

6. S.R. Messenger, R.I. Walters, G.P. Summers, "Effect of Carrier Concentration on the Properties 

of Irradiation-induced Defects in p-type Indium Phosphide Grown by Metalorganic Chemical 

Vapor Deposition", 1. Appl. Phys. 71. p.4201-4207, 1 May 1992. 

212 
N95-20525 

QUALIFICATION TEST RESULTS FOR BLUE-RED REFLECTING SOLAR CELL COVERS 3t:J ??~ r 

W.T. Beauchamp 

GCLI 1'0 ('-

Santa Rosa, California 

Recent market forces and design innovations have spurred the development of solar cell covers that 

significantly reduce the solar absorptance for a cell array. GaAs cells using Ge as the substrate host 

material, can have a significantly higher output if the solar absorbtance of the cell array is reduced. 

New optical coating design techniques have allowed the construction of covers that reflect the 

ultraviolet energy (below 350 nm) and the near infrared energy (above 900 nm) resulting in the 

beneficial reduction in absorbtance. Recent modeling' suggests three or more percent output increase 

due to the lowered temperature with such a device. 

Within the last several months we have completed the testing of production samples of these new 

covers in a qualification series that included the usual environmental effects associated with the routine 

testing of solar cell covers and the combined effects of protons, electrons and solar UV as would be 

encountered in space. 

For the combined effects testing the samples were exposed to 300 sun days equivalent UV, 

5 x 1 014/cm2 of 0.5 MeV protons and 1 0'5/cm2 of 1.0 MeV electrons. Measurements of the reflectance, 

transmission, emittance and other appropriate parameters were made before and atter the testing. As 

measured by the average transmission over the cell operating band, the change in transmission for the 

samples was less than or about equal to 1 %. 

The details of the testing and the results in terms of transmission, reflectance and emittance are 

discussed in the paper. 

Key Words: 

Space Power, Photovoltaic System, Space Qualification, Solar Reflector, Reduced Solar Absorptance, 

Temperature Reduction, Output Efficiency 

INmODucnON 

The results of this series of tests give quantitative data on the nominal performance of blue-red 

reflecting (BRR) solar cell covers designed for use on GaAs and silicon solar cells before and atter 

exposure to the normal environmental tests for such devices, as well as the stability of that 

performance in simulated space environmental effects testing. The tests performed included the 

normal application specific performance parameters and environmental exposures that are important to 

the use of the devices in service. 

The test program used samples prepared using standard production methods and procedures. The 

samples were manufactured in production lots from at least two different coating runs for each coating 

type and were divided into test groups in a randomized manner with each group being exposed to a 

specific series of tests and analyses. Many of the tests and characterizations were performed at OCLI. 

However, the space environmental exposure testing using UV and high energy protons plus electrons, 

was done at the Boeing Corporation. Low energy proton exposures and evaluations were done by the 

Hughes Aircraft Co., Space and Communications Division, and by the Martin Marrietta, Astro Space 

Division. Samples prepared for this test series were mostly2 characterized at OCLI for the appropriate 

performance parameters before testing and atter testing to assess the changes induced by the 

environmental exposures. 

, See BluelRed Reflecting Solar Cell Covers for GaAs Cells; Proceedings of the Twenty Third IEEE 

Photospecialist Conference; Louisville, KY; May 1994 

2 Some pre and post measurements of performance, as well as, the particle exposures were done by Martin 

Marrietta for the low energy series. 

213 
Product Descriptions and Codes 

The products explicitly tested were blue-red reflecting solar cell covers for GaAs and silicon cells. For 

Cell, the basic code used to designate a blue-red reflecting cover is BRR. If the cover is intended for 

use on a GaAs cell then a (g) is appended to this code to result in the BRRlg designation. If the cover 

substrate material is Coming 0213 glass - a cerium doped micro sheet material- then the designation 

becomes BRRlg-0213. Similarly, if the application requirements dictate the use of fused silica as the 

substrate material and the cell will be GaAs the designation would be BRRlg-FS. A BRR cover for use 

on a silicon solar cell would be deSignated as a BRRls-0213 or BRRls-FS. Another version of the 

basic conceptual design for these products deletes the UV reflection feature and substitutes an AR. 

This version would be designated as an AR-RRlg-0213, etc. 

In this qualification series, samples of the BRRlg-0213 and the BRRls-0213 covers were tested and 

evaluated. 

Product Descriptions 

The function of coatings on a reflecting solar cell cover is to reduce the solar absorption of the array 

with little or no change in the electrical output of the system. The reduction in solar absorption is 

accomplished by reflecting as much as possible of the incident solar energy that is at wavelengths 

outside of the response wavelength band for the cell. GaAs solar cells are responsive to energy at 

wavelengths from about 350 nm to about 900 nm. For silicon solar cells the response region is from 

about 350 nm to about 1100 nm. If the cover is designed to reflect both the UV energy below 350 nm 

and the near-infrared energy above 900 nm or 1100 nm, the lower solar absorbtance will allow the 

system to remain cool and the cell efficiency will be less degraded by temperature effects. Of course, 

the reduction in solar absorbtance will result in a lower temperature only if there is no significant 

change in the emittance of the array surface due to the cover and coating. 

A typical BRR cover reflectance for both types of cover - for GaAs and silicon cells - is shown in Fig. 1. 

In the response regions of the cells, the transmission is very high. In the out of band regions above 

900 nm or 1100 nm and below 350 nm, the reflectance is as high and as broad as practicaJ.3 If the 

solar absorbtance is measured for a BRR cover adhered to a typical GaAs cell, the nominal solar 

absorbtance is on the order of 75% with this BRRlg-0213 design. (The absorption of the cell and the 

back contacts, as well as the glass, coating, and cement are included in the solar absorbtance value for 

the complete cell-cover system.) 

By comparing the spectral plots in Figure 1, the difference between a GaAs BRR and a silicon BRR can 

be seen to be in the breadth of the transmission band. The poSition of the reflector for the NIR 

wavelengths is independent of the position of the reflector for the UV wavelengths. In addition, for the 

changes necessary to adapt from a reflector for use on GaAs to one for use on silicon, the transmission 

at wavelengths in the cell response bands can be kept very high. Therefore, the transition edge - from 

transmitting to reflecting - can be independently placed at a best position for each cover. The value of 

1100 nm for silicon and 900 nm for GaAs on the long wavelength side of the transmission band and 

350 nm for the short wavelength edge for both cell types was chosen based on extensive testing of 

celVcover combinations for use on pointing array~ . 

The typical construction for a BRR cover is to place the UV reflector on the front surface of the cover 

and the red reflector on the rear surface. Because the 0213 glass is highly absorbing at wavelengths 

less than 350 nm, the UV reflector must be on the front to be functional. The red reflector could also be 

on the front of the cover (under the UV reflector) but this might result in a significant reduction in the 

emittance for the cover. The coating(s) on the front surface will modify the emittance due to the optical 

properties of the materials in the far infrared. In particular the materials in the coating can add to 

3 There is a great flexibility in the width of the reflectance band for the cover in both the UV and the NIR. What is 

practical is determined by a host of fadors induding the reduction in solar absorbtance for broader reflectance 

measured against increased cost of manufacture, the possible effect on transmission over the cell response 

region associated with the more complex design. and possible greater warpage of the cover due to stress in the 

coatings as well as other technical factors. 

4 For the use of a BRR on a spinner satellite power system, the effects of angle of incidence on the coating must 

be taken into account This results in the use of a NIR reflecting band positioned slightly farther out in 

wavelength than for the pointing array case. 

214 
Reststrahlen reflectance or Fresnel reflectance in the Infrared resulting in a reduced emittance. This Is 

particularly true for moderately thick coatingsS such as the red reflector. 

Figure 1. 

--IILO 

... 

'I'ILO 

I l1/.li 

IILO 

-.0 

IILO 

IIUI 

- V 

\ 

411 

Nominal Reflectance as a Function of Wavelength for a BRRlg-0213 Solar Cell 

Cover (a) and for a BRRls-0213 Solar Cell Cover (b). 

1/""'\ IIUI 

... r 

1\ 

I 

..., I \ .... \ 

IIUI \ 

..... 

IIUI 

r ... / 

\ / JIL I 

.-l .... "- J ~ 

u lUI ~ loG 110 Ul UlOCUI J-. uo l.8 .a ... ~ ......... MOerJf uoo.e waaa ....... 

HANDWiJx .... 

Note: These scans have different wavelength scales. 

The covers evaluated in this test series were all constructed on 0.006 inches (6 mils) thick Coming 

0213 cerium doped microsheet. The 0213 glass was taken from stock material used at OCLI for 

routine manufacturing of solar cell covers and qualified for space use as a cover material, uncoated 

and in conjunction with a UVR coating on the front, in a qualification test discussed in the 1989 OCLI 

Qualification Report APD 89011. 

Part Preparation and Sample Selection 

The parts tested for this qualification were prepared using manufacturing procedures for the particular 

design, and were deposited in equipment regularly used for manufacturing of covers as well as other 

products. The samples were, as much as-practical, selected at random from two separate lots and/or 

two separate coating runs for each coated surface on the covers. Control samples in the tests included 

uncoated 0213 glass and the UVR covers. These control samples showed changes consistent with, 

and in statistical agreement with, the data contained in the OCLI qualification test report for the UVR's 

and Coming 0213 microsheet glass (report APD 89011). 

TESTS PERFORMED AND QUALIFICATION TEST GROUPS 

The parameters measured and the environmental tests performed are shown in Table 1. This list of 

evaluations dictated the sequence of tests for each group and the data to be gathered. 

Because of limited space in the test facilities at Boeing, the number of samples that could be exposed 

to combined space effects was further limited. To insure statistical validity in the results, a minimum of 

5 samples of each design type were tested in two independent sets. 

Facilities and Equipment Used for the TestIng and Exposures. 

For the measurements of the optical performance properties of the covers before and after the 

exposures, a Shimadzu 160 spectrophotometer was the primary instrument used. For measurements 

of the infrared reflectance, a Perkin Elmer 983 was used (thiS was primarily for the emittance 

determination). The regular environmental exposures for humidity and other durability tests were 

performed in the CClI testing laboratory to the military specifications pertinent to the individual tests. 

5 The red reflector is of the order of 5 to 6 times as thick as the UVR 

215 
I 

I 

I 

L 

Table 1 Qualification Test Matrix for the BRRlg-0213 (GaAs BRR) and the BRRls-0213 

(Silicon BRR) Qualification Series. 

Test GrOUp 

REQUIREMENT A B e 0 E F G H I 

OTY S+S S+5 5+5 S+S S+5 5+5 S+S 4+4 

2an 2an 2an 2an 2an 2an 2an 2an ested 

x x x x x x x x by 

SIZES 4an 4cm 4em 4em 4cm 4em 4an 4cm Qtt1ers 

Surface Quality; Workmanship; 80-SO reQuirement X X X X X X X X X 

CoatlnQ Orfentatlon X X X X X X X X X 

Noonal Enittance (1- R 5 to SO microns) X X 

~nlCutoff Wavelength X X X X X X X X 

Ultravlofet ... .... • ~ 85% @ 300 nm +/. 20 nm X X X X X X X X 

Transmft1ance X X X X X X X X 

Humidity A·72 hours@12O"F' ~ 95% R.H. X 

Humidity B·10 days per MIL·STD-810B X 

Thermal Shock·LN21 hr. & 35O"F 1 hr. X 

Tempera1Ure eycfe·l 000 cydes; -180oe to +19Soe X 

Salt Fog-48 hours X 

Ultraviolet Exposure-300 sun days X 

Radiatfon Reslstance-20 to 30 KeV protons X 

Radiation Reslstance·SOO KeV protons X 

Radiation Reslstance-1 MeV electrons X 

Abrasion Reslstance-2O rub X X X 

Adheslon-Slow tape X X 

lS' BofllSlowtape X 

1 .... - ~ r.lltnn )( )( )( )( )( )( x. X 

The radiation and UV exposures were done at the Boeing Company Physical Sciences Laboratories 

using their X-200 Solar Simulator (at Organization 9-5574, Boeing. Renton. WA) and the Dynamitron 

Accelerator at the Physical Sciences Research Center (Seattle. WA). The 1200 total hours UV 

exposure started on November 1. 1993 and finished on December 23. 1993. The UV exposure was 

done at about 6x the solar intensity (Air Mass Zero (AMO) Solar Spectrum) for a resultant 300 sun days 

equivalent exposure. This was followed by the exposure at the Dynamitron to 0.5 MeV protons to > 5 x 

1 o 14/cm2 and then to 1 MeV electrons to >101S/cm2. 

Radiation exposure at low energies (20 KeV to 30 KeY) for protons was done by the Astra Space 

Division of the Martin Marietta Co. and the Space and Communications Division of the Hughes Aircraft 

Co. 

During exposure the radiation test samples were kept at or near room temperature by actively cooling 

the mounting plates with flowing water. 

Measurement Error Analysis 

Repetitive measurements of known wavelength and photometric standards have established the 

following measurement errors for transmission: 

Wavelength: 

Photometric: 

Accuracy 

± 0.5 nm 

0.05% 

216 

RepeatabilitY 

±0.1 nm 

0.02% 
QUALIFICATION TEST RESULTS FOR THE BRR/GaAs COVERS 

All parts were measured for transmittance, UV reflectance, and NIR reflectance before test and met the 

oeLl Product SpecifICation No. 6067001-08. 

Before and after measurements of transmittance were done for all test groups except group F. Group 

F only Involved observations of the surface quality and coating orientation for the samples. 

Measurements of reflectance are included in the data after testing for completeness, but no data was 

taken for this group before. 

Only the samples in test group H were measured before and after the test for emittance as well as for 

the transmittance values. The values for the emittance of the covers were calculated as one minus the 

reflectance in the range 5 ~ to 50 ~. 

Summaries of the before measurements of transmittance for the BRRlg-0213 test samples are shown 

in Table 2. The values of the arithmetic average transmission for various wavelength ranges are 

shown in the table. Data is presented for each of the test groups (however, the thermal cycle tests -

Group D - have not been completed as of this writing). 

The after exposure transmittance measurements for each of the test groups is shown in Table 3. In 

this data set the values for the transmittance of the group F samples are also included, as measured 

after test, for reference. The changes in transmittance. absorbtance, reflectance and emittance for the 

test sample sets in each group are given in Table 4. 

The surface quality of all test parts in the test groups was ~ 80-50 per MIL-0-13830A. 

The normal spectral emittance for the BRRlg covers was found to be unaffected, well within the 

accuracy of determination of the emittance, by the combined effects testing. Nominal values for the 

normal spectral emittance as determined by these measurements, were in the range of 86% to 87% 

The average transmittances in the 400 nm to 900 nm wavelength band were also unaffected by the 

normal environmental exposures. 

The 50% cut-on wavelengths (transmission) of all BRR/GaAs coatings were 359.5 nm ± 1.7 nm at the 

UV position and 972 nm ± 7 nm at the near infrared positiot'!. All measured BRRlg-0213 parts had less 

than 1% average transmission between 200 nm and 320 nm before and after testing. 

The BRAlg-0213 parts in test group H were subjected to combined UV exposure for 300 AMO UV sun 

days equivalent (at 6x sun intensity), followed by proton irradiation (0.50 MeV) to 5 x 101<4/cm2, in a 

vacuum of 1 x 10.6 torr and then electron irradiation (1.0 MeV) to 1.0 x 101S/cm2. The average 

transmission from 400 nm to 900 nm was lowered an average of 1.12%. This change is a little higher 

than was expected. However, it is near the change limit specified for other cover products such as the 

UVR only. The specification limit for change in solar transmittance is usually set at less than 1%. The 

change in the cut-on edge at the UV was about 5.6 nm on average and at the NIR edge was 0.69 nm 

on average. 

Summary of Qualification for the BRR/GaAs 

For all qualification test groups discussed above, transmission changes due to the testing were less 

than 1%, except for the samples irradiated in test group H. The change in transmission for the samples 

in test group H was about 1.12% on average. 

QUALIFICATION TEST RESULTS FOR THE BRR/SILICON COVERS 

The matrix of qualification test groups for the BRRlsilicon (BRRls-0213) covers is identical to the matrix 

for the BRR/g-0213 covers shown in Table 1. All parts were measured for transmittance, UV 

reflectance, and NIR reflectance before test and met the oeLl Product SpeCification No. 6067001-09. 

As with the other test sets, the before and after measurements of transmittance were done for ail test 

groups except group F. Group F only included observations of the surface quality and coating 

orientation for the samples. Measurements of reflectance are included in the data after testing for 

completeness, but no data was taken for this group before. 

217 
Table 2 T ransmlttance & Emittance for the BR RI fa· -02 1 3 S amples Before Testlna. 

(Summary - Set 1 Run Numbers 48~168611688) 

Test 50% Trans Cuton Peak at Avg Avg Emitt 

Group Lower Upper 200-230 400-900 600-800 

A Average 359.86 960.96 0.61 200.60 96.42 97.62 

Std Dev 1.25 5.91 0.05 0.55 0.21 0.05 

B Average 360.02 949.34 0.57 201.00 96.36 97.70 

Std Dev 1.66 5.43 0.01 0.00 0.04 0.04 

C Average 357.96 975.82 0.60 200.60 96.53 97.69 

Std Dev 0.64 15.02 0.06 0.55 0.10 0.12 

0 Average 358.60 975.66 0.52 200.20 96.41 97.68 

Std Dev 0.31 12.59 0.02 0.45 0.47 0.10 

E Average 359.24 989.92 0.54 200.00 96.52 97.83 

Std Dev 1.67 4.67 0.05 0.00 0.03 0.05 

F Average No Data Taken 

Std Dev 

G Average 358.38 958.58 0.54 200.00 96.65 97.68 

Std Dev 0.96 7.17 0.05 0.00 0.11 0.03 

H Average 359.25 961.33 0.58 200.00 95.47 96.68 86.57 

Std Dev 0.52 3.26 0.02 0.00 0.28 0.24 0.27 

Table 3 T ransmlttance & E I h B m ttance for t e Rig· RI-023S 1 amples After T est ng. 

(Summary - Set 1- Run Numbers 466-168611688) 

Test 50% Trans-Cuton Pea!§, at Ayg Avg Emitt 

Group Lower Upper 200-230 400-900 600-800 

A Average 360.16 960.28 0.13 320.00 96.28 97.57 

Std Dey 1.32 5.87 0.03 0.00 0.03 0.03 

B Average 361.60 948.88 0.09 320.00 96.31 97.69 

Std Dev 2.47 5.90 0.02 0.00 0.06 0.03 

C Average 358.02 976.14 0.14 320.00 96.46 97.67 

Std Dev 0.46 15.65 0.02 0.00 0.10 0.11 

0 Average TBD TBD TBD TBD TBD TBD 

Std Dey 

E Average 359.44 990.32 0.12 320.00 96.28 97.59 

Std Dey 1.58 4.38 0.00 0.00 0.23 0.25 

F Average 359.50 971.08 0.47 320.00 96.37 97.62 

Std Dev 0.31 1.20 0.01 0.00 0.23 0.19 

G Average 358.48 958.40 0.15 320.00 96.49 97.57 

Std Dey 0.97 7.76 0.00 0.00 0.10 0.03 

H Average 364.48 962.00 0.25 320.00 94.27 96.39 86.25 

Std Dey 0.45 .030 0.00 0.00 0.12 0.03 0.19 

• Set 1 of two sets. Second set not displayed to save space. Data and conclusions are the same for Set 1 and 

Set 2. For an expaination of the TBD's in group D see the text. 

218 
I 

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I 

I 

I 

I 

l 

I 

I 

I 

I 

I 

I 

I 

I 

- --- - ---

Table 4 Ch an~es I T n ransm ttance an dE m Itt ance f h or t e BR R/ ta-()213 S amples 

Set 1 Run Numbers 466-168611688 

Test 50% Trans Cuton Peak Avg Avg Emitt 

Group Lower Upper 200-320 400-900 600-800 

A Average -0.30 0.68 0.49 0.14 0.05 

B Average -1 .58 0.46 0.48 0.06 0.01 

C Average -0.06 -0.32 0.46 0.06 0.02 

0 Average TBO TBO TBO TBO TBO 

E Average -0.20 -0.40 0.42 0.24 0.24 

G Average -0.10 0.18 0.39 0.16 0.10 

H Average -5.23 -0.67 0.33 1.21 0.29 I 0.32 

.. A negative number signifies an Increase In value 

Space Environmental Tests 

The BRRls-0213 test group H was also the sample set sent to Boeing for particle irradiation. This test 

group was measured before and after test for emittance as well as for the transmittance values. The 

values for the emittance for the covers are calculated as one minus the reflectance in the range 5 J,lln to 

50 J,lln. 

Summaries of the before measurements of transmittance for the BRRls-0213 test samples are shown 

in Table 5. The arithmetic average transmission for various wavelength ranges are shown in the table. 

(as mentioned before, the thermal cycle tests - Group 0 - have not been completed as of this writing). 

The values for the reflectance of the cove~ in the UV and the position of the maximum value are also 

given. Data is presented for each of the test groups (the test group I has not been completed as of this 

writing.) Values for the emittance of the covers calculated as one minus the reflectance in the range 

5 J,lln to 50 J,lln, are also shown for each cover in the H test group, as measured before testing. 

The after exposure transmittance measurements for each of the BRRls-0213 test groups are shown in 

Table 6. Again the values are the average or minimum transmittances in the selected wavelength 

ranges. In this data set, the values for the transmittance of the group F samples are also included, as 

measured after test, for reference. The change in transmittance, absorbtance, reflectance and 

emittance for the test sample sets in each group are given in Table 7. 

The surface quality of all BRRls-0213 te::;t parts in the test groups is ~ 80-50 per MIL-0-13830A. 

All parts were measured for transmittance before test and met the OCLI Product Specification 

No. 6067001-09. 

The normal spectral emittance for the BRRls-0213 covers was found to be unaffected, well within the 

accuracy of determination of the emittance value, by the combined effects testing. Nominal values for 

the normal spectral emittance as determined by these measurements was in the range of 86% to 87<'10 

The average transmittances in the 450 nm to 1100 nm wavelength band were also unaffected by the 

normal environmental exposures. 

The 50% cut-on wavelengths (transmisSion) of all BRRlSilicon coatings were 359.5 (± 15) nm at the UV 

position and 1160 (± 10) nm at the near infrared position. BRRls-0213 test covers were placed into a 

test chamber at a control/ed temperature for 72 hours at a temperature of 49°C ± 2.5°C and > 95% 

• Set 1 of two sets. Second set not displayed to save space. Data and conclusions are the same for Set 1 and 

Set 2. 

219 
relative humidity per MIL-C-67SA (test group A). Mer humidity all parts were tested for 20 rub eraser 

abrasion resistance per MIL-C-67SA and slow tape adhesion per MIL-M-13S08B. No physical 

degradation was observed on any part and the change in transmission for the wavelengths 450 nm to 

1100 nm was less than 0.S7% on average. The change in the SO% transmittance cutoOn edge was less 

than about 0.4 nm to 1.2 nm on average at the W position. 

The BRR/s-0213 parts in test group H were subjected to combined UV exposure to 300 AMO UV sun 

days equivalent at 6x sun intensity, followed by O.S MeV proton irradiation to S x 1014/cm2 in a vacuum 

of 10-45 torr and then 1.0 MeV electron irradiation to 101S/cm2. The average transmission from 450 nm 

to 1100 nm was lowered an average of 0.36% .. The change in the cutoOn edge at the UV was about 3.4 

nm on average and at the NIR edge was about 5 nm on average. 

It is Interesting to note that the design of the BRR/silicon coating is the same as the design for the 

BRRlGaAs design, with the exception of the increase in the red reflector coating thickness to shift the 

reflectance band out beyond 1.1 Jl1T1. This represents a thickening of all of the layers in the design by 

about 15%. If the changes shown are rofJ1)ared between the two designs, however, it is not dear why 

the GaAs version should have changed so rnJch while the silicon version changed so little. 

Summary of Qualification for the BRR/Slllcon 

For all qualification test groups in the siliron BRR evaluations, the change in transmission was less 

than 1 %. This is within the specification for the BRR rovers and the BRRls-0213 covers can therefore 

be qualified for space applications. 

aUALlACAnONS BY SIMILARITY 

Based on the results of this testing and flight experience with other cover/coating rombinations, it is 

possible to qualify other specific products by similarity. These include the blue-red reflecting solar cell 

covers using fused silica as the substrate and an AR-red reflector cover. 

Other Versions of The Red Reflecting Solar Cell Cover 

Among the alternate product forms for red reflecting solar cell covers are products that substitute other 

substrate materials such as fused silica, and product fonns that substitute an antireflective coating on 

the front surface for the UVR. 

As noted in the introduction, the version ot the design that substitutes fused silica for the substrate is 

designated by appending an FS to the product code in place of the 0213. The difference between the 

BRRlg-FS and the BRRlg-0213 and between the BRRls-FS and the BRRs-0213 solar cell covers is the 

substrate material and the addition to a small amount of UV absorbing material in the front coating (the 

UV reflector). The 0213 glass, qualified in the testing described in oeLl report no. APD 89011, 

changes very little when exposed to the particle irradiation, the other tests described in that report, and 

in this series of tests. However, fused silica is even more stable in these same tests. Therefore, the 

use of fused silica as the substrate would result in an even more stable cover product. 

The coating deSign and material used for the front surface roating on the BRRIx-FS product is also a 

space qualified product that has flown for many years. This coating on fused silica is the blue reflector 

used in the ARlBR deSign that was the staple of the solar cell cover market for the years prior to the 

introduction of the Ce02 stabilized glasses such as Coming 0213. 

The reflectance and transmittance for these product forms are nearly identical to the glass based 

products. The solar absorbtance for the fused silica based BRR products is a little less than that for the 

glass based product. 

220 

I 

I 

I 

I 

I 

I 

I 

I 

I 

~ 
I 

I 

I 

I 

I 

Table 5 T ransmlttance and Emittance for the BRRI s-02 13 S amples Before Testlna. 

Set 1 Run Numbers 466-168511688 

Test 50% Trans Cuton Peak at Avg Emitl 

Group Lower Uooer 200-230 450-1100 

A Average 363.36 (NO) 0.71 200.60 96.49 

Std Dev 0.38 0.05 0.55 0.07 

B Average 364.68 (NO) 0.73 200.40 96.57 

Std Dev 1.11 0.05 0.55 0.08 

C Average 363.02 (NO) 0.66 200.80 96.55 

Std Dev 2.88 0.02 0.45 0.06 

0 Average 362.54 (NO) 0.70 201.00 96.64 

Std Dev 1.85 0.04 0.00 0.14 

E Average 366.72 (NO) 0.72 200.60 96.72 

Std Dev 2.43 0.05 0.55 0.08 

F Average No Data Taken (NO) 

Std Dev 

G Average 365.84 (NO) 0.68 200.60 96.68 

Std Dev 2.76 0.05 0.55 0.12 

H Average 364.40 (NO) 0.59 200.00 95.84 86.28 

Std Dev 1.44 0.01 0.00 0.15 0.06 

Table 6 T ransm Itt ance an dE m Itt ance f h ort e BRRI s-0213 S amples Aft er T est na. 

Set 1 Run Numbers 466-168511688 

Test 50% Trans Cuton Peak at Avg Emitt 

Group Lower Upper 200-230 450-1100 

A Average 363.78 1209.02 0.15 320.00 96.12 

Std Dev 0.40 4.79 0.03 0.00 0.19 

B Average 365.28 1212.28 0.10 320.00 96.35 

Std Dev 1.03 10.79 0.02 0.00 0.43 

C Average 363.10 1214.48 0.12 320.00 96.43 

Std Dev 3.03 13.45 0.03 0.00 0.11 

0 Average TBD TBD TBD TBD TBD 

Std Dev 

E Average 366.74 1202.72 0.12 320.00 96.67 

Std Dev 2.46 1.32 0.04 0.00 0.09 

F Average 364.40 1227.92 0.36 320.00 96.67 

Std Dev 2.60 7.62 0.06 0.00 0.24 

G Average 366.34 1228.08 0.10 320.00 96.59 

Std Dev 3.08 10.32 0.03 0.00 0.15 

H Average 368.45 1182.05 0.30 320.00 95.99 86.29 

Std Dev 2.47 1.16 0.01 0.00 0.11 0.50 

• Set 1 of two sets. Second set not displayed to save space. Data and conclusions are the same for Set 1 and 

Set 2. For an 8xpaination of the TED's in group D see the text. 

221 
Table 7 Changes In Transmittance and Emittance f ort h e BRRls-0213 Sa mples 

Set 1 Run Numbers 466-168511690 

Test 50% Trans Cuton Peak Avg Avg Emitt 

Group Lower Upper 200-320 450-1100 400-450 

A Average, -0.42 (NO) 0.56 0.36 0.12 

B Average -0.60 (NO) 0.63 0.22 -0.30 

C Average -0.08 (NO) 0.55 0.11 -0.15 

0 Average TBO (NO) TBO TBO TBO 

E Average -0.02 (NO) 0.60 0.05 0.27 

G Average -0.50 (NO) 0.58 0.09 -0.40 

H Average -4.05 (NO) 0.29 -0.15 5.45 I -0.01 

.. A negative number signifies an Increase In value. ND Indicates no data due to no before measurements 

of the cut on edge. For an expaination of the TBD's in group D see the text. 

. " .... 

The second alternate product form with an AR on the front has the designation AR-RRlg-0213 when 

the coatings are on the Coming 0213 microsheet glass. This product differs from the BRRlg-0213 

product because the front coating is the conventional single layer magnesium fluoride antireflection 

coating that has flown in space for many years. In the same manner, the difference between the 

AR-RRls-0213 product and the BRRls-0213 cover is the same single layer antireflection coating 

design. 

It may be possible, because of the flight history for parts of these designs and the results of the testing 

discussed here, to consider these alternate product forms qualified by similarity. Due to the high cost 

associated with such testing and the time required for a test program, it is hoped that the similarity of 

the products to the BRRlg-0213 and BRRls-0213 will be sufficient for qualification. However, if there is 

sufficient justification for the specific testing of any of these alternate product forms, oeLi will entertain 

the suggestion that a test series be conducted for the AR-RR. 

CONCLUSION 

Based on the testing discussed here, the blue-red reflecting solar cell covers can be considered 

qualified for space use. 

Set 1 of two sets. Second set not displayed to save space. Data and conclusions are the same for Set 1 and 

Set 2. 

222 

._-- ------ -
l 

37'777(.} 

N95-20526 

UV TESTING OF INTELSAT-VII, VilA, AND VIII SOLAR CELLS1 

A. Meulenberg 

COMSA T Laboratories 

Clarksburg, Maryland 

Extended Abstract 

A 4000 hour experiment, conducted in late 1992 through mid 1993, confirmed 

earlier results on the ultraviolet damage effects in covered solar cells of various types 

being used, or proposed for use, in INTELSA T programs. Two different UV test systems 

were used to identify systematic errors and to study the effects of UV source-bulb age on 

degradation rate. 

After correction for contamination and UV source-bulb aging, the extrapolated 

degradation rates for irradiated and unirradiated INTELSAT-5, -6 single AR (SAR) 

coated cells and INTELSAT-7, -7 A, -8 double layer AR (DAR) coated cells2 in both the 

1993 tests confirm the following hypotheses resulting from the 1992 experiment. 

a. Irradiated cells display significantly more UV degradation than do the 

unirradiated cells for tests exceeding 2000 hours. [The new data indicates that 

degradation effects from electron irradiation are proportional to t2 (the square of 

the UV hours), at least for times 5 3000 hours.] 

b. This difference does not depend upon antireflective coating, cell resistivity, 

or manufacturer within the sensitivity and reproducibility of the experiment. 

c. There is a clear difference in degradation rate between single AR coated 

cells (TiOx) and double layer AR coated cells (SiOx and AI203?). At 100,000 hours 

(11.4 years) the DAR coated cells display more degradation than do the SAR 

coated cells, even though at 1,000 hours the DAR cells display less degradation. 

d. UV degradation rates, to modern covered silicon solar cells, at the beginning of 

bulb life drop from -2 times the average rate to near zero after 2000 hours 

(average end-of-life for the xenon short-arc lamps used in the tests). 

The effects of 1 MeV electron irradiation (1015 e-/cm2) prior to UV exposure are 

clearly indicated in the plot of percent change in cell open circuit voltage (Voc) versus 

percent change in short circuit current (lsc) during the UV test and post-test cleanup of the 

cells (Figure 1). The heavy lines indicate the trends of the data for both unirradiated and 

pre-irradiated cells (different cell types and resistivity show the same trend). The slopes of 

1 This work was supported by the International Telecommunications Satellite Organization (INTELSAT) 

and by COMSAT Corporation. 

2 The tested INTElSAT-5, -6, and ·7 cells are German and the INTELSAT-7A and -8 cells are Japanese. 

223 

--~-- -------------------' 
the trend lines differ between the unirradiated and irradiated cells during UV exposure but 

not during the post-test removal of contamination. Clearly, extended UV testing produces 

a permanent photo-induced redegradation of previously irradiated cells. However, this 

photo-induced redegradation may be caused by the long-wavelength light, not the UV 

light. 

0 .0 

-1.0 

-2.0 

C.) 

en -3.0 

c: 

-4.0 

CIl +---t-.,....,t-----;- Irradiated Cells 

C) 

c: 

IV -5.0 ~ 

0 

~ 0 -6.0 

L __ _ 

-7.0 

-8.0 

-9.0 

-1.6 

Figure 1. 

Unirradiated Cells 

-1.4 -1.2 -1 .0 -0.8 -0.6 -0.4 -0.2 0.0 

% Change In Open Circuit Voltage 

The % change in cell open circuit voltage versus % change in Isc durin g the 

UV test and post-test cleanup of unirradiated and irradiated (labeled), 2-

and 10- ohm-em (clear and filled symbols respectively) solar cells. 

Figure 2 is representative of the reduced data obtained in the two 1993 tests 

reported here. The corrections to the data include: normalization against control cells, to 

adjust for any long-term intensity or spectral drift of the solar simulator; modification of 

the time base, to adjust for changes in the damaging portion of the UV test source 

spectrum; and, subtraction of the contamination that accumulated on both the quartz 

window and coverslides during the extended test. 

A comparison of the results for SAR and DAR coated cells from the one 1992 and 

the average of two 1993 Tests (in the table) indicates their level of agreement and the 

spread in data and extrapolations of the tests. The very high degradation seen in the 

extrapolated result for irradiated DAR coated cells is partially an artifact of the 

assumption that photo redegradation of preirradiated cells does not saturate. 

224 

-  J 
I 

l 

I 

I 

0.5 

0 

-0.5 

-1 

-1.5 

0 

(/) -2 

c: -2.5 

CI 

en -3 c 

C\J 

.r:. 

(.) -3.5 

0 ~ -4 

-4.5 

-5 

-5.5 -

-6 

---Zero % Degradation 

- - .. - - 1-5 Cells 10 OHM-CM 

--- -A--- 1-7 Cells 10 OHM-CM 

1-7 Cells 2 OHM-CM 

- - - -0 - - - 1-7A Cells 2 OHM-CM 

- - - Single AR Coated Cells 

\ ~ 

\ 

, 

\ 

, 

- - - - Double AR Coated Cells t-----t------t------i 

, 

10 100 1000 10000 100000 

UV Solar Hours 

Figure 2. Corrected UV degradation, data and extrapolations, for the unirradiated 

cells of Test 1 in the 1993 experiment. 

Comparison of 1992 and 1993 Test Results for SAR and DAR Coated Cells 

Cell Configuration % Msc at % Msc at % Msc at 

1000 hours 10,000 hours 100,000 hours 

UNIRRADIATED 1992 /1993 1992/1993 1992/1993 

SAR -1.7/ -1.6 ± 0.5 -3.8 / -2.5 ± 1 -5.9 / -3.3 ± 1.5 

DAR -0.5 / -1.5 ± 0.5 -2.5/ -3.5 ± 1 -6/ -6 ± 2 

IRRADIATED 

SAR NA / -2.2 ± 0.5 NA / -4 ± 1 NA / -5.7 ± 1.5 

DAR -0.8 / -1.5± 0.5 -5 / -5.2 ± 1 -14/-10±2 

Eleven years exposure to the space UV environment should degrade unirradiated 

1-5,6 cells by 4.5 ± 1.5% and 1-7,8 cells by 6 ± 2%. At 4000 hours, 10 and 2 ohm-em, 

preirradiated cells display -1.8 and -2.4% greater loss than do the corresponding 

unirradiated cells. Therefore, degradation of the preirradiated cells at 100,000 hours 

should be ~ 6.5 ± 2% for 1-5,-6 cells and ~ 8 ± 2% for 1-7,-8 cells. It is not yet determined 

which set of laboratory data corresponds to space experience in a radiation environment. 

225 
37'177'3 

N95-20527 

PARASITIC CURRENT COLLECTION BY SOLAR ARRAYS IN LE01 

Victoria A. Davis and Barbara M. Gardner 

S-Cubed Division of Maxwell Laboratories 

San Diego, California 

SUMMARY 

Solar cells at potentials positive with respect to a surrounding plasma collect electrons. Current is 

collected by the exposed high voltage surfaces: the interconnects and the sides of the solar cells. This 

current is a drain on the array power that can be significant for high-voltage arrays. In addition, this 

current influences the current balance that determines the floating potential of the spacecraft. One of the 

objectives of the Air Force (PUGPS) PASP Plus experiment is an improved understanding of parasitic 

current collection. As part of the PASP Plus program, we are using computer modeling to improve our 

understanding of the physical processes that control parasitic current collection. 

BACKGROUND 

Solar arrays provide power for nearly all space systems. Traditionally, solar arrays have operated in 

the 30 V range to avoid complex interactions with the plasma environment. As space systems become 

more complex, more power, therefore higher voltages, is needed. 

Typically, the negative side of each solar array is grounded to the spacecraft chassis. This makes the 

exposed metal and semiconductor of the arrays positive with respect to the spacecraft body. The 

equilibrium potential of a spacecraft with respect to the plasma is the potential at which there is no net 

current to the spacecraft. As illustrated in figure 1, there are several components of the current to a 

spacecraft. Portions of the solar array attract ions and other portions attract electrons, depending on the 

local potential. Ions are attracted to exposed conductors on the spacecraft surface, as the spacecraft 

body is negative with respect to the plasma. Particle beams and other emitters also contribute to the net 

current. Since electrons are more mobile than ions, spacecraft float negative unless the potential is 

actively controlled. 

It may be necessary to keep the spacecraft body near zero potential with respect to the plasma. For 

example, an instrument to measure the low energy plasma environment may need to be near plasma 

ground. Anodization arcing and negative potential arcing are potentially disruptive at potentials greater 

than 50 to 100 V negative with respect to the plasma (refs. 1 and 2). Particle beams and plasma 

contactors are sometimes used to maintain the spacecraft potential. If the solar arrays are at 150 V and 

the spacecraft chassis is maintained at no more than 50 V negative with respect to the plasma, the most 

positive portions of the solar arrays are at 100 V, with respect to the plasma. Therefore, it is important to 

consider current collection by the portions of solar arrays at positive potentials. 

A traditional solar array is an array of solar cells 2 to 8 cm in size and connected in series by 

millimeter-sized metal interconnects. The interconnects can be metallic meshes or sets of wires. Each 

solar cell is protected from the environment by a cover glass. Figure 2 shows the regions between cells in 

a typical array. Current is collected by the portions of the metallic interconnects and the semi-conducting 

solar cells exposed to the plasma environment. 

1 This work is supported by the Air Force Materiel 'Command. 

227 

PAGi eLAN NOT FtLMED 
At low potentials, a barrier forms that prevents electrons from reaching the high voltage surfaces. The 

barrier is created by the surrounding insulating surfaces, primarily the coverglass, that float slightly 

negative. The current collected is the high energy tail of the distribution and therefore depends 

exponentially on the barrier height. The barrier height depends on the potential, the plasma, and the 

geometry. 

At high potentials, typically over 200 V, the current rapidly rises due to a phenomena called snapover. 

Snapover was first observed at NASNLeRC (refs. 3, 4, and 5). Snapover can occur whenever there is a 

biased surface adjoining an insulating surface and the bias is above the first crossover of the secondary 

yield curve of the insulating surface (refs. 6 and 7). 

When the solar cell potential is below the first crossover of the secondary yield curve, the insulating 

surface floats slightly negative (a few times the electron energy) and collects as many ions as electrons 

from the ambient plasma. If the potential of the insulating surface is above the first crossover, each 

electron generates more than one electron. The secondary electrons may either return to the surface or 

move across the surface under the influence of surface electric fields. The equilibrium potential at each 

location is the potential at which the net current of electrons from the plasma and secondary electrons 

arriving at and leaving the location balance. The exposed conductor is a sink for the secondary electrons. 

The surfaces adjust so that a potential gradient exists across the insulating surface attracting the 

secondaries to the conductor at the highest potential. The net effect is that the high potential area and the 

collecting area increase. There is a range of applied bias values for which the insulating surface may 

either float slightly negative or be snapped over. Experimentally, hysteresis is observed. The size of the 

snapped over region, and therefore the current, depends on the local geometry as well as the potentials 

and the plasma. 

The Photovoltaic Array Space Power Plus Diagnostics (PASP Plus) flight experiment will explore high 

voltage current collection by solar arrays (ref. 8). PASP Plus is the principal experiment integrated onto 

the Advanced Photovoltaic and Electronics Experiments (APEX) satellite bus. APEX will be launched this 

summer into a 700 elliptical orbit from 360 to 1950 km. The spacecraft attitude is such that the solar array 

test panels will always face the sun. The experiment will test twelve different solar array designs. The 

experiment will investigate negative potential arcing, parasitic current collection, and long term radiation 

damage. Parasitic current collection will be measured for eight of the designs under various operational 

and environment conditions. The arrays will be biased from 75 to 400 V. Previous space experiments that 

examined parasitic current collection include PIX I (ref. 9), PIX II (refs. 10 and 11) and SAMPlE (ref. 12). 

OUR APPROACH 

The computation of the current collected by a specific solar array can become intractable. The gap 

size is of the order of tens of mils while the solar cells are a few centimeters and the entire array can be 

meters. Each solar cell is at a slightly different potential. The current depends on the geometry of the gap, 

the geometry of the entire array, the spacecraft, and the plasma conditions. 

We are interested in improving our understanding of which aspects of the problem are most important 

and developing a tool or at least an algorithm to assist spacecraft designers. Our approach is to look in 

detail at current collection at a single cell gap. Using the computer we can vary each parameter 

independently. We then develop formulas that estimate the current collected by a single gap. We then 

incorporate the formulas into a tool that can add up the current from all the gaps to give the current 

collected by an array. Information on the array geometry and how it influences the current can also be 

added to the tool. 

To compute the current to a single gap, we are using the Gilbert computer code. Gilbert is a generalpurpose, two-dimensional, plasma analysis code. It can be used to solve for the electrostatic potential 

about an object, with flexible boundary conditions on the object and with space charge computed either 

228 
fully by particles, fully analytically, or in a hybrid manner. For this study Gilbert is used to first compute the 

electrostatic potentials in space around the solar cells and then to compute the electron trajectories in the 

previously computed potentials. The space charge was computed using an analytic formulation. The 

analytic formula includes charge density variation due to acceleration and convergence (ref. 13). 

Figure 5 shows a typical grid for the Gilbert calculations. The gap and the space above the gap are 

gridded. The surfaces provide the boundary conditions. This grid represents solar cells with mesh 

interconnects. An idealized geometry was chosen because we are interested in understanding the 

relationship between the variables. The interconnect is modeled as a diagonal line extending across the 

bottom of the gap. The side of one solar cell is at a high potential along the left side of the gap. The sides 

of the coverglass on both sides of the gap are included. All the insulating surfaces have a special 

boundary condition that represents the snapover condition when appropriate. 

Figure 6 shows potential contours and a few trajectories for a typical case. The potentials are 

computed using an analytic representation of the charge density. Macroparticles are emitted from several 

locations along each sheath segment. At each location, a set of particles with a thermal distribution of 

velocities is created. These macroparticles are tracked. The collected current is the current carried by the 

macroparticles reaching the high potential surfaces. 

Current collection can be either orbit limited or space charge limited, depending on the debye length of 

the plasma and the size of the collecting area. 

There are two serious limitations to computing the current in this way. Implicit in the technique of drawing 

a sheath edge and tracking current from this sheath edge is the assumption of a sharp sheath edge. This 

assumption is not necessarily valid when the sheath size is of the order of a debye length. The computed 

sheath sizes are smaller than a debye length for some geometries under some plasma conditions, 

particularly for geometries without an interconnect and at low potentials. 

The other limitation is the assumption that current can reach the sheath. At lower cell potentials, the 

coverglass surface potential dominates the long distance potential. To reach the "sheath edge," an 

electron must pass through a negative potential region. Only the electrons in the high energy tail of the 

distribution are collected. A different technique is needed to compute the current in this regime. 

A different approach is also needed at the highest potentials. When the entire surface of the array is 

snapped over, the sheath is dominated by three-dimensional effects. 

RESULTS 

We have done calculations for three geometries: cell-to-cell gap region with an interconnect, without 

an interconnect, and a single cell edge. The calculations span the space of anticipated plasma conditions, 

applied potential, and first crossover potential. The current rises rapidly when the applied bias is four to 

five times the first crossover potential. Figure 7 shows how the current varies as a function of the applied 

bias for a single solar cell edge. 

We now have analytic formulas for the dependence of the current on the primary problem variables: 

applied voltage, plasma conditions, and, to some extent geometry. These formulas have been 

incorporated into the EPSAT computer code. 

EPSAT is an analysis tool for determining the performance of power systems in both naturally 

occurring and self-induced environments (ref. 14). EPSAT is an engineering spreadsheet that allows rapid 

"what if' analysis of the effect of parametric changes on a space-based system's performance. EPSAT 

provides information on the plasma and neutral environment anticipated on orbit. The code presently 

models many of the interactions of the space environment with a power system, including sheath 

229 
formation, v-x B potentials, particle beams, and sheath ionization. The code permits the user to do 

mission studies and evaluate the importance of the various interactions with a specific spacecraft design. 

The user can then adjust the design and examine the effect of the change on the ability of the system to 

operate in natural or hostile environments. EPSAT uses a unique architecture for integrating analysis 

models that allows modeling capabilities to evolve with changing needs. 

A new solar array module for EPSAT has been developed. The current to a solar array is computed 

by summing the current to each solar cell edge, interconnect, and gap without interconnect, while not 

double counting. The current to a single edge or gap is computed using a fit to the two-dimensional 

computations. 

Figure 8 shows the current to PASP Plus array #1, mesh interconnect design, under typical space 

conditions for different values of the first crossover potential. Figure 9 shows the current to PASP Plus 

array #3, space station design, under typical space conditions for different values of the first crossover 

potential. 

Figure 10 shows how the current computed in this manner compares with laboratory measurements made 

on the PASP Plus array #8 . Array 8 has a wrap-thru interconnect design. 

CONCLUSIONS 

We are refining our two-dimensional calculations and examining their limitations. As needed, we will use 

other techniques to extend the range of our calculations. The results are continuing to be incorporated 

into EPSAT. We look forward to comparing the flight measurements with the calculations. Once our 

understanding is validated, the formulas developed can be used to improve the design of solar cells to 

minimize complications due to this interaction. 

REFERENCES 

1. Vaughn, JA, et al. : Extrapolation of Electrical Breakdown Currents from the Laboratory to Space 

Station. AIAA 92-0822, AIAA 30th Aerospace Sciences Meeting, Reno, Nevada, 1992. 

2. Ferguson, D.C.: The Voltage Threshold for Arcing for Solar Cells in LEO - Flight and Ground Test 

Results. NASA TM 87259, 1986. 

3. Purvis, C.K.; Stevens, N.J.; and Berkopec, F.D.: Interaction of Large High Power Systems with 

Operational Orbit Charged-Particle Environments. NASA TM X-73867, 1977. 

4. Stevens, J.N.: Interactions Between Spacecraft and the Charged-Particle Environment. Spacecraft 

Charging Technology - 1978, NASA Conference Publication 2071, AFGL-TR-79-0082, 1978, pp. 

268-294. 

5. Kennerud, K.L.: High Voltage Solar Array Experiments, NASA Rep. CR-121280, 1974. 

6. Chaky, R.C.; Nonnast, J.H.; and Enoch J.: Numerical Simulation of Sheath Structure and CurrentVoltage Characteristics of a Conductor-dielectric Disk in a Plasma. J. Appl. Phys., vol. 52,1981, pp. 

7092-7098. 

7. Katz, I.; and Mandell, M.J.: Differential Charging of High-Voltage Spacecraft: The Equilibrium 

Potential of Insulated Surfaces. J. Geophys. Res., vol. 87, 1982, pp. 4533-4541. 

230 
r-----

8. Guidice, D.A.: PASP Plus: An Experiment to Measure Space-Environment Effects on Photovoltaic 

Power Subsystems. Fifth Annual Workshop on Space Operations Applications and Research 

(SOAR '91), NASA Conference Publication 3127, vol. II, 1991 , pp. 662-668. 

9. Grier, N.T.; and Stevens, N.J.: Plasma Interaction Experiment (PIX) Flight Results. Spacecraft 

Charging Technology - 1978, NASA Conference Publication 2071 , AFGL-TR-79-0082, 1979, 

pp. 295-314. 

10. Grier, N.T. : Plasma Interaction Experiment II (PIX II): Laboratory and Flight Results. Spacecraft 

Environmental Interactions Technology, 1983 , NASA Conference Publication 2359, AFGL-TR-85-

0018, 1985, pp. 333-347. 

11 . Ferguson, D.C.: RAMIWAKE Effects on Plasma Current Collection of the PIX II Langmuir Probe. 

Spacecraft Environmental Interactions Technology, 1983, NASA Conference Publication 2359, 

AFGL-TR-85-0018, 1985, pp. 349-357. 

12. Hillard, G. S.; and Ferguson D.C.: Solar Array Module Plasma Interactions Experiment (SAMPlE): 

Science and Technology Objectives. J. of Spacecraft and Rockets, vol. 30, 1993, pp. 488-494. 

13. Mandell, M.J.; and Davis, VA: User's Guide to NASCAP/LEO. LEW-15641, SSS-R-85-7300-R2, 

1990. (Available as LEW-15641 from COSMIC, U of Georgia.) 

14. Kuharski, RA, et al. High Voltage Interactions of a Sounding Rocket with the Ambient and SystemGenerated Environments. IEEE Transactions on Nuclear Science, vol. 37, 1990, pp.2128-2133. 

Electrons attracted to positive 

portion of solar arrays 

Ions attracted to negatiVE 

portion of solar arrays 

Ions attracted to 

spacecraft body 

Figure 1.-At the spacecraft floating potential the net current is zero. 

This current has several components. 

Conducting 

Interconnect 

Support 

Figure 2.-Cross section of the region between traditional solar cells. 

231 
O.2V Barrier to -..... p(KA-A..'A 

Incident A \A 

Electrons ,i "A 

A 'A 

A ~ A A 

A A 

f(. 

A 

A 

t 

i 

Marked Levels 

A -0.20 

B -0.01 

C 0.10 

o 2.00 

E -0.10 

Figure 3.-At low potentials, barrier formation dominates current collection. 

Plasma of 1012 m-3 and 0.1 eV. Cell potential of 25 V, surface potential of -0.1 V. 

First crossover above 25 V. 

No Snapover: 

dielectric potential - 0 

dielectric 

conductor 

+SDV 

Dielectric surface collects 

as many ions as electrons 

from the ambient plasma. 

+3DDV 

Snapover: 

dielectric potential> E 

Dielectric surface emits as many 

secondary electrons as it collects 

primaries and secondaries. 

Figure 4.-When a conductor is at a potential above the first crossover of the secondary yield curve 

of an adjoining insulator, the insulator surface can develop positive potentials in a phenomena 

known as snapover. 

232 
I 

I 

 

 

f 

r 

I 

I 

I 

, 

I 

I 

Figure 5.-Grid for two-dimensional Gilbert calculations. 

101~ __________ ...., 

0.0 

0.0 

Marked Lovels 

A 0.00 

F 0.07 

Marked Levels 

A 0.00 B 40.00 C 80.00 

o 120.00 E 160.00 F 0.07 

-4.E·03 ·2.E·03 0 2.E-03 4.E-03 

Figure 5.-Potential contours and a few representative trajectories. Plasma of 1011 m-3 and 0.1 eV. 

Cell potential of 150 V, surface potential of -0.5 V, first crossover of 40 V. 

233 
100 

~ 80 

"0 o Q) 

60 13 

Q) 

(5 

() 40 -

c 

Q) 

..... 

..... 

::J 20 () 

0 

0 100 200 300 400 

Applied Bias (V) 

Figure 7.-Current to 1 m along 5 mil thick solar cell edges as a function of applied bias. 

Plasma of 10 10 m-3 and 0.1 eV. First crossover of 40 V and surface potential of -0.5 V. 

The points are the results of Gilbert calculations and the line is a fit. 

100 

6 1 st crsovr = 20 

~ 10 ---0---1 st crsovr = 40 

E-

"0 -<>--1 st crsovr = 80 Q) 

13 

~ 

(5 

0 

c 

Q) 

t: 0.1 ::J 

() 

0.01 

10 100 1000 

Applied Bias (V) 

Figure 8.-Current to PASP Plus array #1 , mesh interconnect design, as a function of 

applied bias. Plasma of 1.9 x 1011 m-3 and 0.088 eV. Surface potential of -0.5 V. 

Three value of the first crossover of the secondary yield. 

234 

--- -------- - ---
rI . . 

I 

I 

I 

I 

I 

I 

I 

I 

I 

« 

E 

'0 

Q) 

U 

~ 

(5 

() 

-

c: 

Q) 

..... 

..... 

:::l 

() 

10 

Il 1 st crsovr = 20 

-0--1 st crsovr = 40 

---<>- 1 st crsovr = 80 

0.1 

0.01 

0.001 ¥------------t------------i 

10 100 

Applied Bias (V) 

1000 

Figure 9.-Current to PASP Plus array #3, space station design, as a function of applied bias. 

Plasma of 1.9 x 1011 m-3 and 0.088 eV. Surface potential of -0.5 V. 

Three value of the first crossover of the secondary yield. 

1.00E-2 

~ 1.00E-3 

E 

Q) 

..... 

..... 

:::l 

() 1.00E-4 

'0 

Q) 

U 

~ 1.00E-5 (5 

() 

[] Experiment 

----Calculation 

1.00E-6 +-U4'----+-----+-----t-------i 

o 100 200 

Array Bias (V) 

300 400 

Figure 10.-Comparison of model of parasitic current collection with laboratory 

measurements for PASP Plus array #8. The environment is 4 x 1011 m-3 at 0.5 eV. 

The first crossover potential is assumed to be 25 V. 

(Unpublished experimental data courtesy of N. T. Grier of NASNLeRC.) 

235 

----- - ---- - _1 
SESSION IV 

FLIGHT RESULTS 

PRtlOIDtN .. AGE N NOT FILME 
,

_ ._~ 

-- --

30;7795J' 

N95- 20528 

FINAL RESULTS OF THE ADVANCED PHOTOVOLTAIC EXPERIMENT 

David J. Brinker 

NASA Lewis Research Center 

Cleveland, Ohio 

and 

John R. Hickey 

The Eppley Laboratory, Inc. 

Newport, Rhode Island 

SUMMARY 

The Advanced Photovoltaic Experiment was designed to generate laboratory reference 

standards as well as to explore the durability of a wide variety of space solar cells. In 

addition to the cells, it was equipped with an absolute cavity radiometer to measure solar 

intensity, a spectroradiometer to measure the spectral content of this radiation and a sun 

angle sensor. Data from the solar cells and various sensors was obtained on a daily basis 

during the first eleven months of the 69 month flight. In this paper we compare pre-flight 

and post-flight laboratory measurements with on-orbit calibration data. Pre-flight and 

post-flight calibration data of the cavity radiometer as well as on-orbit data demonstrated 

the accuracy and durability of the Eppley Labs. instrument flown on APEX. 

INTRODUCTION AND EXPERIMENT DESCRIPTION 

The Advanced Photovoltaic Experiment (APEX) is a space flight test designed to 

provide reference cell standards for laboratory photovoltaic performance measurements as 

well as to investigate the solar spectrum and the effect of long term exposure of solar cells 

to the space environment. Toward this end, 155 solar cells of the widest available variety of 

design and material were incorporated into the experiment along with sensors to measure 

total solar irradiance and sun angle. Experimental measurements were made on a daily 

basis, dependent upon the achievement of proper sun angles, and recorded on an on-board 

magnetic tape recorder. The experiment was designed around the original flight time of one 

year, with battery capacity the principal lifetime limiting factor. Useful data was in fact 

obtained for 325 days, at which time the voltage of the batteries supplying the data 

acquisition system fell below the threshold necessary for calibrated operation. Details of 

the design and operation of APEX have been previously published (ref. 1,2). 

When the announcement of opportunity for LDEF experiments was released in 1976, a 

launch of about 1980 was envisioned. As a result, the solar cell samples prepared for APEX 

represented the state-of-the-art in space cell technology as of 1979, as well as samples of 

cells in use on a variety of satellites. A Shuttle-caused delay in the launch of LDEF by 

several years provided both the opportunity and necessity for updating the sample set, to one 

including the most recent advances in technology. The cell investigators were invited in 

mid-1982 to submit new cells. Of the 136 calibration cells (120 Isc and 16 IV cells), 69 

were replaced. Many of those which were not replaced were either standards previously 

calibrated by other techniques or representative of cells in use on a variety of satellites. 

239 

flftS<*l1N8 AGE LANK NOT FILMED 
These APEX solar cell investigators and the number of cells each supplied are: 

A.F. Wright Aeronautical Laboratory 

Applied Solar Energy Corporation 

COMSA T Laboratories 

European Space Agency 

Jet Propulsion Laboratory 

NASA Lewis Research Center 

NASA Marshall Space Flight Center 

Solarex Corporation 

Spectrolab, Inc. 

8 

1 4 

7 

9 

34 

56 (Includes 19 sensor cells) 

1 1 

7 

9 

Each group provided cells representative of technologies which were either in 

development or production. The experiment was designed to accommodate a total of 155 such 

cells , including the silicon cells which were employed as sensors for the spectral 

radiometer portion of the experiment. All cells were permanently mounted on aluminum 

plates with a thermistor in contact with the rear of the cell. 139 cells were designated as 

Isc cells, 120 to be calibrated as reference standards and returned to the investigators, 

eighteen for use as spectral radiometer sensors and one as a night sensor to signal the data 

acquisition system that conditions were correct for the requisite periodic calibration of the 

cavity radiometer. For these cells , the short-circuit current was converted to a voltage 

through the use of a precision load resistor. In most cases a 0.1 Q value was used. The 

remaining sixteen cells were designated IV cells , that is the entire current - voltage 

characteristic was measured through the loading of the cell by a series of five appropriately 

sized resistors. 

At that time, Si cells were the only type in production, with the development of gallium 

arsenide in its early stages. This is reflected in the distribution of these semiconductor 

types in the APEX complement, which is summarized by cell type and size below: 

Silicon: 105 

21 

2 

15 

-.1 

144 

2 x 2 em 

2 x 4 em 

5 x 5 em 

5.9 x 5.9 cm 

6 x 6 cm (module) 

Gallium Arsenide: 10 2 x 2 cm 

---1 1.3 x 1.6 cm 

1 1 

The cells were mounted on 127 aluminum plates of twelve different sizes and 

configurations. 28 of the mounts each held two 2 x 2 cm cells. Each mount was equipped 

with a Yellow Springs Instruments Type 16429 thermistor (10,000 Q @ 25 °C). An 

additional thermistor monitored the Eppley absolute cavity radiometer. 

SENSOR RESULTS 

A detailed examination and recalibration of the various sensors and instruments on 

APEX was performed. The Eppley Type HF absolute cavity radiometer ,«as of particular 

interest in that it is the only radiometer ever returned from an extended stay in space. It is 

identical to an instrument that has provided over 12 years of data as part of the Earth 

Radiation Budget experiment on Nimbus 7. After detailed examination at the Eppley 

240 

L ____ _ 
Laboratories, it was intercompared with standards at the World Radiation Center in Davos, 

Switzerland (ref. 3). As seen in Table I, the 69 months in-orbit had little effect on the 

radiometer. Its sensitivity was essentially unchanged, as was the its reflectance , in 

agreement with the results of visual inspections . 

. A spectroradiometer was included in APEX to measure the spectral content of the 

extraterrestrial sunlight. After return, it was intended that the instrument be used for 

calibration of laboratory solar simulators. The spectroradiometer consisted of sixteen 

identical silicon cells with narrow bandpass optical filters. The wavelength center of the 

filters ranged from 325 nm to 1100, with the last filter covering the infrared out to 2000 

nm. Figure 1 compares values of solar irradiance from the APEX flight data with the World 

Meteorological Organization (WMO) standard. A comparison of pre-flight and post-flight 

transmittance of the first thirteen filters is seen in Figure 2 The drastic change in 

transmittance in many filter is clearly seen, although there was no major shift in 

wavelength band. This degradation was, of course, unexpected and has made the instrument 

unsuitable for further use in the laboratory. An explanation of the effect or its probable 

cause has not been forthcoming from the filter manufacturer (ref. 4). 

The Digital Sun Angle Sensor (DSAS) was removed and returned to the manufacturer, 

Adcole Corporation, for recalibration. Its performance was unchanged due to the flight. One 

of the two graticules was cratered from a collision with a debris/micrometeoroid particle. 

Shadowing from the crater was evidence at the extreme angle limit of the DSAS (32°), but 

did not otherwise effect its performance. The instrument operated flawlessly during the data 

taking portion of the flight. 

SOLAR CELL RESULTS 

The overall condition of the cell sample set was excellent. A contaminating film seen 

over much of LDEF was present to a varying degree on APEX, the thickness of the layer 

dependent upon location. No loss of cell coverglass nor significant changes in color or 

appearance was observed. Several of the cells were cratered from micrometeoroid and/or 

debris impacts, with the range of damage spanning from microscopic craters in the 

coverglass surface to penetration of the coverglass and cell and crate ring of the underlying 

aluminum mounting plate. However, even the few cells in which the crate ring extended into 

the solar cell itself, or caused a crack in the coverglass and cell, electrical continuity was 

maintained. Loss in current proportional to the damage area and increase in fill factor due 

to cell cracking was observed. The electrical leads from the mounting plate feedthrough to 

the cell front and rear contacts were found to open in six cells. A silver ribbon of about 3 

mil thickness was used for these cells. Where the flat portion of the ribbon faced the ram 

direction, the ribbon was severely eroded, creating an open circuit. In most cases the 

ribbon twisted through 90° at the feedthrough so that the narrow (3 mil) edge faced the ram 

direction; here the silver ribbon remained intact. Examination of the flight data indicates 

that the erosion did not occur to any extent that would affect cell performance during the data 

recording portion of the flight, the first eleven months. Post-flight performance testing of 

these cells was accomplished by direct probing of the cell contacts, no significant change 

from pre-flight performance was seen. 

The first post-flight electrical test performed was measurement of the short-circuit 

241 

I 

I 

___ I 
current utilizing the precIsion load resistor mounted on each cell for the flight. The 

resistors were soldered to the cell mounting plate electrical feedthroughs on the underside of 

the cell mounting plates. These measurements, as well as subsequent current-voltage (I-V) 

tests, were carried out in the Solar Cell Evaluation Laboratory at Lewis Research Center 

using a Spectrolab X-25L solar simulator. This simulator employs a short-arc xenon lamp 

as the light source and provides uniform, collimated illumination. The intensity of the 

simulator was set using an aircraft calibrated silicon standard which is identical to the 

standard used at Eppley Laboratory for pre-flight testing, where a xenon arc lamp 

simulator was also utilized. Cell temperature was monitored using the flight thermistors. 

One thermistor was found to be open. An examination of the flight data showed abnormal 

readings from it, indicating that the failure occurred before launch. With this sole 

exception, all of the thermistors functioned properly, providing values in close agreement 

with a temperature sensor used in controlling the laboratory test fixture. The shortcircuit current values obtained in these test are useful in comparison with both pre-flight 

performance and flight data. The values obtained were in most cases in good agreement with 

pre-flight values, with the exception of those cells without a coverglass. 

Upon completion of the measurement of short-circuit current, the load resistor was 

removed from the circuit by cutting one of its two leads. If LDEF had been retrieved on 

schedule and the value of the cells as calibration standards was retained, the load resistors 

could not have been removed. However, the absence of data from the last five years on-orbit 

negates their usefulness as standards. The complete I-V characteristic of all cells were then 

measured at 25 °C and recorded. Table II compares flight data with pre- and post-flight 

simulator data for a small representative sampling of the silicon cells. All data presented 

here is corrected to 25 °C and one AMO sun (136.7 mW/cm2). All of these cells are n-p 

type, as were most of the silicon cells flown. Also included is the value of short-circuit 

current as measured on-orbit. 

The loss in voltage and current in the unglassed cells (cells ISC 63 and ISC 83) was 

consistent with the proton radiation flux of the 250 nm, 280 inclination orbit. The 

variation in pre-flight to post-flight Isc for the last four cells of Table II is typical of the 

entire cells set; the variation was within about ±2% It is believed that this variation was 

due to inherent differences is the LeRC simulator and data acquisition system used for postflight testing and those at Eppley Labs used for pre-flight testing. Pre-flight testing was 

not performed at LeRC as it was expected that the APEX cells would be used as calibrated 

reference cells; the evolution of the experiment into a long term durability test was neither 

expected nor ever contemplated. The I-V characteristic of 16 cells was measured by first 

recording the open-circuit voltage and then switching in five load resistors and recording 

cell current and voltage for each. The excellent agreement of the on-orbit I-V data (black 

squares) with the post-flight laboratory data of a Solar Maximum Mission cell in shown 

(Figure 3). 

A summary of results from GaAs cells is shown in Table III. Of the eleven GaAs cells 

flown, ten were made by Hughes Research Laboratory using liquid phase epitaxy, the only 

type of GaAs cells then available. These cells had junction depths of either 0.35 11m or 0.50 

11m and were flown with either a 12 mil fused silica coverglass or unglassed. The 

performance degradation, particularly in open-circuit voltage, of the uncovered cells is 

consistent with the known radiation characteristics of these cells and the mostly proton flux 

of the orbit. The eleventh GaAs cell was a metal-oxide-semiconductor design which was 

242 
1- - ------ ----

under study at that time as a low fabrication cost alternative to epitaxial growth for III-V 

compounds. It has a very thick coverglass of unknown thickness and material and degraded 

significantly in short-circuit current. The accuracy of the preflight Isc value is suspect. 

Detailed discussion of other cells in the set, particularly those which sustained 

physical damage from micrometeoroids/debris, can be found in references 2 and 5. 

CONCLUSION 

The solar cells and sensors flown on APEX survived well their nearly six years in 

low earth orbit. Post-flight testing of the Eppley cavity radiometer shows that it is 

essential unchanged from its pre-flight condition. The solar cell set survived equally well, 

with no evidence of electrical or physical changes except in those cases where debris or 

micrometeoroids cratered the cells. Pre-flight and post-flight performance measurements 

are in very good agreement. 

REFERENCES 

1. Brinker, D.J.; Hart, R.E. and Hickey, J.R.: Preliminary Results from the Advanced 

Photovoltaic Experiment Flight Test, 21 st IEEE Photovoltaic Specialists Conference, 

1990, pp 1213-1218. 

2. Brinker, D.J.; Hickey, J.R. and Scheiman, D.A., Advanced Photovoltaic Experiment, 

S0014: Preliminary Flight Results and Post-Flight Findings, 1 st LDEF Post-Retrieval 

Symposium, June 1991, NASA CP 3134, pp 1395-1404. 

3. Hickey, J.R., Passive Exposure of Earth Radiation Budget Experiment Components, 

LDEF Experiment AO-147: Pest-Flight Examinations and Tests, 1st LDEF PostRetrieval Symposium, June 1991, NASA CP 3134, pp 1493-1510. 

4. Hickey, J.R.; Brinker, D.J. and Jenkins, P.P., Studies of Effects on Optical Components 

and Sensors: LDEF Experiments AO-147 (ERB Components) and S-0014 (APEX), 2nd 

LDEF Post-Retrieval Symposium, June 1992, NASA CP-3194, pp. 1375-1388. 

5. Brinker, D.J.; Hickey, J.R. and Scheiman, D.A., The Effect of the Low Earth Orbit 

Environment on Space Solar Cells: Results of the Advanced Photovoltaic Experiment 

(S0014), 2nd LDEF Post-Retrieval Symposium, June 1992, NASA CP-3194, pp. 

1291-1302. 

243 

________________ __ _________________ _____ J 
Table I - Cavity Radiometer Tests and Comparisons 

Pre-Flight Post-Flight 

Heater Resistance (0) 152.2 152.2 

Thermopile Resistance (0) 354.4 354.7 

Power Sensitivity Ratio 1.2146 1.21"26 

(Vac/Atm) 

Intercomparison With World Radiation Reference 

Cell Number 

ISC 63, NA-10 

ISC 83, B-21R 

ISC 86, B2S0F 

ISC 95, M-5 

ISC 112, B-2R 

ISC 114, B-4R 

APEX Instrument 1 .00069 

EPLAB Reference Instrument 

Reflectance Measurement (WRC) 

APEX Instrument 

New cavity 

1.00002 

250 ±80 ppm 

270 ±80 ppm 

Table 2 - Silicon Cell Data 

Description Coverglass Isc (mA) 

Pre-Flight Flight 

Solarex, BSF/R No Cover 146.9 144 .6 

LeRC AlC AStand. No cover 150.1 152 

COMSAT, Intsat 5 6 mil F.S. 288.2 300 .9 

ASEC, 6x6 W/A 6 mil F.S. 1199 881 .3 

COMSAT Blue 30 mil V-Groove 160 178.6 

COMSAT Non-Refl. 12 mil F.S. 193 .1 202 .7 

% Change 

0.00 

0.08 

-0.16 

Post-Flight 

133.5 

(t.voc=-65 mV) 

145.4 

(.Woc=-46 mV) 

292 .3 

1195 

163.7 

189.3 

244 l I 

-~---------------------- --------~ 
Table III • Gallium Arsenide Cell Data 

Cell Number Description Coverglass 

Pre-Flight 

ISC 71 , NB-15L HRL, Dj=0.5 ~m 12 mil F.S. 122 .5 

ISC 76, NB-29R HRL, Dj=0.5 ~m No cover 117 .2 

ISC 77, NB-29L HRL, Dj=0 .35 ~m No Cover 117.3 

ISC 111, A-2 JPL, AMOS Unknown 17.6 

E 

:1. 

N 

E 

~ 

a> 

u 

c 

CIl 

"0 

~ ... 

2000 

1500 

1000 

500 

, , 

.. • 

, , , , 

, , 

. \ 

, 

, 

, 

, 

, 

\ 

, 

~,~\ , 

, 

-" 'II 

, 

, 

, 

, 

, 

WMO Spectrum 

----.--- APEX Filters 

\ 

, , , , i.-• 

Isc (rnA) 

Flight 

116.1 

111 .6 

, 13 .1 

33 .6 

o~~~--~~~~~~--~~~--~~~~~~ 

o 500 1000 1500 

Wavelength (nm) 

Post-Flight 

108 

(LWoc= 10 mV) 

95 .5 

(tNoc=-65 mV) 

93 .6 

(to Voc=-85 mV) 

22 

Figure 1 • APEX Filter Data Compared With WMO Spectrum 

245 
w 

o 

z 

<{ 

f--

f--

~ (/) 

z 

<{ 

a: 

f--

13 

.4 ...... .. ; . ... . .. .: .......... . 

.2 

< .§. 

f--

Z 

LU 

a: 

a: 

::::> 

u 

475 525 575 625 675 725 775 825 8 75 925 975 

WA VELENGTH (nm) 

Figure 2 - Transmittance of APEX Filters 1 to 13 

Cell : IV. 7 8 · 1L 

Relerence Cell : A- I 04 

Atea : 4 sq em. 

Temperalure : 26.5 cleg C 

Atr Mass Zero 

File Name : IV.7INl 

200 

160 

120 

ao 

.1 .2 .3 .4 .5 

VOL TAGE (volls) 

Isc _ -16052 mA 

Voc _ -579.5 mV 

1m ... _ -143.a rnA 

Vmax _ -473. 1 mV 

Pmax _ 68.05 mIN 

F.F. _ 73.1 

EN. _ 12.4 "f. 

.6 .7 

Figure 3 - I-V Curve of Solar Maximum Mission Cell 

246 

J 
r- -~- ------ -

3 Q /)'J77 

N95- 20529 

PRELIMINARY RESULTS FROM THE FLIGHT OF THE SOLAR ARRAY MODULE 

PLASMA INTERACTIONS EXPERIMENT (SAMPlE) 

Dale C. Ferguson and G. Barry Hillard 

NASA Lewis Research Center 

Cleveland, Ohio 

SAMPlE, the Solar Array Module Plasma Interactions Experiment, flew in the Space Shuttle 

Columbia payload bay as part of the OAST-2 mission on STS-62, March, 1994. SAMPlE biased 

samples of solar arrays and space power materials to varying potentials with respect to the surrounding 

space plasma, and recorded the plasma currents collected and the arcs which occurred, along with a 

set of plasma diagnostics data. A large set of high quality data was obtained on the behavior of solar 

arrays and space power materials in the space environment. This paper is the first report on the data 

SAMPlE telemetered to the ground during the mission. It will be seen that the flight data promise to 

help determine arcing thresholds, snapover potentials and floating potentials for arrays and spacecraft 

in LEO. 

INTRODUCTION 

Solar cells in low Earth orbit (LEO) environments have been shown to arc into the surrounding 

plasma when they are at a potential highly negative of the plasma (Ref. 1) and to collect anomalously 

large currents from the plasma (to "snapover") when at high positive potentials (Ref. 2). The space 

flight experiments PIX and PIX II showed that these phenomena are not confined to a laboratory 

environment, but occur naturally in space plasmas. Using data from these experiments and ground 

tests, Ferguson (Ref. 3) derived a threshold potential for plasma arcing from standard silicon solar 

cells, which has been the object of several theoretical treatments (Refs. 4 through 6). Based partly on 

these treatments, several means of mitigating the arcing phenomenon have been proposed (see Ref. 7). 

In addition, it has been shown that the currents collected by high voltage array strings in LEO will 

force negatively grounded power systems 10 "float" at high negative potentials relative to the plasma, 

where solar array arcing and/or breakdown of dielectric coatings may occur, causing possible power 

disruptions, EMI, and surface damage (Ref. 8). To prevent such occurrences, a plasma contacting 

device (a plasma contactor, or "PC") has been baselined for the International Space Station (lSSA), 

and operates by actively controlling the ISSA floating potential. This complex and expensive solution 

to a simple problem makes one desire simpler, cheaper methods of predicting and/or controlling 

spacecraft floating potentials or arcing thresholds. SAMPlE was flown to investigate the arcing and 

current collection phenomena in space, to enable the understanding and mitigation of arcing and other 

undesirable plasma effects. 

SAMPlE 

Figure 1 shows the electronics box and sample tray of the SAMPlE experiment. It was mounted on 

the top of a Hitchhiker cross-bay experiment carrier in the payload bay of the Space Shuttle Columbia. 

Inside this box (Figure 2) were the electronics that controlled the experiments, biased samples on the 

sample tray to high potentials, measured, stored and telemetered the resulting data, and measured the 

surrounding plasma conditions. On the side of another Hitchhiker attachment plate were two plasma 

diagnostic probes (a Langmuir probe and a vehicle potential, "V-Body", probe) to determine the 

plasma conditions during the biasing experiments. A pressure gauge monitored the payload bay 

pressure at one end of the electronics box. 

247 
Figure 3 shows a blowup of the electronics box and sample tray (experiment plate) assembly. On the 

sample tray were the samples which were biased in the LEO plasma conditions (Fig. 4). In all, there 

were 37 different experiment configurations which were tested in LEO. Two high voltage power 

supplies (HVPS-l and HVPS-2) in the electronics box biased the experiments to predetermined 

voltages relative to the Orbiter. An electrometer in the HVPS-l circuit measured the currents 

collected, and transient current detectors on both power supply circuits counted the arcs that occurred, 

on the biased samples. Both ion and electron collection currents were measured at the appropriate 

biases. The maximum biases attained were -600 V and + 300 V, relative to the Shuttle chassis. 

Thanks to the large collecting area of the Shuttle main engine nozzles, Columbia stayed within a few 

volts of the surrounding plasma during all of the bias and plasma conditions encountered. 

Of interest to the Space Photovoltaic Power community were several solar array experiments. A 

twelve cell series coupon of the Advanced Photovoltaic Solar Array (APSA) was flown . A four cell 

string of Space Station technology cells were also flown. As the backs of these ordinarily flexible 

arrays are covered with Kapton, precluding significant plasma interactions from the array backs in 

LEO, the array backs were attached to the experiment plate by adhesive. These arrays were loaded to 

near their maximum power points by load resistors. In addition, three separate concentric strings of 

standard 2x2 cm silicon solar cells (similar to those flown on PIX and PIX II) were flown to serve as 

a control, and to investigate the dependence of current collection on the presence of surrounding solar 

cells. 

Three sets of modified quarter-cells of Space Station design were flown to investigate the dependence 

of arc rate and current collection on solar cell parameters such as the presence or absence of adhesive 

on cell edges, the degree of adhesive coverage, and the overhang of cell coverslides past the cell 

edges. Previous work by many authors (Refs. 9 through 11) had predicted that by varying these 

parameters, cell arcing and/or current collection would be modified or prevented. To determine the 

dependence of arcing on exposed conductor properties, samples of pure metals were flown, with 

insulator strips or adjacent ground rods to encourage arcing. Since the floating potential of ISSA will 

depend to a great degree on the possibility of current collection by its radiator thermal control paint 

(Z-93), which had been shown by Hillard to be non-conductive in ground plasma tests (Ref. 12), a 

small sample of Z-93 was also flown . To investigate the arcing behavior of the ISSA structural 

material (anodized aluminum), a sample prepared to ISSA truss specifications was flown . The 

snapover effect, where insulators surrounding exposed conductors at high positive potentials start 

collecting currents as if they were conductors, was investigated by biasing conductors behind 

insulation pinholes of various sizes. Finally, a sun sensor on the sample tray confirmed the vehicle 

attitude. A more complete description of SAMPlE's experiments may be found in Ref. 13. 

Preliminary data were telemetered back through the OAST -2 payload operations control center at 

Goddard Space Flight Center. During the experiment, an anomaly with the HVPS-l circuit midway 

through the 37 hours of scheduled data-taking necessitated reconsideration of the experiment timeline, 

and new instructions were telemetered up to SAMPlE on orbit. In all, about 62 hours of data were 

obtained, stored on-board, and recovered after Columbia returned to Earth. Data were taken in the 

bay-to-earth, bay-to-ram, and bay-to-wake orientations. Initial inspection of the data show that they 

are surprisingly noise-free, and easily constitute the largest and best set of data on the interactions of 

solar arrays and space power materials with the LEO environment ever obtained. 

PRELIMINARY RESULTS 

Inspection of Figure 5 shows just how noise-free the SAMPlE data are. These curves of data on the 

electron collection of the APSA and SS (Space Station) arrays in the bay-to-earth orientation represent 

248 

l ___ _ --------- - --- --~~-- ~~---- -- ~~ ------
comparison of not just the two array types, but also of the maximum and minimum measured currents 

at each bias step. Because they were obtained at plasma densities that varied by as much as a factor 

of ten, and because the Langmuir probe data have not yet been reduced, the different array types can't 

be naively directly compared or calibrated, but it is clear that the APSA array, with a smaller array 

area, collects significantly more current than does the array of Space Station cells at all potentials less 

than about + 100 volts. Both arrays go into snapover at voltages between +200 V and +300 V, and at 

+300 V nearly the entire kapton-covered surface of the sample tray is snapped over in both cases. In 

Figure 6, the increase in APSA electron current at comparable voltages may be due to increased ram 

collection or to increased plasma densities. When the Langmuir probe data are reduced, the answer 

will be known. 

Figure 7 shows the electron current collected in the ram condition, near the time of maximum orbital 

plasma density, for the Space Station array. It is believed that the plasma densities shown in this 

figure are overestimates by about a factor of 3.6. These data show that the Space Station cells, 

despite their lack of exposed interconnectors, can collect significant currents from the plasma, and thus 

influence the Space Station floating potential. It has been estimated that to prevent large negative 

potential excursions on ISSA, an exposed conducting area of 1000 to 2000 square meters would be 

required, were the plasma contactor not baselined for ISSA. The data also support the conclusion that 

the maximum current presently being considered for ISSA PC operation (10 A) will be sufficient to 

control the ISSA potential at all times. Space Station cells that have been modified to have varying 

degrees of coverslide overhang do indeed collect different amounts of current, as shown in Figure 8. 

Preliminary looks at the data seem to imply that a coverslide overhang of 11 mils would reduce the 

electron current collection of the ISSA arrays by at least an order of magnitude, reducing the need for 

a plasma contactor. Similar effects may be obtained by ensuring complete coverage of the cell edges 

with adhesive, although this paradoxically may increase their arc rates (Ref. 7). Reduction of the 

SAMPlE data on arcing of cells without exposed adhesive may help settle this question. Comparison 

of the data with those taken in ground plasma tests (Ref. 14) will also be instructive. 

Figure 9 shows some of the arcing data on the anodized aluminum sample. As is typical of dielectric 

breakdown, the arc rate shows a strong dependence on the potential (in this case, we have fit the 

dependence by a power-law). Although these data have been normalized to a standard plasma 

condition and plate area, absolute rates must await reduction of the Langmuir probe data. It may be 

possible to determine an arcing threshold for this thickness of anodized aluminum in a plasma by 

further reduction of the complete data set. 

During the SAMPlE arcing experiments, every sample biased to high negative potentials arced at least 

once, except for a sample of gold nearby a grounded rod. It may be hypothesized that the lack of 

surface oxides and/or contaminants prevented arcing in this case, although this is still speculation. In 

proximity to insulating Kapton strips, even gold showed arcs. 

In addition to the experiments conducted to further our knowledge of arcing and current collection, 

SAMPlE also obtained payload bay pressure measurements to aid in the interpretation of the 

Experimental Investigation of Shuttle Glow (EISG) experiment data. Figure 10 shows some of the 

pressure data obtained for EISG, including times when the thrusters fired and gas releases took place. 

These events are clearly seen in the pressure record. 

SUMMARY AND CONCLUSIONS 

SAMPlE obtained excellent data on the arcing and current collection characteristics of solar arrays and 

space power materials in LEO. Further analysis of these data may furnish us with a better 

249 
understanding of, and better means to control, arcing on solar arrays and spacecraft in the LEO 

environment. The result could be better solar array or power system designs, ensuring years of safe 

and reliable operation in the harsh LEO plasma. 

In order to determine the behavior of even newer and more diverse materials and solar arrays in the 

LEO environment, it is planned to refly SAMPlE with different samples on the sample tray. This 

low-cost reflight may be one of our best near-term opportunities to learn more about LEO arcing and 

current collection behavior. 

ACKNOWLEDGEMENTS 

The authors would like to thank the Office of Advanced Concepts and Technology for footing the bill 

for SAMPlE. Their foresight and forbearance led to the great success that SAMPlE has become. Our 

thanks also go to the great SAMPIE team, all the way from technicians to project management, who 

built SAMPlE and ensured that it would work on-orbit. Like any great team, the watchword for the 

SAMPlE team was teamwork. Individual egos not withstanding, the SAMPlE team worked 

cooperatively and selflessly to see the project through to completion. Thanks to Thomas L. Morton 

for his fine reduction of the SAMPlE Langmuir probe and modified Space Station cell data. We 

would also like to acknowledge the contributions of Steven C. Zook, a brilliant young software and 

electronics engineer who met an untimely death before SAMPlE was completed. We will miss him. 

Finally, we acknowledge the excellent OAST -2 mission control team at Goddard Space Flight Center. 

We could not have asked for a more responsive and supportive team during our real-time 

reprogramming of the SAMPlE timeline. 

REFERENCES 

1. Grier, N.T. (1983). "Plasma Interaction Experiment II: Laboratory and Flight Results," in 

NASA CP-2359, p. 333, 1985 . 

2. Gabriel, S.B., Garner, C.E. and Kitamura, S. (1983). "Experimental Measurements of the 

Plasma Sheath Around Pinhole Defects in a Simulated High-Woltage Solar Array," AIAA 

Paper 83-031l. 

3. Ferguson, D.C. (1986). "The Voltage Threshold for Arcing for Solar Cells in LEO - Flight and 

Ground Test Results," NASA TM-87259. 

4. Jongeward, G.A. et al (1985). "The Role of Unneutralized Surface Ions in Negative Potential 

Arcing," IEEE Trans. Nucl. Sci., vol. NS-32, no. 6, Dec., pp. 4087-4091. 

5. Hastings, D.E., Weyl, G., and Kaufman, D. (1990). "Threshold Voltage for Arcing on 

Negatively Biased Solar Arrays," 1. Spacecraft and Rockets, 27, p. 539. 

6. Mong, R.L. (1993). "Arcing Mitigation and Predictions for High Voltage Solar Arrays," 

Master's Degree Thesis, Massachusetts Institute of Technology. 

7. Ferguson, D.C. (1993). "Interactions Between Spacecraft and their Environments," AIAA 

Paper 93-0705. 

8. Ferguson, D.C., Snyder, D.B., and Carruth, R. (1990). Final Report of the Joint Workshop on 

Evaluation of Impacts of Space Station Freedom Grounding Configurations, NASA LeRC, 

250 

l ______________ _ 
- ---- ----

August 21, 1990. 

9. Upshulte, B.L. et al (1991). "Significant Reduction in Arc Frequency of Negatively Biased 

Solar Cells: Observations, Diagnostics, and Mitigation Technique(s)," in NASA CP-3121, p. 

32-1. 

10. Cho, M. and Hastings, D.E. (1991). "Dielectric Charging Processes and Arcing Rates of High 

Voltage Solar Arrays," 1. Spacecraft and Rockets, 28, pp.698-706. 

11. Hillard, G.B. and Ferguson, D.C. (1993). "The SAMPlE Flight Experiment Final Technical 

Requirements Document," NASA TM-I06224. 

12. Hillard, G.B. (1993). "Experimental Measurement of the Plasma Conductivity of Z93 and 

Z93P Thermal Control Paint," NASA TM-I06284. 

13. Hillard, G.B. and Ferguson, D.C. (1993). "The Solar Array Module Plasma Interactions 

Experiment (SAMPlE): Science and Technology Objectives," J. Spacecraft and Rockets, 30, p . 

488. 

14. Bozich, R.C. (1994). "Characterization of Solar Cells and Metal Samples in a Plasma 

Atmosphere Using the SAMPlE Proto-type Model System," Master's Degree Thesis, Cleveland 

State University. 

251 
ORiGINAL PAGE 

BLACK AND WHITE P TO R~PH 

Figure 1. The Solar Array Module Plasma Interactions Experiment (SAMPlE) electronics box 

and sam pIe tray. 

Figure 2. The SAMPlE electronics box card cage and interior 

components. 

252 

I 

I 
EHCI..OIIUAE COIIEII _lilLY 

_PlATE 

IIIClNAL CONDITIONER 

IIOUN1lIIG IIIIACI(ET 

END PlATE AISEIIILY 

SPACER 

IIIClNAL CONDITIONER 

Figure 3. An exploded view of the SAMPlE electronics box and sample tray. 

STANDARD SlUCON 2an x 2an ColI 

MODIFIED SPACE STAllON S.Dan x ' .Dan C4Ia 4 Col C<lo.IpM 

_SVI8OR 

--

Figure 4. The SAMPlE sample tray. 

253 
APSA and SS Currents, Bay-to-Earth 

0,01 

.-.. --- 0,001 APSA Max 

~ 

(J) 

~ 

...... APSA Min c 

(]) 0,0001 ... - ... 

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0 SSF Max c 

4:i 

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0 

c 

e 

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1 E-09 

10 100 1000 

Biao Voltage (V) 

Figure 5, Electron currents collected by the APSA and Space Station arrays In the 

bay-to-earth orientation, 

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10 100 1000 

Bias Voltage (V) 

Figure 6, Electron currents collected by the APSA array In the bay-to-ram 

orientation. 

254 
Space Station Currents, Bay-to-Ram 

Experiment 20 

1 E-02 .0 ......... 

<C 

'-" 

~ 1 E-03 .5 .... 

() 

..Q1 1 E-04 

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.... 

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Figure 7. Electron current collection by the Space Station array in the bay-to-ram 

orientation. Plasma densities are overestimated (see text). 

·1 

From Exp's 10-13 

Ovemsng Dille 

4~------------------------~------------------------~ 

-11+--------.--~-L--~~--~_r------_.--------,_------_1 

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Figure 8. Electron currents collected in the bay-to-ram orientation by Space Station type 

cells modified to have varying amounts of coverslide overhang. 

255 
- - .--- - ---

SAMPlE DATA -ANODIZED AL 

Ram data #2 

1 E -1 3 ~ ~~ ~ ~ ~: ~ ~ ~ ~: ~~ ~~~~~~~~~ ~ ~ ~~:::::: ~ t::::::::::: ~ :~:~:~: ~::::::::::f ~ ~ ~: ~:::::t:::::: ~~: ~: ~~: ~ z:::::!:::: 1:::: ~ 

::::::::::::::::::::::::::::::::::~::::::::::::::: :::: :l::::::::::::f::::::::;:::::::f:::T::f::f::: 

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......................... ···>········t······7·····~·····~· ··· · 

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··:f::::::::::f::::::::j::::::f::::f::::r:::: 

• 

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.................................. ~., ...... ~ ....... . ·············l··········t······t·····t·····l····(·· 1E-16+-----------~------+---~--~--~~--~~ 

100 1000 

- Voltage 

Figure 9. One set of normalized arcing data on the SAMPlE anodized aluminum 

sample in the bay-to-ram orientation. 

Raw Pressure Data - SAMPlE 

MET day 11 

1 00 ::::::::±::::::::r:::::::I:::::::E:::::I::::::::t::::::::E:::::::±::::::::l::::::::: ·········>·· ···· ··-i-········i ··· ·· ··~······ ···i ·········>······· i·········>········i········· 

:::::::::f::::::::+::::::::t::::::::t:::::::t:::::::::f::::::::t::::::::t::::::::f::::::::: 

11 

::J 

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~ :: ::+::::::: ::: ::::!:: .. ~:·:····::f:::::::::~::::::::f::::::::+::::::::j::::::::: 

"-;FI FIFi L \ 1+--,~~--r-1--~~r--+--r--+-~ 

o 4 8 12 1 6 20 24 28 32 36 40 

MET minutQS aftQr 19:53 (4- /- 10 sQC) 

Figure 10. Raw telemetered pressure data taken during EISG experiment 

operation. 

256 
r

39196 ? 

/If N95· 20530 

PHOTOVOLTAIC ELECTRIC POWER APPLIED TO UNMANNED AERIAL VEHICLES (UAV) 

Jack Geis 

Wright Patterson Air Force Base 

Dayton, Ohio 

and 

Jack H. Arnold 

Rockwell International 

Canoga Park, California 

ABSTRACT 

Photovoltaic Electric-Powered Flight is receiving a great deal of attention in 

the context of the United States' Unmanned Aerial Vehicle (UAV) program. 

This paper addresses some of the enabling technical areas, and their 

potential solutions. Of particular interest are the long-duration, highaltitude class of UA Vs whose mission it is to achieve altitudes between 60, 

000 and 100,000 feet, and to remain at those altitudes for prolonged 

periods performing various mapping and surveillance activities. Addressed 

herein are studies which reveal the need for extremely light-weight and 

efficient solar cells, high-efficiency electric motor-driven propeller 

modules, and power management and distribution control elements. Since 

the potential payloads vary dramatically in their power consumption and 

duty cycles, a typical load profile has been selected to provide 

commonality for the propulsion power comparisons. Also, since missions 

vary widely with respect to ground coverage requirements, from repeated 

orbiting over a localized target, to long-distance routes over irregular 

terrain, we have also averaged the power requirements for on-board G&C 

power, as well as ground control and communication link utilization. 

In the context of the national technology reinvestment program, 

wherever possible we modeled components and materials which have 

been qualified for space and defense applications, yet are compatible with 

civilian UAV activities. These include, but are not limited to solar cell 

developments, electric storage technology for diurnal operation, local and 

ground communications, power management and distribution, and control 

servo design. 

And finally, the results of tests conducted by Wright Laboratory on 

ultra-light, highly efficient MOCVD GaAs solar cells purchased from EPI 

Materials Ltd. (EML) of the UK are presented. These cells were also used 

for modeling the flight characteristics of UAY aircraft described in Section 

3.0 and Table I. 

1.0 Solar-Electric VA V Background and History: Solar powered 

human flight has been accomplished, and needs no elaborate historic 

257 

-------- - - -------------
overview. It evolved from the ultra-light technology of AeroVironment's 

human-powered aircraft, and evolved from retrofitting similar vehicles 

with single-crystal silicon solar cell arrays and electrically-driven 

propellers. They sustained low-level flight for limited periods, and 

remained airborne largely at the whim of prevailing weather conditions. In 

fact, as a class they can be described as propeller-augmented sail planes. 

Later, in the fall of 1991, Eric Raymond's Sun Seeker aircraft, using Sanyo 

amorphous silicon cells deposited on polymeric film, logged a cross-country 

flight of 2,467 miles during a series of hops which totalled 119 hours of 

electrically-augmented flight. But again, at the risk of under-emphasizing 

the significance of these remarkable accomplishments, this aircraft was 

controlled more by the weather than the pilot, and the aircraft electrical 

propulsion system. 

1400 

1350 -C\I 

< 

::!: 1300 

~ 

~ 1250 

CI) 

z 1200 

w 

tz 1150 

a:: 

:5 1100 

0 

CI) 1050 

1000 

.JIll 

: 

: ~ --- : ./ 

,. 

· 

· / 

: 

: / 

'f 

J 

: 

· I 

: 

: / 

- SEALEVEL 65KFT 85KFT SPACE 

Figure 1: Relationship between altitude and solar intensity 

2.0 Insolation vs Altitude: If the extended mission VA V must takeoff 

and achieve operational altitude as a conventional aircraft, it must operate 

through a wide spectrum of solar intensity which varies with altititude. On 

the ground, even though the batteries are fully charged, the array must be 

sized to operate the propulsion subsystem at Air Mass 1 (AM-l.O) 

However, as the vehicle climbs higher, and begins to rise above the nearground atmosphere, it will eventually be receiving solar insolation which is 

80-90% of AM-D. 

258 
Figure 1 depicts a general spectrum of solar intensity vs altitude, and 

shows an approximate gain of 25% in solar intensity at operating altitude. 

This prompts the designer to seek various means of hybrid or aircraftassisted takeoff and climb-out augmentation, but in this paper, only 

electrical storage and solar power were modeled. 

3.0 Impact On Aircraft Performance of Solar Cell Efficiency and 

Weight: As a general observation, until recently the efficiency of solar 

cells was inversely proprtional to their weight. Space-qualified single 

crystal silicon cells, which are moderately efficient (14-16%,) are 

comparatively heavy, and require large wing-surface mounting areas in 

order to achieve a given power level. In later developments, when 

engineers were enticed by amorphous silicon cells deposited on thin, 

flexible films, because they were dramatically lighter in weight, it was 

necessary to devote 60% more surface area because of their lower 

efficiency. The applicability of using standard GaAs-on-Ge technology for 

higher efficiency (18% @ 28°C, AM-O)) will not improve the aircraft weight 

due to the fact that the resulting array weighs even more than singlecrystal silicon of a comparable thickness. It should be noted that certain 

U.S. companies are working on 3-mil GaAs-on-Ge technology, but this was 

considered too far term to use for a near-term comparison. However, in the 

present timeframe, a British company is demonstrating high-efficiency, 

thin-film GaAs solar cells which are comparable in weight with a-Siliconon-Kapton blanket material, or conventional blanket consisting of GaAs 

deposited on germanium substrates, then bonded to a Kapton blanket. The 

relative merits of cost for per watt for the various options are not 

addressed in the paper. 

It was considered useful in our analysis to review an existing UA V 

experimental aircraft (Pathfinder) and the associated weight and power 

algorithms that are derived from it, then calculate aircraft performance 

over a range of solar cell efficiencies and array mass densities. We then 

revised or updated aircraft component characteristics such as propulsion 

and storage conversion efficiencies, and solar celli wing area ratios, then 

compared the new over-all airplane performance using state-of-the-art 

solar cells with the new model using vastly improved solar cells. 

For this purpose, the authors interfaced with EPI Materials Ltd (EML) of 

the UK and used their ultra thin, ultra-lightweight MOCVD GaAs solar cells 

modeled for application to the new "Pathfinder" baseline aircraft. The 

results of this modeling are presented in Table 1. 

259 

~~-~ - - - - --- --
TABLE I 

EFFECT OF PHOTOVOL TAlC OPTIONS ON ELECTRIC AIRPLANE 

PERFORMANCE USING EPI MATERIAL LTD. (EML) THIN FILM GAAS 

CELLS 

FLIGHT CONDITIONS· NEW EML EML BIFACIAL 

·60 KFT BASELINE STANDARD THIN-FILM 

• Winter Solstice PATHFINDER THIN-FILM GAAS 

·24 Hr. Operation GAAS (ALBEDD-.24) 

Cell Efficiency (%) 21 21 24.6 

Cell Mass Density .0457 .0228 .0228 

(lbs!Tt2 ) 

Payload Wt. (Ibs) 100 100 100 

Propulsion Wt. (Ibs) 233.9 179.6 153.1 

Misc. Wt. (Ibs) 117.4 89.5 76.3 

Solar Cell Wt. (Ibs) 132.5 51 43.5 

Energy Storage Wt. (Ibs) 1060.4 822 701.1 

Airframe Wt. (Ibs) 274.9 233.7 197.6 

Total Aircraft Wt. (lbs) 1919.1 1465.8 1271.6 

Wing Area (ft ~) 3636.3 2790.5 2339 

Wing Span (ft) 208.9 183 169 

Storage Energy Density 126 126 126 

(Whrsllb) 

Payload Power (w) 250 250 250 

Misc. Power (w) 115.5 88 75 

Propulsion Power (w) 8352.7 6413.8 5467.9 

Total Power (w) 8718.2 6751.8 5792.9 

• Wright Laboratory Spreadsheet Model 

It should be noted that the UAV design challenge is not merely 

sustaining marginal near-ground flight conditions. It involves the 

optimization of an un-piloted aerodynamic vehicle and electrical 

propulsion system which will permit the UA V to operate completely under 

the control of the ground station, at extremely high altitude, for extended 

260 
periods, bearing a payload, which for purposes of our comparison, weighs 

100 pounds. 

The aircraft, depending on mISSIOn requirements and design restrictions, 

may be capable of taking off conventionally on its own power and climbing 

out to design altitude, or if practical, it may be assisted by auxilliary 

propulsion, or carried aboard a larger mother aircraft where it is released 

at altitude. A computerized ground-control system will then "pilot" the 

aircraft robotically on a prescribed course over specified surface targets on 

a real-time basis. The electric propulsion and control systems would be 

capable of maneuvering the vehicle on an accurate, pre-specified flight 

plan during normal environmental conditions, thus requiring an adequate 

margin of power for all-weather aerodynamic stability. 

It is clear from examining Table I that the use of EML ultra-lightweight, 

single-sided (high efficiency) solar cells reduces the present weight of a 

"Pathfinder" type aircraft significantly. Please note that Table I does not 

include performance for an aircraft with the type of solar cells currently 

used on the LLNL "Pathfinder" (mass density of 0.103 Ibs./ft.2.) For winter 

solstice flight conditions, this aircraft would weigh in excess of 6000 Ibs., 

and have a wing area greater than 10,000 ft.2, which is clearly an 

impractical design. 

Column No.1 of the table is for a hypothetical baseline aircraft, similar to 

"Pathfinder," but with significant changes and/or improvements to aircraft 

component performance characteristics, as follows: 

A. Propulsion efficiency is increased from 64% to 70%, which is considered 

feasible using available technology 

B. Energy storage efficiency has been increased from 45% to 65%, which by 

today's space and aircraft technology standards is quite conservative 

C. Solar cellI wing area ratio has been upgraded from 0.55 to 0.799 by 

mounting solar cells on surfaces inside a transparent wing (see Fig.3) 

D. Wing aspect ratio has been changed from 25 to 12 

E. Cell performance has been increased from 19.5% to 21 % efficiency, and 

cell density decreased from 0.103 to 0.0457 lbs./ft 2, which the authors 

note is presently unsurpassed by several potential domestic options other 

than the cells produced by EML 

261 

~ __________________ ~ ______ J 
F. The lift coefficient has been reduced from 1.0 to 0.78 to assure that the 

airplane travels at 20 ft./see above the theoretical stall speed 

In addition, the analyses were made for planes flying at worst optical 

conditions (winter solstice,) rather than summertime conditions. 

It should be noted that despite an arbitrary increase in "pathfinder" 

solar cell efficiency from 19.5 to 21 %, and reducing the cell mass density 

by 55%, the weight of this model could be reduced even further if EML cell 

technology were employed. For example, if the present cells were replaced 

by the EML single-sided cells, another 453 pounds could be saved, and if 

EML bifacial cell technology were employed, almost 650 pounds could be 

saved. It is also shown that the power requirements can be reduced from 

8.7 kW to 6.7 kW using the EML single-sided cells, and to 5.8 kW for the 

bifacial cells. 

7000 

6000 

5000 

4000 

3000 

2000 

1000 

o 

SOLAR-POWERED UAV'S 

21% SOLAR CELL EFFICIENCY 

126 W-HR/LB STORAGE EFFICIENCY 

Total wt(lbs) 

II .1 03 Lblft2 

~ .0457 Lblft.2 

~ .0228 Lblft.2 

Figure 2: Impact of improved solar cell weight densities on 

total aircraft weights 

Figure 2 illustrates graphically the effect of cell mass density on aircraft 

total weight for a 21 % efficienct cell. It is also evident from Table I that 

significant gains in energy storage round-trip efficiency will dramatically 

reduce the overall aircraft weight. For example, present advances in space 

technology now yield upwards to 80% charge/discharge efficiency, which 

when coupled to improved storage capacity density in watt-hrs./lb, could 

262 

~- - - ~ 
I 

I 

make a profound impact on total aircraft weight. Clearly, each small 

improvement in aircraft component weight results in a cascade effect on 

total aircraft weight, and the end performance the aircraft can achieve. 

A reduction in cell mass density means that the aircraft wing will weigh 

less, and the propulsion power (and weight) will be reduced to maintain 

the aircraft at nominal flying speed. Therefore, the aircraft can be 

proportionately downsized, including reduced wing area, resulting in still 

additional weight savings. For example, if the solar cell mass density is 

reduced by onlyl %, the aircraft weight drops by over 9 lbs. Similarly, 

when the solar cell efficiency is raised by 1 %, the aircraft weight drops by 

almost 54 lbs. 

4.0 WRIGHT LAB TESTING OF EML SOLAR CELLS : A 3.98 cm X 3.98 

cm, 16-volt, high-efficiency, thin film GaAs solar cell substring was 

delivered to Wright Laboratory for performance verification. The general 

configuration of the cells are shown in Figure 3. The substring is produced 

by creating a single solar cell which is 4 cm square, and then 

photolithographically dividing it into sixteen 1 cm square solar cells which 

are interconnected in series to provide the 16-volt nominal output. Clearly, 

the device could be left as a single solar cell, with even higher efficiency, 

but in most UAV applications the higher voltage substring would provide a 

convenient electrical building block. 

The conversion efficiency was measured and recorded at the Air Force 

Institute of Technology by PhD candidate Kitt Reinhardt using a l-kW Oriel 

Xenon solar simulator at AM-O conditions, calibrated using a JPL balloonflown standard GaAs cell. The I-V curve and the efficiency calculations are 

presented as Figure 4. As can be seen, the conversion efficiency of the 

planar solar cell was 20.3%. The Voc was 16.25 V, while the Isc was 31.85 

rnA, resulting in a fill factor of 0.84. 

The general structure of the cell is shown in Firgure 5 The cell has a 

significant weight advantage over standard cells inasmuch as bonding the 

cell to a I-mil coverglass will result in a cell mass density of less than .03 

lbs/square foot, which represents a factor of 3.5 improvement when 

compared with the current "Pathfinder" cell. 

263 
r t4--------':l3 98CM-~1 

3.98 CM 

Figure 3: General Configuration of EML GaAs Solar Cell Substring 

264 

L ~ __ ~_~J 
,-

-

c( 

E 

-

~ 

Z 

W 

a:: 

a:: 

::;) 

(J 

45 

40 

35 

30 

25 

20 

15 

5 

o 

Ultrathin 16 V EML GaAs Solar Cell AMO 

Conversion Efficiency Test Results & Calculations 

o 

Total Cell Area= 15.8 cm2 

Fill Factor = 0.839 

Dmax= 16.25 Voc X 31 .85lsc (rnA) 

= 0.518 W X 0.839 (ff)= 0.434 W 

0.434 W/1S.& cm2: .0275 W/cm2 

EIf.= .027 W/cm2/.1353 W/cm2: 20.3°,4 

, 

5 10 

VOLTAGE (V) 

~ 

15 20 

Figure 4: Current vs Voltage curve for EPI Ultrathin GaAs Cell 

265 

-------------------
25.4~ (1 mil) COVERGLASS 

SO.811 (2 mil) 

0.1 11 GaAs (etched away except for electrical contacts) T 

O.S 11 P-Doped Window Layer 80% AU200kGaAs ----1-------------------4 4-SIl 

O.SIl P-DopedGaAs-O.SIl-Dffusn Lngth=O.5-11l ~ 

-~--~ 

3 ·11"" 0 N-doped GaAs-Diffuslon Lngth.=211 

Back Electrical Contacts 

Figure 5: General structure of EML Ultrathin GaAs Solar Cell 

5.0 Bifacial Solar Cell Sunlight Conversion: As previously described, 

bifacial solar cell performance is achieved when the solar cell is designed 

in such a way that it can receive and convert solar energy to electricity on 

either surface. In a VA V application, the wing-mounted solar arrays could 

be constructed in a manner that would allow the Earth's albedo, which is 

20-30% of direct sunlight, to illuminate the lower surfaces of the solar 

cells, while the direct sunlight illuminates the upper surfaces. 

The level of albedo intensity is a function of the Earth's local surface 

radiometric properties, but it is significant up to an altitude of 1000 miles, 

and intensified by snow and other highly reflective surface conditions. If 

employed, this "bifacial" feature could add about 20 to 30% to the 

conversion efficiency of the solar cell array, thus permitting it to be 

proportionately smaller and lighter in weight. Conversely, this makes the 

design of the wing structure more complex, thus requiring a trade study to 

determine the relative merits of bifacial power generation and its impact 

on construction complexities. Figure 8 illustrates how bifacial solar cell 

technology could be applied to a VA V wing construction. 

266 
I~ 

I 

EML GaAs BIFACIAL SOLAR CELLS ARE 

ULTRA-LIGHT AND EFFICIENT (24%) 

TRANSPARENT FILM 

DIRECT 

SUNLIGHT 

ALBEDO/EARTHSHINE 

(20-30% OF SUNLIGHT) 

BIFACIAL 

SOLAR CELLS 

Fig. 6: UAV Wing Construction Can Support Bltaelal Solar 

Cell Insolation 

6.0 Advanced Control and Sensing System: An advanced solid state 

control and sensing system would monitor the sunlight and maneuver the 

aircraft, when permitted, to maximize the solar-electric conversion process. 

Another feature of the automatic sensing/control system will monitor 

current loading and tachometer rotational speeds of the propeller motor 

drives. When the motors are developing higher torque than necessary to 

maintain the prescribed flight characteristics, more of the power delivered 

by the solar arrays will be switched to the battery recharging system for 

night operation. Under these conditions, when the battery reaches a fully 

charged state, portions of the solar arrays can either be automatically 

267 

_J 
switched off the load buses, or shunted to a thermal rejection radiator .. In 

this manner, during favorable weather conditions, the propeller motor life 

will be extended by minimizing over-current and voltage conditions. In 

summary, the solar power/control system will (1) drive the propeller 

motors during daylight hours; (2) replenish the battery to a fully charged 

capacity; (3) provide required power to the housekeeping and payload 

buses, and (4) optimize the torque of the propeller drive motors to 

conform to the conditions of maintaining the flight plan under variable 

weather conditions. 

7.0 Summary: In summary, the use of extremely light-weight, highlyefficient, MOCVD GaAs solar cells, as described herein, represents an 

enabling technology which, in fact, allows solar-electric VA V aircraft to 

perform their mission, bearing reasonably heavy payloads, and of taking 

off in a conventional manner, climbing to high-altitude, and remaining 

aloft for long durations, whereas when they previously employed standard 

Si or GaAs cells they could not achieve these goals. 

8.0 Conclusions and Recommendation: Although the adaptation and 

retrofitting of existing ultra-light aircraft to unmanned solar-electric flight 

is perhaps the most economic and expeditious method of near-term 

demonstrations, it will not bring the VA V community nearer to achieving 

the goals of high-altitude, long duration, stable and reliable flight. What 

seems to be required to realize these goals is designing an aircraft from 

scratch, using existing 1994 aerostructure, control, and propulsion 

technology transferred from the space and advanced aircraft industry, that 

can takeoff from the ground, climb to high altitude, and remain there in 

stable and reliable flight for extended periods. This, of course, requires the 

enthusiastic support and dedication of a strategic government/industrial 

alliance in which key national Superlabs and industrial firms not only 

identify desirable mission architectures, but fully collaborate in these 

efforts in a comprehensive national, and perhaps international, VA V solarelectric initiative. 

268 

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:3 I 

N95- 20531 

POST-FLIGHT INVESTIGATION PROGRAMMES OF RECENTLY RETRIEVED 

SOLAR GENERATORS 

Lothar Gerlach 

European Space Agency-ESTEC (XPG) 

Noordwijk, The Netherlands 

Abstract 

In 1993 two ESA solar power generators were successfully retrieved from Space. 

• EURECA with its 10 panel rigid array in August 93, after 11 months in a 500 km orbit. 

• One of the two flexible Hubble Space Telescope (HST) arrays in December 93, 

after almost 4 years in a 600 km orbit. 

Both solar generators are undergoing separate Post-Flight Investigation Programmes (PFIP). These 

programmes cover investigations of all solar array (SA) components and mechanisms. Since both 

programmes have much in common, most of the component and material investigations are the same. 

Extremely valuable information on numerous essential subjects, such as atomic oxygen , radiation, 

meteoroid and space debris environment resulting damage, low cycle fatigue, material degradation etc. are 

expected to be obtained for both types of arrays. We will also be able to explain and understand the 

anomalies experienced on both solar arrays in orbit. The paper will outline both Post-Flight Investigation 

Programmes and will concentrate on reporting the first results and findings. 

Keywords: Solar Arrays, Post-Flight Investigations, Thermal Fatigue, Power Degradation, 

Meteoroid Damage, Space Debris, Material Degradation 

1 Introduction 

The EURECA and the HST solar generators are the first solar arrays brought back from space after being 

exposed to the LEO environment for a significant duration (EURECA: 10.8 months, HST: 43.3 months) in 

a well known orbit and orbit orientation. 

After its first mission the Eureca-SA was originally not planned to be investigated in detail. EURECA-SA 

was designed for 5 missions and the plan was to immediately refurbish it and prepare for are-launch. 

However due to excessive power degradation during the first mission a post-flight investigation programme 

(PFIP) has been defined on short notice. EURECA-SA PFIP benefited from the work already done to 

prepare the HST -SA PFIP. Already in 1992, its preparation started and the content was optimised during 

the remaining period until retrieval. Unfortunately only one of the two HST-SA wings could be retrieved 

during the HST's first servicing mission (Dec. 1993 ) and brought back to Europe for investigation. 

This paper is intended to give an insight into what is being investigated, reports on the preliminary results 

and what kind of results are expected, but concentrating on photovoltaic related issues only. 

2 Post-Flight Investigation Programmes 

It is of prime importance to the ESA to study these generators in detail. This provides a unique opportunity 

to study in depth the mechanical and electrical integrity of a retractable rigid panel array as well as of the 

flexible HST generator, following exposure to the severe LEO environment. It is expected to obtain 

extremely valuable and reliable information on numerous aspects, such as atomic oxygen (ATOX), 

meteoroid and space debris damage, low cycle fatigue, material degradation etc. It will also be possible to 

explain and understand the anomalies experienced in orbit. It will help to improve future solar arrays for 

both the rigid panel concepts and flexible arrays resulting in more reliable design and better protection 

against damaging effects such as random failures (e.g. short/open circuits) etc. The in-orbit performance of 

future arrays will be more predictable, the predictions more reliable and consequently the SA's can be 

designed and operated more cost effectively. The knowledge gained from the investigations is also 

important for the space debris and meteoroid community. The total surface of the retrieved HST wing and 

the EURECA solar-array is about 170 m2 , and provides a unique opportunity to study the craters of 

269 

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impacting particles, also to assess the impacts and potential damage and to refine the current meteoroid, 

and debris models. 

Both PFLIPs are co-ordinated and managed by ESA. The investigation team for the HST SA-PFLIP 

consists of all parties who were involved in the development and manufacturing of the solar arrays, i.e.: 

ESA/ESTEC, British Aerospace (BAe), Dornier, Contraves and DASA (formerly TST). For the EURECA-SA 

the team consists of Fokker Space & Systems, DASA Wedel and the same ESTEC team as for HST. They 

have the expertise for disassembly of the parts they have developed, minimising the risk that evidence is 

destroyed when handling or investigating the hardware. Where required, institutes and universities (i.e. for 

meteoroid and debris investigations) as well as the European Space Tribology Laboratory (ESTL) will be 

involved in the investigations and evaluations of the flight hardware. Most of the material investigations are 

planned to be done in the ESTEC Materials division. 

The key guidelines of these investigations are to ensure that no evidence is destroyed during SA storage 

and handling, or when samples are being removed. Both PFIPs are now in progress. The EURECA-SA 

activities will be completed in Aug. 94. The HST-SA PFIP is expected to be completed in the first half of 

1995. In May 95, the results will be presented at an HST-SA PFIP symposium at ESTEC. 

Although both generators are different in their mechanical construction (rigid panels and flexible roll-out 

blankets) they have a lot in common. The solar cell assemblies are the same (only cell dimensions are 

different, table 1) and most of the materials and components used for the electrical network are the same. 

Base material - Crucible crown silicon 

Base resistivity - 10 Ohm cm 

Junction - n-on-p shallow diffused 

Back surface field - p+ doping 

Back side reflector - Aluminium layer 

Contact system - Ti(Pd)Ag 

Anti reflective layer - TiOx 

Surface roughness - $1.5 microns 

Solar cell dimensions HST - 20.8mm x 40.2mm 

Eureca, Charge - 20.8mm x 59.4mm 

Eureca, Load - 26.8mm x 50.1 mm 

Table 1: Solar cell assembly (SCA) characteristics 

4 := __ ST ~!l_~_.: ___ . __ ~ __ . __ ._ .. ____________ ~ ______ ._, _______ ._., 

Table 2: SCA performance data (HST size) 

Silicon thickness 

BOL SCA efficiency 

SCA absorptivity 

SCA emissivity 

Cover slide material 

- HST: 250 jlm, EUR 150jlm 

- 14% (1AMO, 25°C) 

- 0.75 (unloaded) 

- 0.83 (hemispherical) 

- Soft Boro-silicate base 

glass (CMX) 

- 150 microns 

- Silicon, DC 93500 + Primer 

Cover slide thickness 

Adhesive for cell bonding 

Cover dimensions HST - 21 .0 mm x 40.5 mm 

Eureca, Charge - 21.0mm x 59.7mm 

Eureca, Load - 27.1 mm x 50.4mm 

.......,1'-"----""1"' silicone adheSive 

7-----~ silicone adhesive 

substrate 

Figure 1: Interconnected SCA 

Compared to EURECA, HST SA-1 investigations have the advantage that sufficient hardware can be 

provided for tests and investigation because there are no plans to re-use the first HST solar array_ It is 

foreseen to completely disassemble and study the mechanisms. All blanket buffers and at least one SPA 

are planned to be completely dissected and analysed. The remaining parts will be subject to specific 

examination with small areas cut out as required. 

The first activities in the programmes were the in-orbit inspections. Great emphasis was given to the 

photographic and video coverage during retrieval covering all SA areas and special close-ups of SAhighlights. A similar intensive photographic coverage were carried out after shuttle landing as well as initial 

insulation and continuity checks. Special attention will be given to the relative position of moving parts and 

other key features. 

Based on previous experience and knowledge the investigation began at the earliest possibility at KSC, 

where reference samples were cut from both solar arrays, sealed in dry nitrogen and hand-carried to 

Europe for test and immediate investigation. 

A further objective of the PFIP is to study changes of material properties and effects having a direct 

influence on SA power generation. 

Silicon adhesive has largely been used for the blanket protection against ATOX. Those protective layers 

will experience changes in their material properties and their thermo optical characteristic. This is caused by 

the radiation environment including UV radiation , thermal cycling and attack by ATOX. Corresponding 

material investigations will study these type of synergistic effects. 

270 
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As may be seen from the design description HST-SA has identical substrate surface coatings (DC 93500, 

fig. 6) on the front and rear side of the flexible blankets. Since the solar cell side was always sun oriented, 

the effect of the contribution of UV radiation to the synergistic environmental effects can be evaluated. 

UV radiation also tends to cause darkening or reduction in transparency of the transparent silicone 

adhesives if they are not shielded. DC 93500 is used for gluing the cover slides to the solar cells and a 

darkened adhesive will reduce the power generated. One important question is, did the CMX cover glass 

sufficiently shield the adhesive from UV radiation? 

Due to the low operation solar-cell string voltage of 37 volts (== 100 volts in Voc) plasma interaction or 

plasma sputtering is not expected to have an influence on the solar array hardware, but nevertheless it will 

be checked. 

It is of interest to check the silicone adhesives for surface hardening and embrittlement, including depth 

effects, and to which degree the polymerisation chains have been changed, and also if changes caused 

the unbonding of the glass-fibre cloth from the Kapton foil (fig. 6 and 11). 

How well the MoAg interconnector and its interconnections withstood the aggressive environment will be 

investigated. Are they showing signs of thermal fatigue (loop and weld areas) and to what extent did the 

silver erode in the ATOX exposed areas? 

Comparison between shielded areas (e.g. inside stress relief loop) with respect to the surfaces exposed to 

the ram direction of the ATOX flux, and how the thermal movement of the interconnector influences the 

silver erosion are important questions to be answered. The clamped and embedded parts of the 

interconnectors (e.g. between cover slide and cell) will allow the study of possible creep erosion effects. 

Further investigations being performed in the framework of the PFIPs are presented in section 4 together 

with preliminary results. 

3 Solar-Array Design Description 

The build-up of the two solar generators are only described to the extent which is needed to understand 

this paper or expected to be of interest to the SPRAT community. This implies that details of the 

deployment and retraction mechanisms are not discussed in this paper. For more design details refer to 

[ref. 1 to 6]. 

3.1 HST Solar-Array Design 

The Hubble Space Telescope (HST) was launched into a 614 kilometre low earth orbit (LEO) on 24 April 

1990. The HST is a joint NASAIESA project. ESA provided two major elements towards the project one 

being the Solar Arrays (SA). Under the prime responsibility of the European Space Agency, a group of 

European contractors led by BAe, with DASA (formerly TST) providing the flexible blankets, have 

developed and built the largest flexible solar generator to date. 

It comprises two double roll-out solar array wings (fig. 2, 3) which are deployable and retractable. Each wing 

is equipped with two solar array blankets (2.4m x 6.34m) carrying the solar cells which are protected from 

each other by an embossed Kapton cushion whilst they are stowed on a common storage drum. 

To unfurl the blankets an actuator motor for each wing drives the four bi-stem booms out (two per blanket, 

fig. 4), which in turn draw the blanket from the drum by means of a spreader bar fixed between the ends of 

the two bi-stem booms. The first set of solar array wings were successfully deployed on 25 April 1990 with 

the Space Shuttle Discovery and HST in a 614 km orbit. 

The HST-SA is designed to survive intact for at least five years in a ==600 km low earth orbit (30000 cycles, 

+/-10aOC). HST -SA is required to deliver at least 4.4 kilowatts of electrical power at 34 volts after two years 

in orbit. On 5 Dec. 93 (day of SA retraction) the solar array delivered 4.8 kW which is 6 % above predictions. 

Each of the four solar array blankets is made up of five identical power generating sections, known as the 

Solar Panel Assemblies (SPA, figure 8, 11) and four Buffer Assemblies which act as mechanical and 

electrical interface to the deployment mechanism. 

Each of the power-generating SPAs (1.1 m x 2.4m ) are equipped with 3 solar cell strings, each having 106 

solar cell assemblies (SCAs) in series, two of them with 8 single solar-cell rows in parallel and one of them 

with 7 single cell rows in parallel. The single cell rows for each string are connected via MoAg tapping bars in 

groups of 14, 15 or 16 cells. All of these groups are protected by flat solar cell shunt diodes (Si wafer 2 cm x 

2 cm with Au coated Ag in-plane interconnector and CMX cover) for shadow protection. They are 

271 
Figure 2: Solar array hot case configuration 

Figure 5: Solar-panel assemblies (SPA) 

Figure 3: Solar array in cold case 

configuration 

Fig. 4: Bi-Stem Boom 

Glass fibre filled wilh DC93500 Adhesive (DP46971) 

Kaplon H (12.5 microns) 

Adhesive (DP46971) 

~~~~~~~~ Silver Mesh 

~~~~~~~~ji~~ Adhesive (DP46971) 

Figure 6: Exploded view of the 0.21 mm - thick, 

flexible carrier substrate of the SPA, including 50 

micron-thick silver mesh for power transfer 

Figure 7: Upper IBA of STSA-1 with Ag power tracks for string 

connection, discolouration due to UV and location of anomaly 

Igure 8: Close-up of short 

circuit location (rear side) 

272 
mounted in the tooling gaps between the solar cell strings. On the Inner Buffer Assembly (IBA), figure 7, 

the individual solar cell strings are electrically connected by means of 75 micron thick silver foil strips to main 

and redundant connections for each SPA. They are routed to a flexible printed circuit board which serves 

as the interface to the harness attached to the deployment mechanism. 

The basic carrier substrate of the flexible blankets consists of a 210 11m thick atomic oxygen (ATOX) 

resistant glass fibre/Kapton compound (Fig 12). 

The power is provided by 48760 solar cell assemblies (table 1, 2 and fig. 1) consisting of silicon cell, cover 

and MoAg interconnector. They are bonded with silicone adhesive RTV-S 691 to the 20 SPA substrates. 

3.2 EURECA Solar-Array 

The European Retrievable Carrier (EURECA) is a unique, re-usable, user oriented space facility developed 

to meet the needs of both scientific and application oriented users. The Eureca solar array consists of 2 

fully interchangeable wings of 5 panels each (fig. 9. 10). 

The panels consist of a rigid aluminium honey comb structure with carbon fibre face sheets of 1.4 m width 

and 3.4 m length. The thickness of the substrate is", 22 mm (fig. 11). 

The panel size requires 6 hold down pOints of 100 mm diameter in-side the panel area for stowed 

conditions and 4 edge brackets for the deployment mechanism. On the front side the panels are equipped 

with the solar cell network bonded with the standard silicone adhesive RTV S-691 to an insulation layer (fig. 

11) which protects the structure against the atomic oxygen environment. For the solar cell layout ATOX 

resistant MoAg interconnector (Ag-coating required for welding) have been chosen. In the area exposed 

to the environment the silver has been removed from the cell interconnector (except Panel F5 which was 

planned to be a qualification panel). On the rear side an Aluminium/Kapton/ITO layer is bonded to the panel 

for protection against ATOX erosion and to improve the thermo-optical behaviour of the panels. 

The power conditioning design required that the panels are split into a load (main bus) array part and a 

charge array part as shown in figure 12. 

Figure 9: EURECA-SA wing after retrieval at Fokker Space & Systems 

273 ON ·IN l PAGE: II 

OF '-0 . Al;1T'( 
40%charge sections 

and load array 

Figure 10: EURECA solar-array configuration (deployed: 1 Aug. 92. retracted: 24 Jun. 93) 

1 50llm CMX cover slide " with 30llm DC 93500 

1 80llm BSFR solar cell 

261lm MoAg dE5~~~5C5~~:::~;;' RTV $·691 interconnector ~ 551lm glass fibre with XB 3052 

12.51lm FEP 

39B 44s 

lp lp 

Kapton H 

2.51lm FEP 

22mm AI·honeycomb 

120llm CFRP (non·reinforced areas) 

Figure 11: Panel cross-section 

+ + _ + _ _ + lp I ~~~----------~ load .. -..... ~ load· 

section charge section 

section 

Figure 12: Solar cell string layout of solar array 

panel (panel half) 

Each panel is equipped with 4 charge strings with 124 cells in series (124s). These charge strings are built 

up of 3 single cell rows with intermediate parallel connection via MoAg - tapping bars in groups of 6 to 12 

cells. All these groups or shunt intervals on panels F1. F5. F6 and F10 are protected by flat solar cell shunt 

diodes (0.2mm x 20mm x 40mm) against local hot spots since shadows were possible on those panels. To 

reach the optimum cell packing factor the shunt diodes are located on the rear side of the panels. 

typical load 

array WCP 

stress relief 

loop after 

every two 

5 tre 55 re II e f 10 a p 

wi dth Ihlckneaa 

~ :: l&O!Unc""mlt.' ... n' .... m ......... ., . 1251!m1l5,.."", .. n/75,..AG/2!omKII", ... FIPI 

~z:::;m;m;:=====4=========j. 60~ (2It.am ~on/2~ KtiplDn FEP) 

not to scale 

solder joints stress relief 

loop radii = O.2mm 

typical charge =====~ array WCP 

Figure 13: Wiring Collecting Panel (WCP) arrangement Figure 14: Design details of WCP 

The build-up of the load array is following the same shadow protection concept with its shunt intervals but 

different number of cells per string in para"el. Due to geometrical constraints the cell dimensions for the 

load array are different to the charge array type. 

Load array string are build-up with 91 and 88 solar cells (26.8 mm x 50.1 mm). respectively. connected in 

series to provide the required power of 2500W after 1 year at 30 V at any condition during the sun-lit 

phase. 

274 
- - - - - 

- - - - - --- - - -

Each of the solar cell sub-strings (Fig. 5) or cell matrices is equipped with silver bus bars (coated with RTV 

8-691 for ATOX protection) at their ends to form a solar cell string. The solar array strings of the load- and 

charge array are separately wired through holes in the panel to a wiring collection panel (WCP) on the panel 

rear side where the individual solar cell strings are connected in parallel (solder connections) to form 

electrical sUh-sections. The WCP design and wiring connection concept is given in figures 13 and 14. 

4 Post-Flight Investigation Programme - First Results 

4.1 HST-SA In-Flight Power Generation Anomalies 

During its almost 4 years in-flight operation there were 5 power generation related anomalies on the 4 

blankets. Two disconnected solar cell strings from which one recovered, a short within a solar cell string of 

SPA -CC (fig. 3) which also recovered, a short between a power circuit and a temperature sensor circuit and 

a short between two SPAs and structure. The intermittent short circuit on SPA -CC and the last anomaly are 

on the retrieved wing (-wing or -V2 wing). 

Until now there are no results available on the -CC anomaly, however the last failure seems to be traced. 

The failure appeared on 5 December 1993 while astronauts attempting to close the -V3 Aft Shroud door 

during HST's First Servicing Mission (FSM), EVA-1 . 

Having reviewed the anomaly history of both solar array wings together with the recent in-flight data it was 

concluded that there was a short from the positive rail of SPA-C (SA section 4) to the return rail of SPA-E 

(SA section 3) and a short from the newly formed SPA -C/-E circuit to structure (fig. 15). 

The above shorts are theoretically possible at several locations, due to the physical layout of the blankets. 

This is limited to SPA-C and the IBA of the upper blanket (-wing). Other theoretically possible failure 

locations (PCU, diode box, harness) were investigated but the in-flight data was never in agreement with 

possible failure scenarii. 

ARRAY 

TRIM 

RELAYS 

BYPASS 

CCC 

RELAYS 

K31 

K32 

K42 

BATTERY 

RETURN BUS 

section 3 ,- - --- I"III~ ,... -- - - . 

r--,'" on' , 

, ~ "-- ...JI ' 

section 4 RELAYS ,- - - - -

L..--------++---rL..HIIIIIIt-: ----- . 

Figure 15: STSA-1 Section 3 and 4 logic with 

short circuits indicated 

LOAD 

LOAD 

Figure 16: Drum with cut-out and bum mark 

After HST SA-1 retrieval electrical health checks were performed at KSC and at BAe. There was a short 

circuit between SPA-C and -E as well as low resistance between SPA-E to ground. The visual evidence 

was only available after completion of flash testing and after disassembly of the blankets from the drum. 

There was a burn mark on the IBA where the SPA-C positive rail crosses the return rail of SPA-E (fig. 7 & 8). 

The cross-section of the IBA substrate is the same as for the SPAs (fig. 12), except that the Ag-mesh is 

replaced by: 75/lm thick silver strip/ 25/lm Kapton/ 75j.l.m Silver strip. The location of the burn mark on the 

IBA is more visible on the rear side of the blanket at the interface to the flexible harness. This flexible 

harness is routed to a printed circuit board (PCB) located in a cut-out slot of the drum (fig. 16). Where the 

IBA is in contact with the edge of the cut-out, the burn mark on the IBA coincides with a burn mark on the 

drum. 

This indicates that the observed shorts during FSM are on the upper blanket at the expected locations, 

showing also that both shorts (between SPAs and the short to ground) were caused by a single failure. 

275 
Presently the following failure modes and conditions (single and combined) are being investigated: 

a) Sharp edges on silver wiring strips due to manufacture by guillotine. 

b) Manufacturing induced effects (e.g. pre-damaged, torn/bent silver bars, foreign metal inclusion) 

c) Wear of IBA, and flexible harness and cushion surfaces caused by cut-out edge and PCB cover. 

d) Creep of Kapton (25 micron insulation between power tracks) due to local pressure on the IBA and 

by the interface to the PCB cover. 

e) Thermal cycling effects. 

f) Electrical fields. 

All non electrical anomalies which were observed are not discussed here but most of them (e.g. tension 

sensor failure, SA jitter and stick-slip effect) are described in ref. 7. 

4.2 EURECA-SA In-Flight Anomalies 

Since 14 Aug. 92 (day 227) the solar array has continuously experienced anomalous power degradation 

due to unexpected in-orbit failures (one short circuit and numerous open circuits).Two examples are given 

in the following: 

On day 263 of 1992 it was observed that during the first minutes after entering eclipse the current sensor 

of section 2.2 indicated a current from the battery to the solar array (no blocking diodes were put between 

the SA and battery circuits). Figure 17 shows the actual current profile. The current profile with its obvious 

sensitivity to temperature changes is typical for a short circuit. An evaluation of the in-orbit data indicated 

that the short is present at about half the string length of a solar cell string (40% branch of section 2.2) with 

three solar cells in parallel and 124 cells in series. 

14 

12 

10 

8 

'li6 

E 

::;4 '" 

c 

~ 2 ::> 

u 

~o 

~ -2 

I 

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CA 1.2 current 

CA 2.1 cu rrent 

CA 2.2 current 

~~~~~~~~~~~~~~~~~~ 

520 530 540 550 560 570 580 590 600 

time (minutes/day 263/92) 

Figure 17: Charge current profile at time short circuit 

16 ,---------------------------------~ 

charge current' . , 11 .. ····· ................ :::~:~:.::···· charge current 1.2 / ·' .. ···· .. ·· ........ 1 ......... : 

'.J I ~ 

121t~--~~====~~-------tV~--~~~~=-==~~~~!.:~1-1 

~ 8 -tt--------i-_It----t------i--+-i 

::> 

u 

'" e> 

~ 4 -tt-------t-~t---_It------~+--l u 

~~_r~~~\· ·~ .. ·~ .. ·F=r=r=~~~-r-r~~ .. ·~ .. ·±·~ 

670 690 710 730 750 770 790 810 

time (minutes/day 265 of 1992) 

Figure 18: Orbit with first open circuit failure 

830 

After retrieval a large burn mark was found at the expected short circuit location on panel F7 (fig. 19). The 

size of the burn-out area is unexpectedly large and indicates that very high temperatures must have 

occurred at this spot (>2000 °C). In some areas the glass from the cover and the silicon is melted. 

Adhesives and insulation foil are carbonised. X-ray photographs also reveal that the honeycomb of the 

panel has locally been damaged below this area. Experts on arc progression are involved in the 

investigation of this unusual burn mark. The triggering mechanism(s) of the burn mark is not yet found. 

Tests are planned to reproduce this anomalous event. 

Between 21 Sept. 92 (day 265) up to the end of the mission a certain number of open circuits were 

observed, first in the charge array sections and later also in the load array. From 21 Sep. 92 onwards 

sudden drops of == 1.5 and == 3 amps were observed in the charge current profile (1.5 amps corresponds to 

the current generated by one solar cell charge string). In the beginning wing 1 charge sections were 

affected and only towards the end of mission wing 2 solar cell strings started to fail. In some cases the loss 

was immediate, in other cases the string toggled before the single or double string was lost. In all cases the 

first drop in current generation was within the first three minutes after entering orbit day. A typical sample of 

recorded current profiles, showing an open circuit failure is provided in figures 18. The cumUlative charge 

array power loss during the mission is given in figure 21 . 

276 
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Figure 19, left: Burn mark at short circuit 

location ($hown area is '" 1 0 cm x 10 cm) 

50,---------------------------, 

§: 

., 

., 

40 

E 30 

C 

~ :l 

U 

.. 20 0> 

:;; 

s:: 

u 

10 

totat charge current loss 

wing 2 -not yet fully explained 

wing 1 (open circu~s) 

wing 2 (open rcu~s) 

wing 2 (short circuit) 

, ....................... .1 ......... _ .. . : ....... .. 

:_r-.. ·i" .... · .. 

!.! 

__ '. ~.r- -' -------_. -' _. _. -. _. _. -_. ---r" 

o 30 60 90 120 150 180 210 240 270 300 

MET (days) 

Figure 21: Cumulative charge array 

Figure 20, right: Part of a WCP showing early 

effects from arcing in stress relief loop 

power loss during the mission Figure 22: Completely destroyed WCP 

Due to the numerous anomalies in the charge array it was also expected that in the load array several solar 

cell strings could be lost. There is no easy method to measure the load array output because the required 

power from the load is most of the time below the actual capabilities of the load array and the string/section 

currents are not measured. Only at times of high power consumption, above the capabilities of the load 

array, the actual load array performance can be roughly calculated using the total load current and the delta 

277 

I 

I 

I 

I 

I 

I 

I 

I 
current to be provided by the batteries. The last possibility was in Jan. 93, shortly before all payload 

experiments were successfully completed. It could be calculated that at this time about 24 % of the load 

array capability has been lost. 

From the investigation into the loss of solar cell strings it was found that the power loss is a result of failures 

in the wiring collecting panels (silver bus bars) located on the rear side of the panels connecting solar cell 

strings in parallel. The failure is a common mode failure occurring in all WCPs. The failure occurred in the 

stress relief loop and all parties involved in the investigation agree that the failure is a fatigue problem due 

to inadequate bend radii in the stress relief loop. Figure 20 and 22 give an indication from the degree of 

damage seen on the WCPs. It covers the complete range of damage from broken loops (only detectable 

with X-rays) still conducting current, first sign of arcing until complete destruction of the WCPs with short to 

panel structure. Plasma and arcing experts will also study this type of arc progression. 

The reason why this has not been detected prior to launch is, that the WCPs were neither tested in a fully 

flight representative configuration nor were the loops inspected after test to the required detail (X-ray). 

However, this failure would be easy to correct for a re-flight. The failed components can easily be removed 

and replaced by reliable WCPs. 

Another clear early conclusion from the WCP related investigation is that the acoustic test performed only 

on wing 1 cannot be made responsible for any of the failures. If a proper design would have been chosen 

the effects from the acoustic test would not be visible. In our case, we are only "lucky" that we had this weak 

design allowing us to correlate the acoustic test and the launch environment with fatigue effect as a result 

of in-orbit thermal cycles. This helps us for the future to better consider the total mechanical stress on the 

electrical network. 

However, in light of the extent of the failures, all payload and mission objectives were met. The flexible 

nature of the mission planning and payload together with adaptable operational methods allowed work 

around solutions to be incorporated as the mission and failures progressed. 

4.3 Power I Solar-Cell Degradation Studies 

For all satellites, reliable solar-array power degradation stUdies are needed, and it is important that all 

damaging mechanisms are considered in the power budget predictions. The goal is to fully understand 

these power degradation effects. This give us the chance to minimise their effects on the solar array during 

the design phase, or even eliminate them in some cases. For this reason great emphasis will be put in 

studying all the negative impacts on electrical power generation. 

Extensive radiation studies and tests will be performed to predict the total accumulated fluence and 

evaluate the effects of electron, proton and photon radiation on the solar-cell assemblies. 

The most interesting parameters for the Space Telescope cell type, in combination with the 150 micron 

thick cover and resulting degradation characteristic for 1 MeV electrons/cm2 are shown in table 2. 

These are well known and well established characteristics for solar cells radiated with 1 MeV electrons. The 

radiation environment is also well known, but the uncertainty is, if the proton and electron spectrum is 

correctly converted to the 1 MeV electrons. Ground tests for comparison will be done refining the power 

predictions and power verification method for future solar arrays. 

When, solely the effects of radiation are used to evaluate the power during lifetime the resulting total 

degradation would be under-estimated. 

The monitoring of in-flight generated power at the operation voltage does not allow to differentiate 

between power degradation caused by radiation and other damaging effects. Thus only a total power 

reduction is observed. The solar cell efficiency also changes with the solar cell operation temperature. 

Parameters which alter the array temperatures are the cell efficiency itself and the thermo optical properties 

of the materials used on the outer surfaces of the array. 

In addition to the natural environment (radiation temperature, ATOX) further potential impacts on the power 

output have to be considered. These are solar cell orientation errors, solar cell mismatch, meteoroid and 

space debris bombardment and random failures, which include open circuits as well as short circuits (not 

caused by meteoroid and space debris). These effects are considered in the form of current and voltage 

loss factors applied to the solar cell network. Except for the sun intenSity the individual current or amperage 

.278 
loss factors typically considered were based on ground tests and best engineering estimate, but worst 

case and never expected to occur. Those used for the HST solar generator are summarised in table 3. 

Years in Orbit BOL 1 2 3 4 Table 3: Current or amperage loss factors 

Calibration Error .980 .980 .980 .980 .980 for HST-SA deployment on 25th Apr. 1990 

SCA mismatch 1.00 .998 .996 .994 .992 

Orientation Error .996 .996 .996 .996 .996 

Random Failure 1.00 .982 .970 .960 .950 

Meteoroid/Debris 1.00 .995 .990 .985 .980 

Total .976 .952 .934 .917 .901 

The power profi le prediction for one typical SA section and actual in-flight data are given in figure 23. After 

43.3 months in orbit in-flight data show currents 6% above the prediction. 

The in-fl ight power generation of HST-SA is continuously monitored and recorded. However, only 

readings taken in "cold case" configuration (fig. 3) are suitable for the power degradation studies. Others 

solar array slew angles or HST off-nominal roll angles result in undefined reflections or shadows on the solar 

array. Due to the numerous reorientations of the HST this for the power verification ideal configuration is 

not frequently approached and also angles close to this configuration had to be used to monitor the power 

degradation. Additional power generation due to reflections can reach 4 % (total array output). 

It is getting increasingly important for economic reasons to adjust these worst case factors, which normally 

cover all solar array eventualities, to more realistic degradation factors. STSA-1 offers us the possibility to 

investigate which losses were considered too pessimistic or too optimistic, or may be there are others to be 

considered in the future. 

For example, the effect of particle impacts on the power generation can be checked by a laser scanning the 

surface of the SCA. The distribution and number of impacting particles can be provided by the 

corresponding numerical models. 

A darkening of cover slide and its adhesive by UV, if any, will be verified by means of spectral response 

measurements 

Reliable power/solar-cell degradation studies can only be performed with the in-flight results from HST-SA 

because there was no measurable degradation on the EURECA-SA when measuring on solar cell string 

level using the Flasher Test Equipment. Eureca power output measurements compared to pre-flight data 

were between +5% and -1.6%. With the known uncertainty due to electrical contact difficulties, all results 

are within the measurement accuracy of flash test (+/- 1 %). 

For HST a LAPSS (flash-test) was performed on the integrated solar array blankets. The results are shown 

in figures 24 to 26 together with the pre-flight data. The Pmax data were intentionally left out because at 

this test level they are not suitable for power degradation studies. There are too many unknowns due to 

blanket wiring and harness. The results for the currents are very homogeneous. There are insignificant 

small variations between the SPAs. The measured degradation for the two operation points (short circuit 

and at 34 volts) is 3%. This is including all cell damages (particle impacts etc.). The results are as expected, 

but not in line with the worst case predictions. The worst case relative degradation outlines 10% for these 

operation points. 

For the open circuit voltages the degradation is smaller than expected. It varies between 0.2% and 0.7% 

(fig. 25, due to the short in SPA -C no measurement available for SPA -C). 

24.0 7 .80 -,-------------------, 

23.6 

~ 23.2 

.§. 22.8 

~ 22.4 

B 22.0 

N 21.6 c 

,g 21.2 u 

~ 20.8 

20.4 

o 

• 

• section 2 currents 

q undefined HST reflections 

o 

o 

predicted section current 

• • 

20.0 +-__,_-_r_~___,r_-r---,---r-_r---..,.---...,.---,-__1 

o 365 days in orbit 730 1095 

Figure 23. Power profile of SA section 2 

7 .75 

'ill 

7.70 

'" ";;;' 7.65 --+-

<{ 5i 7.60 

'0 7.55 --- u 

!E 7 .50 

7.45 

LI;l 9 Y ~ 'f w a () CD 

LI;l 9 Y ~ 

<{ <{ <{ <{ ~ <{ <{ <{ <{ 

5i 5i 5i 5i w 5i 5i 5i 5i 

Figure 24. Short-circuit current comparison 

279 
Although above results give already a clear indication of the much smaller cell performance degradation 

when compared with the worst case predictions, more detailed power measurements are planned on SPA 

level, string level and on individual single cell rows in shunt intervals. The results will be compared with preflight data which partly consider the initial matching classes distribution for individual single cell rows in 

shunt intervals. 

Two different standards (JPL ST -OS, and Telecom 10/5) were used in parallel during the HST -SA 1 

programme. Because of the low degradation it will be difficult to reliably assess the actual level, but it will be 

attempted. To evaluate the relative degradation is much easier. 

63.1 Ci) 7.75 co. -It- pOSI landing E 63.0 --+- ~ 7.70 > 

62.9 ..,. 

cry 7.65 

~ 62.8 m 7.60 0 C --+- pre launch > 

;; 62.7 ~ ::> 7.55 -It- pOSI landing 0 u 

> 62.6 c 

.Q 

7.50 

62.5 ~ 7.45 Ql 

a. 62.4 0 7.40 

U;J 9 Y CJ? -;: w 0 0 CD « U;J 9 Y CJ? -;: w 0 0 CD « U;J 9 Y CJ? -;: w 9 y CJ? -;: 

« « « « « « « « « « « « « « « « « « « « a.. a.. a.. a.. a.. a.. a.. a.. a.. a.. a.. 0.. 0.. 0.. 0.. 0.. 0.. 0.. 0.. 0.. (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ (fJ 

Figure 25. Open-circuit voltage comparison Figure 26. Operation (34v) current comparison 

4.4 Results from Material Investigations and General Observations 

HST-SA 

The blanket compound seems to be in a good bonding state. No delaminations of the carrier substrate 

were visible from the front side. This is planned to be verified by a vacuum test on a small coupon sample. 

The piano hinges are in good shape. No rupture of hinge loops were observed . In one case the hinge rod 

penetrated the sliding protection substrate ("" 1 cm). There will be a detailed inspection of the contact pOint 

of the hinge rod and the protection substrate, and a measurement of the force needed to push the hinge 

rod through the restraint tab. 

No delamination of the GFRP stiffeners have been observed. 

The RTV S-691 coatings show different degrees of discolouration . This is very obvious where the 

stiffeners and meander bars are next to each other. The coatings on the stiffeners are generally much 

darker than the coatings applied on the silver bus bars or meander bars. A thickness measurement of the 

coating might explain this difference 

The SCA interconnectors are in good shape. The colour when inspected with unaided eye is similar to the 

interconnectors seen on EURECA. Since Eureca had the silver removed in the ATOX exposed areas it 

would be no surprise when on STSA the silver in the ram direction is completely gone. ESTEC Materials 

division is taking care of this Investigation. 

The shunt diode assemblies are in good shape. There is no obvious darkening of the DC 93500, but 

spectral response measurements will still have to be made (i.e. on SCAs). Considering that the gold coated 

in-plane diode interconnectors were exposed to almost 4 years in the LEO environment they are looking 

very good. The gold coating was obviously efficient and provided sufficient protection against ATOX. 

DC 93500 was also used as ATOX protection for the flexible substrate (fig. 6). Its darkening under UV light 

is known and it was no surprise to see a strong discolourations of the blankets. The most pronounced 

discolouration can be seen on the upper IBA (fig. 7). The portion which was exposed to UV light is very 

dark over a short stripe of == 10 em , at a location where there were reflections from the cushion roller. 

Temperature effects on the discolouration are also seen at the location on top of the power tracks. Other 

areas were less dark. The unexposed areas well' almost as new and had the original colour. 

280 
Contrary to the Eureca-SA, no effects from arcing have been observ8d on the front side (unaided eye 

inspection). The rear side is not yet accessible. In the frame ··.vork of our PFIP potential arcing areas will be 

investigated in detail. 

EURECA-SA 

As for HST-SA there are no further special observations concerning the bonding integrity to report. 

There was no significant change of thermo optical properties on solar cell assemblies (SCAs) other power 

generation related surfaces (i.e. aluminised Kapton ITO). 

The investigation has also revealed that the bus bars and their connection method are containing 

weaknesses in the design which will fail after a re-Iaunch. Some have failed already. However, the total 

extent of the bus bar failures can only be established after evaluation of the X-ray photographs taken from 

all critical and accessible areas. To date only the worst areas are visible, since all bus bars are coated for 

ATOX protection with a non transpare:·,t adhesive (RTV-S691). Metallurgical investigations are also 

foreseen. But from initial observations it can be concluded that the bus bar design is not fully reliable. In 

addition, the external forces induced from the cables soldered to the bus bars have been underestimated. 

For EURECA-SA the conclusion is that without correcting the bus bar problem with its interconnection 

method, a re-Iaunch of the solar array is not possible and a reliable repair method must be found. 

The results from the material investigations have not yet been completed, except for the reference 

samples cut at KSC from panel F5 having silver coatings exposed to ATOX. The results have already been 

published in reference 3. Different types of silver erosion have been found . Changes of surfaces are 

discussed which includes contamination (see also reference 8). 

4.5 Investigation of mechanisms 

HST-SA 

After completion of the first checks (at KSC and BAe, Bristol) and a deployment and retraction cycle all 

three mechanisms (Primary Deployment Mechanism, Secondary Deployment Mechanism, Solar Array 

Drive) will be dismounted for detailed inspection and investigations. 

In general all moving parts will checked against their pre-flight performance, for evidence of wear and 

damage. 

Torques and frictions will be measured where applicable. 

The state of lubricant in each bearing will be assessed. Fluid reservoirs and possible fluid creep will be 

examined as well as surface treatments. Wavy washers/preload devices will be checked for deformation, 

adhesion etc. Gears will be investigated for state of lubricant gear wear and wear of ceramic gear carriers. 

Static adhesion or fretting will be examined on all clamps/end stops. 

Electrical contacts will be checked. Motor currents including speed/torque characteristic will be studied as 

well as brushes and commutator surfaces 

The above list is not exhaustive but reflects the detail to which we intend to study the mechanisms. 

EURECA-SA 

The EURECA-SA deployment/retraction mechanism worked like a clockwork. All deployment - tensioningretraction cycles, prior to launch, in-orbit and post-flight were in excellent agreement. No adjustments are 

needed for are-flight. 

4.6 Meteoroid and space debris investigations 

The investigation of space debris and meteoroid particles and their distribution is one of the major tasks 

and of great help to the space debris and meteoroid community in refining their models and knowledge on 

damaging effects of impacting particles. 

EURECA-SA 

The meteoroid survey on both wings have been completed in Dec. 93 and the evaluation of the craters is 

under way. So far it is known that on the outer panel of +X wing 165 out of 3108 solar cells have been hit of 

which 68 have a larger impact feature than 200)lm (see also ref. 9). No particle has penetrated the panels 

completely. Figure 27 shows typical impact craters from the HST and Eureca solar array. 

281 
Figure 27. Typical particle impacts 

Left: Largest impact feature on EURECA-SA (solar-cell grid finger spacing for all cells = 1.2 mm) 

Centre left: Impact on HST blanket with particle penetrating the blanket at the cell interconnector 

Centre right: Impact feature on HST blanket (rear side hit, no particle penetration) 

Right: Impact feature (== 4 mm) of particle penetrating the HST blanket (rear side view) 

HST-SA 

After an exposure time of almost 4 years in orbit about 40000 particles greater than 10 microns are 

expected on the == 70 m2 of the HST-SA1 wing (counting both sides of the array) from which several 

hundred will penetrate the blankets. 

It is foreseen to examine all solar array units for highlights and use only the two blankets (still == 60 m2) for 

the systematic documentation of crater and impacting particles. The survey of both blanket front sides have 

just been completed (10 June 94). Based on the high number of impacts the plan is to document on the 2 

blankets impact features 2100 microns only. The cataloguing and investigation of smaller particles will be 

limited to one SPA (== 5 m2) and one buffer assembly (== 3 m2). As already mentioned, we will not exclude 

special highlights and the thermal covers on the mechanisms. 

General 

The solar array front sides with the solar cell cover slides are an ideal surface to record particle impacts. 

Depending on speed and size of the impacting particles, the diameter of impact features can be up to 30 

times bigger than the particle size. 

Through dedicated "in-situ" experiments on spare material calibration measurements are performed to 

study crater size versus particle size etc. 

With some luck we may find a sufficient number of trapped particles al:owing us to perform a chemical 

analyses in order to distinguish meteorOids from man-made objects, with a statistically significant 

distribution. 

A major goal of this investigation is to study which size of craters have an influence on the power 

generation caused by particles hitting the SCA surfaces. Also if particles can cause short or open circuits 

between the electrical network and the carrier substrate or trigger some other failure mechanism. 

Until now, there is no evidence that any meteoroid or debris impact has caused any of the observed 

failures. Some of the particles hitting the solar cells have damaged locally the silicon, but the resulting 

effect is that the chipped off portion or broken portion of the cell simply does not contribute anymore to the 

power generation. Power measurements on wing level indicate that the remaining cell area functions 

normally. 

282 
5 Conclusion 

The sOlar-array Post-Flight Investigation Programmes for EURECA and HST have already proven to be 

very valuable for ESA and its future solar-array projects. 

The mechanical performance of the retrieved hardware was in excellent agreement with the pre-flight data. 

The selected materials and coatings for the protection against atomic oxygen are suitable for long duration 

missions in LEO with high concentrations of atomic oxygen. 

Design weaknesses due to thermal fatigue could be identified and design improvements were 

implemented into the running Polar Platform and ISO project. 

The electrical degradation of solar cells was somewhat less than expected. 

Ten thousands of particle impacts are recorded on the arrays with several penetrating the flexible blankets. 

Despite the numerous clearly visible impacts there is no indication that any meteoroid or debris impact has 

caused any of the observed failures or other unusual power degradation. The loss of power due to local 

damage of the silicon is not measurable on panel level. 

All PFIP activities will be completed by May 1995, and above findings are only a snap shot. The results from 

the PFIPs will significantly extend the technical knowledge that will lead to better solar-arrays in the future, 

and are a great help in refining the environmental models and the understanding of damaging effects of 

impacting particles. 

References 

K. Adamson, L. Gerlach, P. Birot, 1994, Electrical Performance, Anomaly Investigation and 

Operational Implications for the EURECA Solar Generator, 

ESA WPP-069, EURECA Symposium, 26-29 Apr. 94, ESTEC, Noordwijk, The Netherlands 

2 L. Gerlach, E. Bongers, T. Mende, 1994, Post-Flight Investigation Programme for the Eureca Solar 

Generator and First Results 

ESA WPP-069, EURECA Symposium, 26-29 Apr. 94, ESTEC, Noordwijk, The Netherlands 

3 A. de Rooii, L. Gerlach, 1994, First Investigation Results from EURECA Solar-Array Samples Cut 

after Retrieval at NASA-KSC 

ESA WPP-069, EURECA Symposium, 26-29 Apr. 94, ESTEC, Noordwijk, The Netherlands 

4 Laurance R.J. 1990, The History of the Hubble Space Telescope and ESA's Involvement, 

ESA Bulletin No. 61, February 1990 

5 Henson B.W. 1990, ESA's First In-Orbit Replaceable Solar Array, 

ESA Bulletin No. 61, February 1990 

6 Gerlach L. 1990, The Solar-Power Generator for the Hubble Space Telescope, 

ESA Journal 199012, Vol. 14 

7 Henson B.W., Jackson B. , Gerlach L. 1991, In-Orbit Performance Overview of the 

HST Solar-Generator 

European Space Power Conference, Florence, Italy, 2-6 Sep. 1991, ESA SP-320 

8 M. v. Eesbeek et all, 1994, Post Flight Investigation of EURECA - Preliminary Findings and 

Recommendations 

ESA WPP-069, EURECA Symposium, 26-29 Apr. 94, ESTEC, Noordwijk, The Netherlands 

9 G. Drolshagen, 1994, EURECA Meteoroid/Debris Impact Analysis 

ESA WPP-069, EURECA Symposium, 26-29 Apr. 94, ESTEC, Noordwijk, The Netherlands 

283 

------------- --~- -----
SESSION V 

BLANKET AND ARRAY 

DEVELOPMENT 
-;? 9?~/¥ 

N95- 20532 

HIGH-PERFORMANCE, FLEXIBLE, DEPLOYABLE ARRAY DEVELOPMENT 

FOR SPACE APPLICATIONS 1 

Russell N. Gehling, Joseph H. Armstrong, and Mohan S. Misra 

Martin Marietta Astronautics 

Denver, Colorado 

SUMMARY 

Flexible, deployable arrays are an attractive alternative to conventional solar arrays for near-term and 

future space power applications, particularly due to their potential for high specific power and low 

stowage volume. Combined with low-cost flexible thin-film photovoltaics, these arrays have the potential 

to become an enabling or an enhancing technology for many missions. In order to expedite the acceptance of thin-film photovoltaics for space applications, however, parallel development of flexible photovoltaics and the corresponding deployable structure is essential. Many innovative technologies must be 

incorporated in these arrays to ensure a significant performance increase over conventional technologies. For example, innovative mechanisms which employ shape memory alloys for stowage latches, 

deployment mechanisms, and array positioning gimbals can be incorporated into flexible array design 

with significant improvement in the areas of cost, weight, and reliability. 

This paper discusses recent activities at Martin Marietta regarding the development of flexible, deployable solar array technology. Particular emphasis is placed on the novel use of shape memory alloys for 

lightweight deployment elements to improve the overall specific power of the array. Array performance 

projections with flexible thin-film copper-indium-diselenide (CIS) are presented, and Government-sponsored solar array programs recently initiated at Martin Marietta through NASA and Air Force Phillips 

Laboratory are discussed. 

INTRODUCTION 

It is evident that a strong trend towards smaller, lighter spacecraft launched on smaller launch vehicles 

has developed, and presently all spacecraft subsystem components are being scrutinized for improved 

performance. Along these lines, the power subsystem, and more specifically, the solar array and associated structure, has been receiving significant attention, particularly in terms of higher specific power and 

reduced stowed volume. The reason the solar array plays such a significant role in this activity is that 

solar array structures are often physically the largest subsystem, and any reduction of weight on an 

extended boom can have significant impact on spacecraft operations. To achieve overall specific power 

levels greater than 100 W/kg, and stowage volumes less than 0.10 m3 for arrays as large as 1 kW, a 

complete system approach must be taken. In most cases, aggressive goals such as these cannot be 

met with evolutionary technologies. Rather, revolutionary advancement of technologies must be accomplished to meet or exceed these goals. 

Both reduced array weight and higher photovoltaic conversion efficiency have an effect on spacecraft 

design in terms of increased payload size/weight. Dynamics of arrays in the kilowatt range can adversely 

affect spacecraft performance. Presently, PV arrays are nominally 20 W/kg to 40 W/kg, where most of 

the array weight is associated with the rigid substrate upon which the cells are mounted. As a result, a 

great deal of effort has been placed on improving efficiency to reduce array size, but low manufacturing 

yield and increased complexity drives device and installation cost significantly. Other efforts to improve 

specific power are related to lightweight structures, including fold-up and roll-up technologies. An additional advantage of these lightweight substrate approaches is the promise of smaller stowage volume. 

1 This Work Sponsored by Martin Marietta Independent Research and Development (IR&D) Project D17R and Other Internal Funding. 

287 

AGE N NOT FILMED 
l_ 

The following paragraphs introduce the incorporation of revolutionary technologies currently under·development at Martin Marietta in both thin-film flexible photovoltaics and in lightweight, reliable 

deploymenVstructural elements. 

Stowage and Deployment of Flexible Arrays - While use of composite materials in array structures to 

improve specific power for reduced weight is becoming more commonplace, they cannot easily accommodate small stowage volume, particularly for large arrays. One of the ideal configurations is a flexible 

array, although implementation of such technology has been limited due to the lack of a commerciallyavailable flexible solar cell design. Some flexible deployable arrays have been flown, most notably on 

the Hubble Space Telescope and the Solar Array Flight Experiment (SAFE) on the Space Shuttle. In both 

cases, the fragile nature of the solar cells (Si and GaAs) demanded significant attention to structural support. As a result, no array which has flown was truly flexible; at best, the array consisted of numerous 

array segments housing the rigid cells hinged together in flexible jOints. 

Due to the advent of flexible solar cells and modules, many flexible array concepts are now possible. For 

example, issues regarding the stowage and deployment of a rollup array similar to that used in the 

Hubble Telescope can be more easily addressed. Difficulty in rolling up rigid, fragile solar cells in a compact stowage volume is eliminated by completely flexible solar arrays and photovoltaics. Because the 

blanket is completely flexible, it is possible to integrate the deployment mechanism and the blanket for 

improved rigidity, thereby eliminating the possibility of buckling noted in the latest Hubble repair mission. 

Martin Marietta is investigating a variety of techniques to deploy flexible PV blankets. One promising 

concept utilizes shape memory alloy (SMA) elements to achieve both weight and performance advantages over other technologies, including self-rigidizing inflatable technologies. Furthermore, because the 

shape memory effect can be activated both electrically and by passive solar heating, structural deformations caused by unanticipated spacecraft maneuvers exceeding design parameters can be autonomously 

repaired in orbit. This key advantage over every deployment scheme can ensure array repair and continued high-level performance over a significant array lifetime. 

In addition to the blanket and deployment schemes, a key aspect in the overall specific power goals is 

the pointing subsystem. Martin Marietta IR&D funding has developed a shape memory gimbal mechanism capable of two-axis tracking with a weight savings of approximately 50% compared to conventional 

technology. 

Photovoltaics - An attractive alternative to conventional crystalline and amorphous silicon PV is polycrystalline thin-film devices. Polycrystalline thin-film PV, such as copper-indium-diselenide (CIS), offer an 

alternative to Si for most space power applications. These devices have the highest tolerance for radiation damage of any crystalline PV material in proton and electron environments because their short diffusion length (compared to conventional Si and GaAs) makes them less susceptible to damage caused by 

radiation. Furthermore, polycrystalline thin-film PV have proven themselves far more stable than singlejunction a-Si [1,2]. Polycrystalline thin-film devices, which nominally do not exceed 5-8 Ilm thickness 

(excluding substrate), have been reported with efficiencies as high as 15.5% (NREL - 1993) in Air Mass 

(AM) 1.5 insolation. Also in a joint Martin MariettaiNREL effort to supply a PV test article, an active area 

efficiency exceeding 13.5% in AMO was observed. The CIS Technology is discussed in an accompanying paper: "Advances in Polycrystalline Thin-Film Photovoltaics for Space Applications". 

LIGHTWEIGHT ARRAY TECHNOLOGIES 

Integral Deployment I Structural Elements - The achievement of high overall specific power places a premium on subsystem mass allocations. Current deployable solar arrays generally use motor driven mechanisms to extend rigid panels or deployable structural support elements (Astromasts, STEMs, telescoping booms, etc.). While structurally efficient, the mass of the deployment drives and mechanisms penalizes the overall solar array specific power. For non-retractable applications, inflatable deployment sys288 
tems are one approach to eliminating the motors and mechanisms, but must include a pressurization system and an envelope or collapsed structure that holds pressure during deployment. Inflatables generate 

reliability concerns from the standpoint of assuring that the integrity of the pressure envelope is not lost 

during final stowage prior to the mission. 

In order to achieve the most reliable, lowest mass system, Martin Marietta has developed a concept 

which combines deployment and structural elements of the solar array into an integral system. We are 

developing integral deployment I structural support elements composed either entirely of SMA, or SMAcomposite laminates. Deployment will be accomplished by one of two methods; 1) heating of the SMA 

member, or 2) using heated SMA to c~ntrol deployment of an elastic member. The integral deployment I 

structural elements, combined with a fully flexible array blanket will allow the attainment of array specific 

power levels greater than 100 W/kg. 

Design Using Shape Memory Alloys - Shape memory alloys such as Nitinol (a Nickel-Titanium intermetallic discovered by the Naval Ordinance Laboratory in 1963) provide attractive engineering properties for 

use in lightweight actuator designs and adaptive composite structures. SMAs undergo a reversible crystalline phase transformation that is the basis of the "shape memory effect". The low temperature phase is 

a twinned, martensitic structure which is capable of large strain deformation (in excess of 10% in some 

alloys) with relatively little stress (approx. 70 MPa). The high temperature phase is a cubic based, 

austenitic structure with mechanical behavior more similar to conventional metals. When the martensite 

is deformed, and then heated, the original heat-treated shape is recovered. However, if the deformed 

martensite is constrained during heating, high recovery stresses evolve (>690 MPa is possible in some 

alloys). A combination of the two effects allows SMAs to produce mechanical work with the application of 

heat. Using wire as an example, Figure 1 demonstrates the shape memory effect. 

Despite their attractive capabilities, the utility of SMAs in the past has been limited due to a lack of understanding of their very interdependent force-length-temperature response and associated non-linear and 

hysteretic behavior, as well as the effects of creep, fatigue, and material property drift which results from 

transformational cycling. These effects have been under study at Martin Marietta to provide the basis for 

effective alloy processing and ''training'' before incorporation in applications. Moreover, recent development 

of analytic modeling theory has made possible effective engineering of optimized mechanisms and devices 

based on experimentally derived parameters from property-stabilized SMA material. 

Shape memory alloys are an ideal solution to the deployment of a flexible lightweight solar array. 

Because the array has an extremely low mass and is deployed in a O-G environment, the load carrying 

demands placed on the shape memory deployment system will be very manageable. In addition, since 

the system requires one mission cycle (and 5 to 10 test cycles) SMA stability and non-linearity is not a 

major issue. However, several design challenges remain, including the use of the SMA material as structural elements in a martensitic state, and forming the required support geometries. As shown in Figure 2, 

the yield strength for a typical NiTi SMA is much lower for martensite than for austenite. As a result, 

SMAs are seldom used in their martensitic state as structural members. However, because lightweight 

arrays experience minimal on-orbit loads, using the SMA elements in their martensitic state is feasible. 

Benefits of reduced mechanism mass and associated interfacing hardware must be traded against the 

increased structural mass associated with the shape memory elements. Our studies indicate that an 

integral SMA structure provides the lowest mass deployment/support system. 

Structural and Deployment Concepts - Our deployment concept employs a novel expansive roll-out 

approach. In the stowed configuration the array blanket and support elements are rolled-up together in a 

bundle approximately 30 cm in diameter. Activation of the primary deployment elements causes the bundle to expand and un-roll. The fully flexible solar blanket (including flexible solar cells) allows a smooth 

deployment with very little resistance applied to the deployment elements. When the deployment elements fully straighten at their full activation, this stage of deployment is complete. 

289 
AUSTENITE 

I 

COOL 

t 

TWINNED MABTENSlTE 

I 

DEFORM 

• 

DEFORMED MARTENSITE 

I 

HEAT 

t 

AUSTENITE 

HEATTBEAT SHAPE 

INJIG 

(a) 

COOL AND REMOVE 

FROM JIG 

(b) 

DEFORM 

(c) 

HEAT UPTO 

RECOVER SHAPE 

(d) 

Demonstration of the shape memory effect -- First the wire is heat treated into the desired shape uSing a 

restraining jig (a) at an annealing temperature well above the alloy's transformation temperature. After the 

austenitic wire cools, it transforms into martensite with no shape change (b). The twinned martensite is easily 

deformed (c) generating the strain shown. When the spring is heated above its transformation temperature, 

the crystal structure reverts back to austenite, whereupon the original shape is recovered (d). Steps c and d 

may then be repeated. Because this transformation proceeds with the speed of sound, the application of heat 

to a shape memory element is usually the rate limiting step. 

Figure 1 Description of the Shape Memory Effect 

1400 

1200 

ro 1000 a.. 

6 

en 800 en 

~ 

U5 600 

400 

200 

0 

........ ~ 

Austenite /. - /~ .I~ 

f V 

1 L 

I .-/" ~ ... Martensite 

I I 

o 2 4 6 8 10 12 14 16 

Strain (%) 

Figure 2 NiTi Stress/Strain CUNe 

290 
The expansive roll-out deployment concept has several inherent advantages over other deployment 

approaches. High reliability can be achieved since there are no moving parts, drive mechanisms, or 

latches (except for launch restraint). Also, the absence of these components and associated interface 

hardware leads to the lowest mass deployment and structural support system achievable. Further, unlike 

systems using deployable masts or other elaborate deployable truss works, the low part count lends itself 

to rapid and efficient production. Finally, the concept allows manual restowage for repeated deployments 

using the actual mission hardware. This feature further improves the reliability of the system by allowing 

functional ground testing of all hardware. 

In order to obtain the greatest structural efficiency, yet maintain a single spacecraft attachment point, the 

array configuration should be approximately square rather than rectangular. However, to achieve reasonable maximum dimensions in the stowed configuration, a nearly square array must use a two stage 

deployment sequence. While a two-stage deployment adds some complexity and risk, use of SMA 

hinges for the secondary deployment makes the additional deployment step simple and reliable. The 

overall deployment sequence is shown in Figure 3. 

Stowed Launch Latch 

Relea.:se 80: Gimbdl 

Rotation 

Secondary Element 

Actuation 1 (utilize 

solar heating or 

hinges ir pnu:tical) 

full'l Deplo'led 

------............. 

Figure 3 Array Development Sequence 

291 

Primary Element 

Actuation 

Secondary Element 

Actuation 2 

- - - -- - -~
L_ 

A variety of primary support element geometries are under development and evaluation. These include 

the deployable element cross-sections shown in Figure 4. Initial analysis of potential cross-sections suggests that while the SMA tube configuration is most structurally efficient, it may be difficult to package 

repeatedly. Alternatively, while the arc cross-section is most easily fabricated and packaged, lateral / torsional buckling of arc cross section makes a suitable design difficult to achieve unless additional lateral 

support is provided. The array blanket interface could provide the necessary lateral support to make use 

of the arc cross-section viable. The two-part, joined cross section is also a candidate, but requires attention to the joint design and reliability, as well as requiring more assembly time. 

SMA Support Element Concepts 

SMA Arc 

SMA Overlapped Tube 

SMA Tube 

Joined SMA Arcs 

SMA Coaled 

Overlapped Tube 

Figure 4 Primary Support Member Concepts 

If deployed loading and natural frequency requirements are greater than about 0.2 G and 0.25 Hz 

respectively, an SMA coated STEM-type element must be implemented. In this concept, the deployment 

force is generated by elastic energy stored in the STEM. An SMA foil layer sputtered on the STEM 

serves to control the deployment. Without an SMA control layer, the STEM would spring out in a violent 

and unpredictable fashion, thereby disturbing the spacecraft and possibly damaging the array blanket. 

Prior to blanket release, the SMA would be heated to hold the STEM rolled-up. With gradual cooling, the 

STEM would overcome the force generated by the SMA coating and deploy in a smooth, predictable 

fashion. In order to evaluate the viability of this concept, the required SMA coating thickness and thermal 

control to avoid post deployment deflection due to solar heating must be determined. 

Secondary deployment can be accomplished using SMA activated flexural hinges. These hinges may be all 

SMA or a coated beryllium-copper flexure (Fig. 5). This approach enjoys the same advantages as those discussed for the primary deployment elements. A potential advantage of the SMA hinges is the capability to 

achieve passive deployment from solar heating. This would further simplify the system but requires a better 

knowledge of the thermal environment. 

SMA actuated deployment is accomplished by inducing a temperature change in the "as stowed" shape 

memory elements. Increasing temperature causes the element to experience a phase transformation 

292 
SMA HINGE 

STOWED 

Figure 5 SMA Flexure Hinge 

from a fully martensitic condition to an austenitic state generating the necessary deployment force and 

displacement. The change in temperature can be generated by passively exposing the array to solar 

radiation or using a control system which regulates current flow to a thin film heater placed in intimate 

thermal contact within the actuation element 

Lightweight Gimbal Technology - The burgeoning demand for lightweight spacecraft and solar array technology requires development of an ultra-lightweight two-axis gimbal drive for solar array positioning. In 

response to this requirement, Martin Marietta's Mechanical Research & Technology Group has developed a new design for a lightweight gimbal to meet small spacecraft solar array pointing requirements. 

The basic requirement for the SMA gimbal was to position a small rigid panel with a slow and smooth 

motion during a fixed orbit with a two-axis rotation of +/- 90· in elevation and +/- 175· in azimuth. Owing 

to small orbital loads, the torque requirement was determined to be 1 N-m; primarily due to the bearings 

and cable management system. In order to keep a 2:1 torque margin the cable management system had 

to be smooth and consistent. As always, weight was an important consideration and a 1 kg maximum 

was set as a design goal. Power and drive efficiency was also considered and limited to 3.0 watts average and 5.0 watts peak. The gimbal was required to maintain position and stiffness with the power off. 

The gimbal design that was developed to meet these requirements uses shape memory alloy springs in a 

unique way to obtain a smooth rotary positioning. Figure 6 shows a photograph of the gimbal engineering development unit. A proprietary SMA drive technique, used for both the azimuth and elevation drives, 

results in significant reduction of the control electronics associated with brushless motor drives. 

Another requirement that must be addressed is gimbal stiffness and holding torque. We have incorporated a brake into the design which passively releases when the SMA drive is activated, and then engages 

after power is removed. There are two brakes per axis and they act as a "no-backdrive" mechanism 

which stabilizes the gimbal position. 

During the evolution of this design, many methods to reduce structural weight were implemented. The 

gimbal housings are constructed using graphite polycyanate composite material combined with the high 

performance composite plastic, Torlon. These two materials are used together, and in some cases cocured, to produce a unique hybrid housing that is strong and very lightweight. This technique allows 

housing design without complex composite lay-ups or inefficient weight reduction machining. 

Flexible Blanket Design - The photovoltaic blanket subsystem for our solar array concepts will be fully 

flexible. Blanket layup consists of a laminate printed circuit, adhesive, and monolithically-integrated CIS 

293 

_J 
l_ 

:.' ;·).:: ..• ( .... .... ;, .....•. 0: .• 

• " • • .' :;. A ..,., .. , ~ .. ; •• '''~'' 

Figure 6 SMA Gimbal Engineering Development Unit 

modules which already include thin-film replacement for conventional coverglass. Because the large size 

of the modules (nominally 30 x 30 cm) allows for passing circuitry underneath while still maintaining electrical connections, it is possible to lay-up the printed circuit in a single layer. Electrical interconnects are 

made from the module to the interconnect pads via soldered flexible jumpers parallel to the rolling axis. 

The printed circuits laminated into the blanket are used to convey array power to a flexible bus connector 

at the edge of each blanket section 

ARRAY FUNCTIONAL AND ENVIRONMENTS TESTING 

Extensive functional and operational testing of flexible arrays is planned. Tests include power generation, deployment, thermal vacuum, and random vibration. The following paragraphs summarize key 

aspects of the planned testing. 

Power Generation - While array mass can easily be measured, solar simulation requires specialized 

equipment. Individual development modules up to 15 cm x 15 cm (6.0 in. x 6.0 in.) will be tested in a 

computer-controlled facility capable of testing these modules in AMO and AM1.5 insolation while monitoring temperature as the modules and array are heated/cooled. I-V and quantum efficiency characteristics 

will be measured by sweeping in both directions, as well as standard test methods developed for space 

arrays by NASA Lewis Research Center. Standard solar cells will also be measured to provide a reference signal. Modules up to 30.5 cm x 122 cm, as well as the complete array, can be tested in Martin 

Marietta's large space chamber/space simulator laboratory. Continuous solar spot size of 4.9 m (16 tt) 

diameter from 0.35 to 1.4 sun insolation can be realized. In addition, because of the uncertainty in time294 

------- ~ --

I 

I 

~ 
• depeodent phenomenon observed in some polycrystall ine materials, modules and array components will 

also be tested in the Large Area Pulsed Solar Simulator (LAPSS) which can accommodate 5.2 m (17 ft) 

spot diameter and the data will be used to develop a standard comparison between the test methods. 

Deployment Testing - A critical aspect of the solar array system is reliable deployment. In order to 

assure high reliability, deployment must be demonstrated using the actual hardware destined for the mission. To this end, we plan to design and conduct deployment tests of the assembled solar array. An 

advantage of our integral SMA structure is the capability to re-stow (manually) the array and repeat the 

deployment sequence multiple times. However, due to the highly flexible nature of our array concept, all 

aspects of the deployment cannot be tested simultaneously. We therefore plan to use a relatively simple 

gravity off-load suspension to test the primary and secondary deployment using separate test setups. 

This will require a test approach utilizing long (> 35 m), lightweight suspension wires supporting the array 

in two alternate test configurations; a primary deployment test, and a secondary deployment test. 

Thermal Vacuum Testing - Thermo-structural array response is of critical importance to successful operation in the space environment. The low mass and stiffness of the structure makes it susceptible to thermal distortion and buckling. Our planned testing includes subjecting flexible arrays and components (in 

deployed configuration) to anticipated low and high temperature extremes (as shown in Figure 7, typical 

extremes are -200°C to SO°C, although actual values will depend upon the optical characteristics of the 

modules) and measuring the thermo-structural response and power output. 

85 

~ 80 

0. 75 E 

CI) 

l- 70 

e( 

CJ) 65 

== ..J 

60 

0 

\.c:. ............. -

A • max albedo" .. tth IR on btlck 

B • min albedo .... rth IR 

C - ma.1MI'th IR on call_ 

.... .. B ---:... ... ~-..:.:..!:..:=-- __ ...:.. ____ _ 

10000 20000 30000 40000 

Altitude (km) 

-40 

~ -90 

0. 

E 

~ -140 

e( 

CJ) 

== ..J 

-190 

-240 

o 

\ 

\. 

, 

, 

" .... ..... ....... 

LWSA 

~:~ ~ E2~r::-}u 

---

o . max .. rth IR on cella 

E • min albedo" Mrth IR 

--- 10000 20000 30000 40000 

Altitude (km) 

Figure 7 Array Temperature Versus Altitude 

Launch Vibration Testing - Vibration and quasi-static loading during launch will produce the highest load 

levels experienced by the solar array in its stowed configuration. Typically, stresses due to random vibration response far exceed those caused by quasi-static accelerations. Therefore, we plan to conduct a 

random vibration test on the stowed solar array system to proto-flight levels using a Delta or Pegasus 

input spectrum. Analyses will be performed to predict peak load levels and verify that combined launch 

quaSi-static and random loads are enveloped. 

PLANNED ARRAY DEVELOPMENT 

Martin Marietta has recently been awarded two Government sponsored solar array programs; a Phase A 

INSTEP study (NASA) to define a thin-film, deployable array flight experiment, and an Air Force Phillips 

Lab contract to develop a prototype lightweight, deployable solar array. 

Our INSTEP program will focus on the thermo-structural and electrical performance of flexible thin-film 

solar arrays. The experiment, as proposed, is summarized in Figure S. 

The Phillips Lab / Martin Marietta LWSA program will combine the technologies discussed above to meet 

295 
L 

-- Transducers 

Deployment Motor 

Biaxial Strain Gages 

PV Blanket 

Blanket Support 

Figure 8 INSTEP Flexible Thin-Film Solar Array Experiment 

the needs of future Air Force spacecraft. We have developed an initial configuration for our concept that 

employs integral deployment/structural elements, an ultra-lightweight gimbal,. and a fully flexible blanket 

using flexible CIS solar cells. This configuration, shown in Figure 9, is sized for at least 750 W EOl 

power generation. The integral deployment/structural elements are sized for 0.1 G load in the deployed 

configuration. Use of the structural innovations previously described, together with the lightweight and 

flexible blanket, allow our concept to achieve a specific power of 160 W/kg EOl under nominal lEO 

operating conditions. 

Estimated total system mass for achieving 500 W, 750 W, and 1000 W power levels is listed in Table I for 

the lEO and GEO environments at both 28°C and 60°C operating temperatures. These mass estimates 

were derived from our 750 W baseline concept. Note that the EOl specific power output exceeds 150 

W/kg after 7 years in the lEO environment and can approach that level after 10 years in the GEO environment if operating temperatures can be kept relatively low. 

REFERENCES 

1. L. Mrig: "PV Reliability Issues," Photovoltaics Workshop, Denver, CO 1990. 

2. J.H. Armstrong, C.O. Pistole, M.S. Misra, V.K. Kapur, and B.M. Basol: "Flexible Copper-IndiumDiselenide Films and Devices for Space Applications," Proc. SPRAT XI , 1991 . 

296 
iStowed Volume = 0.12 m3 

Total Mass = 5.6 kg 

Power Output = 830 W (EOl in GEO@28 OC) 

0.3m ~ 

\~ lm~ 

Specific Power = 150 W/kg 

Deployed 

flexible CIS -----

Array Blanket 

~ 2-Axis Ultra-Lightweight Gimbal 

/ ___ SecondanJ Deploymentl 

Structural Elements 

Figure 9 LWSA Baseline Design Concept 

Table 1 EOL Specific Power Estimates for LEO and GEO Mission Profiles 

Low Earth Orbit (7 yrs, 1000 km) Geosynchronous Orbit (10 yrs, 36000 km) 

Power Temp Area Blanket Struc- Gimbal Total Specific Area Blanket Struc- Gimbal Total Specific 

ture Power ture Power (W) (0 C) (m2) (kg) (kg) (kg) (kg) (W/kg) (m2) (kg) (kg) (kg) (kg) (W/kg) 

500 28 4.0 1.6 0.5 1.0 3.1 160 4.6 1.8 0.7 1.0 3.5 140 

60 4.4 1.8 0.6 1.0 3.4 150 5.1 2.0 0.8 1.0 3.8 130 

750 28 6.0 2.2 1.2 1.0 4.4 170 6.8 2.6 1.5 1.0 5.1 150 ........... ~, ..... ,. ---. 60 6.7 2.3 1.4 1.0 4.7 160 7.5 2.8 1.8 1.0 5.6 130 

1000 28 8.1 3.0 2.1 1.0 6.1 160 9.1 3.4 2.6 1.0 7.0 140 

60 8.9 3.3 2.5 1.0 6.8 150 10.1 3.7 3.3 1.0 8.0 130 

297 
,

.::{ t} r.] I 

N95-20533 

STATIC STABILITY OF A THREE-DIMENSIONAL SPACE TRUSS 

John F. Shaker 

NASA Lewis Research Center 

Cleveland, Ohio 

Abstract - Space Station requirements for power have resulted in a need for photovoltaic solar arrays 

possessing large blanket surface area. However, due to the limited Shuttle payload volume solar array 

designers have been driven to a deployable concept that by nature is extremely flexible. The principal 

support for this array system is the Folding Articulating Square Truss Mast (FASTMast). In order to 

accommodate service loads the FASTMast is expected to exhibit nonlinear behavior which could possibly 

result in structural instability. Presented herein are the results of the Lewis Research Center test and 

analysis efforts performed in an effort to characterize the FASTMast structural behavior in terms of 

stability. Results include those obtained from recent nonlinear testing and analysis involving a 1/10 

segment of the FASTMast flight article. Implications of these results as they relate to expected behavior 

of the flight unit will also be discussed. 

INTRODUCTION 

In order to characterize complex structures with any degree of accuracy it is necessary to update finite 

element (FE) models using appropriate test data. In general the approach involves a two-fold process 

whereby a structural model is optimized utilizing FE modeling methods and response data obtained from 

characterization tests. Test data is used to either validate or modify original assumptions used when 

creating an idealized computational model. In many cases the finite element method (FEM) is limited in 

terms of representation of structural behavior that deviates from linear-elastic response regime. In 

addition to improving the accuracy of response predictions this approach also reduces the cost of 

structural characterization by minimizing the amount of required testing. Once a FE model has been test 

verified it can be exercised repeatedly until all desired response information has been obtained. Although 

computational costs can also be expensive they are generally much less than those associated with 

structural testing. Due the complex nature of the FASTMast structure and the limited resources 

available, this approach was taken to identify the load states at which the strucutre becomes unstable. 

The objective of this initial study was to create a FE model of the FASTMast structure and attempt to 

correlate response states to those obtained to from a static structural test involving various applied loads. 

Utilizing the test results as a guide, appropriate model parameters were updated and analyses repeated 

until agreement between test data and FE results was achieved. It is clear from the results that this 

methodology can be used to effectively treat the stability characterization of this structure. 

In order to deploy large flexible space structures it is necessary to develop support systems that are 

strong and lightweight. The most recent example of this aerospace design need is vividly evident in the 

Space Station solar array assembly. In order to accommodate both weight limitations and strength 

performance criteria, ABLE Engineering has developed the (FASTMast) support structure. The 

FASTMast is a space truss/mechanism hybrid that can provide system support while adhering to 

stringent packaging demands. However, due to its slender nature and anticipated loading, stability 

characterization is a critical part of the design process. Furthermore, the dire consequences surely to 

result from a catastrophic instability quickly provide the motivation for careful examination of this problem. 

Shown in figure (1) is the solar array assembly of the Space Station Freedom. Once fully deployed the 

FASTMast strucutre will provide structural support for the solar array system. A unique feature of this 

299 

N OT F LME ________ . ____ ___ J 
structure is that the system responds linearly within a certain range of operating loads and nonlinearly 

when that range is exceeded. However, this study involves the nonlinear large displacement problem 

only since it yields the lowest load levels leading to a state of instability. A complete examination of the 

FASTMast stability problem is given in reference (1). 

HARDWARE DESCRIPTION 

Due to electric power demands of the Space Station user community it was necessary to provide solar 

array assemblies much larger than normally used for space flight. Limited by the modest payload volume 

of the Space Shuttle, designers were immediately driven to a deployable concept that could 

accommodate packaging and weight constraints. In order to support the solar array blanket it was 

necessary to design a support structure that possessed the strength characteristics of a space truss with 

the mechanistic features of a deployable structure. The answer to this problem was provided by ABLE 

Engineering with the FASTMast deployable mast assembly. 

The FASTMast structure is comprised of thirty-two interconnected bays of mast. A complete flight unit 

will stand approximately 104 feet in length and supports two solar arrays which are a total of 40 feet in 

width. The total weight of this structure including the mast canister and blanket boxes required for launch 

support is 2500 pounds. It is designed to provide 18KW of electrical power to the Space Station user 

community. Shown in figure (2) is a detailed description of the primary components that make up a 

single bay of mast structure. The longhorns are the primary axial and moment load carrying elements of 

the structure. In order to provide additional buckling resistance the longerons in the lower twenty bays 

possess a tapered cross-section. The engineering properties of all major structural components are 

given in tables (1) and (2). 

The principal elements providing resistance to shear and torsional loading are the prestrained stainless 

steel 7X7 wire rope diagonals. In addition to the shear resistance provided by this structural member it 

also provides flexibility that is required of a deployable structure. The current deSign preload level in this 

element is 31 pounds. Supplying the load required for diagonal prestrain is the fiberglass flex batten. In 

order to create the necessary force the flex batten is installed in a post-buckled state. The buckled shape 

of this element is clearly visible from the top view of the mast given in figure (3). A direct analogy to this 

design concept is the energy transmitted to an string from a buckled bow. 

In addition to reacting the preload of the flex batten, the elbow joints provide a pivot point required for 

mast stowage and deployment. Therefore, it was necessary to deSign this joint with a dual-function end 

fixity. In order to facilitate the folding action of the mast the elbow joints act as a hinge in the direction of 

rotation required for stowage. Shown in figure (4) is the manner in which the elbow joint, flex batten, and 

diagonal elements are connected to the mast. Figure (5) shows the diagonal to lange ron connection 

which is made with a bracket and two 4-40 socket-head cap screws. Also given in this drawing is a clear 

view of the folding direction of the longeron/flex batten interface. The pinned and hinged boundary 

conditions at this joint are associated with the Euler buckling and large displacement failure modes 

respectively. The pinned condition exists at this interface when hinge action is not taking place. 

At the top and bottom of each bay of FASTMast are the rigid battens and the corner fittings. These 

structural elements provide a pivot point for the longeron at the top of the bay and anchor the rigid 

battens to the space lattice. Rigid battens provide shear and torsion load resistance by restraining corner 

fitting motion. The taper feature of the rigid batten was incorporated in order to reduce weight and 

increase strength of the element. The diagonals are mounted at the top of the bay in a manner similar to 

that described above. 

300 

I 

-J 
rAlthough the individual elements of the FASTMast structure do not possess large strength capability the 

integrated unit appears capable of withstanding service loads. However, in order to achieve the required 

strength to weight ratio this type of structure presents an obvious stability problem. Further complicating 

this problem is the fact that there exists both a local and global instability modes each influenced by 

deformed mast geometries and applied loading conditions. 

NONLINEAR FAILURE MODE IDENTIFICATION 

The allowable load of a structure is a function of its design and the anticipated failure mode. 

Determination of allowable load requires a clear understanding of structural behavior during loading 

events and identifying the appropriate mode of failure. Once these two goals have been achieved a valid 

analytical model can be constructed and the allowable load of the structure can be determined. 

The FASTMast will be subjected to a combined state of moment (M), axial (A), shear (V), and torsion (T) 

load as graphically depicted in figure (6). From a simple static assessment of the mast it is clear that 

applied shear torsion loads are reacted in essentially the same manner. This figure merely attempts to 

present a simplified representation of all possible applied loads at a system level. Nonlinear behavior of 

the FAST Mast structure is due principally to the changing stiffness state of the structure that results from 

slack diagonals. As was previously stated shear and torsion loading is reacted internally by the diagonals 

and battens. An attempt to illustrate mast reaction to shear load is given in figures (7) and (8). Shown in 

this figure is the action of the flex batten and diagonals due to the action of preload P and shear load V. 

The shear and torsional stiffness of the mast is a result of the post-buckled flex batten force P inducing a 

tensile preload in the wire diagonals. The load state in figure (7) is a result of only preload P while the 

manner in which the mast will react shear load V is illustrated in figure (8). The sum of figures (7) and (8) 

represents the combined action of V and P. It is assumed that each diagonal is prestrained to the same 

level while resisting shear load equal/y. The flex batten reaction to the shear load is zero because it is in 

a post-buckled state and cannot take additional load. Although figure (8) indicates that a set of diagonals 

would be in "compression", physically this equates to a reduction of the force P supplied by the flex 

batten. The limiting state is reached when the load in the "compressed" diagonal becomes zero (slack 

condition), at which time the flex batten begins to pick up additional compressive load and the mast 

begins to move into a fold-up mode required for mast retraction (figure 9). The applied shear level at 

which unwanted mast stowage occurs is that required to overcome the preload in the wire diagonals. A 

shear load of this magnitude is much lower than that required for material yield in either the diagonal or 

flex batten. Therefore the principal failure mode for applied loads involving shear and torsion is structural 

instability. 

Based upon the preliminary failure mode assessment above it is clear that an evaluation of FASTMast 

stability should be undertaken. From initial instability considerations it was determined that the deformed 

state during instability would be either a local or global deformation as given in figure (10). A global or 

system failure event would correspond to the fold-up mode that results from excessive shear/torsion and 

axial load . Fold-up in this instance is defined to be mast action which occurs during a solar array 

restowing event. The instability event involving mast fold-up is mechanistic in nature and is due to the 

nature of the FASTMast design which involves both truss and mechanism elements. Furthermore, due to 

mast stiffness changes which occur during large deflections of the elbow joint, this form of instability is 

nonlinear in nature. On the other hand, the local failure mode involves the classic general instability of a 

pinned-pinned column subjected to an axial load which would occur during excessive moment and axial 

loading on the mast. The local instability event involves Euler buckling of a single longeron as is shown in 

figures (10) and (11) and is a linear response event. The principal difference between these two failure 

modes is that the system event is kinematic in nature. Also, the local event does not involve nonlinear 

behavior indicative of the system type failure mode. 

301 
l_ 

APPROACH 

After considering both problem physics and resource limitations it was decided that a one-tenth segment 

of FASTMast structure could be used to meet study objectives. This one-tenth model of the flight 

hardware is referred to as the 3-bay FASTMast unit. Furthermore, due to the symmetric nature of the 

structural load paths extending the test verified 3-bay theory to the 32-bay flight unit configuration was 

considered plausible. Based upon these initial assumptions parallel test and analysis efforts involving 

FASTMast nonlinear stability assessments were design and executed. 

THREE-DIMENSIONAL NONLINEAR LARGE DISPLACEMENT ANALYSIS 

The source of nonlinearity in this case is due to the changing system stiffness as the structure 

experiences large deflections. This particular large displacement problem includes the effects of large 

translations and rotations while strains remain small. However, in order to accurately predict nonlinear 

behavior it is extremely important the FE model closely match the nonlinear elements of the hardware. 

Therefore, accurate modeling of the flex battens, wire diagonals, and mast joints was deemed critical for 

generation of accurate and valid data. 

The model used for this analysis was created with the ANSYS finite element code. ANSYS was selected 

due to its proven nonlinear capabilities. The entire model consisted of 312 elements, 225 nodes, and an 

estimated 1330 degrees of freedom. Shown in figure (12) is the undeformed preloaded ANSYS model. 

In order to define the entire model five element types and ten real constant sets were required. All mast 

structural properties were based on the most-up-to-date information available. 

The longerons were modeled using three-dimensional tapered beams thus reflecting the design of the 

lower twenty bays of the flight unit. Each rigid batten frame required three-dimensional truss elements to 

model the batten tube and three-dimensional tapered beams to model the corner fittings. The wire 

diagonals were represented by two-dimensional tension only spar elements which included an initial 

prestrain equal to that prescribed by the required tensile preload. Finally, the fiberglass flex battens were 

also represented by three-dimensional beam elements. 

In order to represent joint flexibilities at the elbow and corner fittings hinge elements were introduced to 

these parts of the structure. The ANSYS hinge jOint provides translational and rotational stiffness in all 

six degrees of freedom at the point of application. At each elbow joint there are four hinge elements and 

at each corner fitting there are two. A hinge element consists of a coincident node pair that are 

connected in all but one rotational degree of freedom. Shown in figure (13) is an example of an elbow 

joint modeled using hinge elements. In this example the coincident node pairs are (5,45), (5,35), (5,25), 

and (5,15). At each longeron end there is a hinge that allows for ninety-degree rotation plus a 0.6 degree 

back rotation required to model stopping action of the deploying mast. After engaging the stop the hinge 

is no longer free to rotate and instead behaves as a torsional spring with a rotational stiffness of 1 x1 D· ,nIb/rad. The hinges on flex batten ends do not have rotational limits and possess very high translational 

stitfnesses. An identical connection process is carried out at the corner fittings without the inclusion of 

the flex batten. 

The ability to identify instability points during a geometric nonlinear large displacement analysis is not 

straightforward. Unlike linear static analyses, the structure is loaded incrementally and the equations of 

motion are solved in a piecewise linear manner over subintervals of the response regime. Over each 

loading increment the equations of equilibrium are solved iteratively until the solution converges within 

some specified tolerance ban. During this analYSis both force and moment convergence criteria were 

302 

--- ---------

) 

I 

l 

I 
--- --- --- ------ 

-

used to evaluate the adequacy of the solution before proceeding to the next load step_ Solution iterations 

continue until either convergence is achieved or the analysis is terminated by user request Upon 

successfully satisfying the equations of equilibrium the stiffness matrix is updated and the analysis moves 

to the next interval of applied load_ This process is repeated over the entire range of applied load. The 

fact that the solution is derived in a piecewise manner introduces the first analytical difficulty which 

involves bypassing the point of instability due to an interval selection that is too broad. This situation can 

occur for instabilities such as "snap through" buckling. Furthermore, even if the proper interval has been 

identified there still remains the question of at which QQin1 in the interval instability will actually occur. For 

example, if an instability occurred between ten and twenty pounds the failure load level possibilities 

include eleven, twelve, ... , and twenty pounds. The only explicit conclusion presented by code output is 

that a failure occurred somewhere between the end points of the applied load interval. Also, 

identification of instability pOints from output data can be accomplished by identifying radical changes in 

output values at some predefined characteristic point. In this case large displacements at the top of the 

mast due to a load of smaller or equal magnitude than the preceding step was taken to infer a point of 

instability. An analytical consequence of such an event may result in a set of equilibrium equations that 

prove to be nonconvergent. This type of nonconvergence is due to the fact that the structure has 

undergone a shape change that will not satisfy conditions of equilibrium. However, convergence 

problems may also result from modeling errors that have no relationship to a possible state of instability. 

Therefore, prior to reporting large displacements and nonconvergent behavior as indications of instability 

points structural response must be judged appropriate. The analyst must ensure that the FE model is 

accurately representing nonlinear behavior and results reflect problem physics. In general this is 

accomplished by utilizing engineering insight and structural test data during the data reduction process. 

The focus for this particular model correlation activity resided with the hinge elements. At each node 

involving a hinge element there are six degrees of freedom five of which must be assigned stiffness 

values. The sixth degree of freedom is that of the primary hinge rotational direction which for this study 

was taken to be frictionless_ Of the remaining five degrees of freedom only the lateral and axial 

stiffnesses were updated with the aid of results from stiffness testing. Stiffness values were obtained by 

performing a longeron axial load test and a system level lateral load test. The model hinge stiffness 

parameters were then updated and a nonlinear analYSis of the configuration given in figure (14) was 

performed and compared to test results. The remaining translational and rotational hinge stiffnesses 

were assumed to be either zero or very large and that fact is reflected in the FE model. 

STIFFNESS TESTING 

In order to generate the required stiffness values an axial stiffness of the longeron/elbow joint was 

performed in addition to a tip shear stiffness test. The test configurations are shown in figures (16) and 

(17). Stiffness testing was performed by agents of the Engineering Directorate of Lewis Research 

Center (LeRC) and details are available in references (2) and (3). 

The first model update was performed using stiffness data from an axial stiffness test which was 

configured as shown in figure (17). The test specimen consisted of an elbow fitting and two square 

longerons and was loaded in a series of ten load steps to final level of 4200 Ibs. As the load level 

reached 3000 Ibs nominal yielding of the strucutre had occurred. Yielding of the specimen increased 

dramatically over the range of 3000 to 4500 Ibs at which point the test was terminated. Linear behavior 

was observed throughout the 2400 Ib load case with yield occurring at approximately 2500 Ibs. A linear 

fit of the test data resulted in the load/displacement curve with a slope of 4.3 x 105 Iblin which represents 

the axial stiffness of the hinge. 

Updating of the lateral elbow joint stiffness was made using results from the 3-bay lateral loads test. The 

system level lateral load test was performed in a manner that allowed for sequential loading of the truss 

303 
strucutre until a limit value was achieved. A lateral load of 120 Ibs was applied to the structure in a 

stepped fashion and the displacements were recorded at locations given in figure (15). The nonlinear 

behavior of the FAST Mast is clearly evident in figure (18) which gives the deflection of the mast top as a 

function of lateral load. The system response becomes nonlinear as the applied load reaches a level of 

83 pounds. This bilinear behavior is due to the fact that the diagonal tension has been reduced to zero 

and the flex battens begin to react the applied load. The mast stiffness up to the 83 Ib inflection point is 

approximately 332 Ibs/in and decreases to 28 Ibslin at the onset of nonlinear behavior. In order to 

achieve model correlation an attempt was made to match analytically the test results given in figure (18) 

by adjusting the lateral stiffness of the elbow joint. 

RESULTS AND DISCUSSION 

The results of the structural testing and nonlinear large displacement analyses are given graphically in 

figure (19). In this figure the lateral displacement of the 3-bay FASTMast is plotted as a function of lateral 

load. The top curve represents the system response identified during the 3-bay lateral load test and the 

remaining curves show results from the nonlinear analyses. A total of three FE model updates were 

made in support of this study. For each analysis the lateral load was equal to 120 Ibs applied in twelve 

equal load steps of 10 Ibs. In all three cases there was no evidence of a structural instability. As a 

starting point all hinge stiffnesseses were modeled as rigid connections. As expected this resulted in a 

mast stiffness much higher than that observed in test. System performance with rigid hinge connections 

is shown in the lower-most in curve of figure (19). The results.· indicate a system stiffness of 796 Ib/in in 

the linear regime, and 34 Iblin during nonlinear response. The first update of the model consisted of 

modifying the axial stiffness of the hinge jOint to a value 4.3x1 05 Ib/in. This change resulted in a linear 

system stiffness of 667 Iblin, and a nonlinear stiffness of 33 Ib/in. Finally, the lateral stiffness of the hinge 

elements was updated to a value of 400 Ib/in. This value was determined from a recent hinge jOint test of 

the 3-bay unit. From figure (19) it is clear that this change results in a dramatic increase in system 

flexibility. The slope of the curve depicting these results indicates a linear stiffness of 515 Ib/in and 

nonlinear stiffness of 24 Ib/in. Therefore, the final updated model poss~ss a 54% error in linear stiffness 

and -14% elTor in nonlinear stiffness. For this study a negative error rate implies excessive flexibility 

exists in the model as compared to test results. Although model results improved over the updating 

process more work will be required to obtain a better model correlation. 

A second interesting result shown in figure (19) is that the onset of nonlinear behavior given by analysis 

differs from that found during testing. The test results indicate that nonlinear behavior begins at an 

applied lateral load level of 84 Ibs while the analytical level is approximately 100 Ibs. This discrepancy 

could be the result of several possible test and/or model anomalies which may include: (1) variation in 

anticipated test article diagonal preload, (2) model hinge stiffness inaccuracies, (3) excessively stiff model 

of the rigid batten, and (4) inaccurate value of Young's modulus for the flex batten. Each of these items 

will be examined during the remainder of the model correlation effort in order to identify the source of this 

discrepancy. 

In terms of structural performance it is clear that resistance to applied loads will decrease dramatically 

when the FASTMast enters a nonlinear response regime. Furthermore, if the predominant failure mode 

is instability collapse will be sudden and catastrophic. Therefore, it would prudent to restrict mast 

operations to the linear regime. 

304 
SUMMARY 

The results presented herein indicate that FE model updating techniques can be successively employed 

when analyzing large flexible structures. However, reliable updates to complex elements such as rotating 

hinges and preloaded elements can only be achieved utilizing test data and sound updating procedures. 

Once the FE model has been test verified, characterization of system responses can be achieved with 

greater levels of accuracy and validity. 

REFERENCES 

1. Shaker, John F., Static Stability of a Three-Dimensional Space Truss, M.S. Project,Case 

Western Reserve University, 1994. 

2. 3-Say FASTMast Shear Test and Analysis, Lewis Research Center Engineering Report 

No. SAS 93-003, April 1993. 

3. PV Structural Analysis Space Station Tension Test of FASTMast Elbow Joint, Lewis Research 

Center Engineering Report No. 93-3. 

305 
r 460~ --- Top blanket boxes 

In~~ j-.--

Extendable 

mast 

(fastmast) ---'" 

Bottom 

blanket 

boxes - .... .... 

/' 

/ 

__ Blanket 

1290 

In. 1400 

In. 

/~'~>~,--------------------~-- CanlSter.-/ / , 

Beta joint --' 

/ '- Base of mast 

Figure 1 - Space Station Solar Array 

r 

--

1\ 

I \ 

Figure 3 - FASTMast (top view) 

306 

• ~N;e.rcru. 

ProvlCe stJf':nes,s 

Arrchcr non-n.x.in9 

dJ.o1COnaJs --

P,..load dlaCCtUlb - _ _ _ _ _ _ _ 

·';nibl,battens --------1 Provide shear st:ren~.l'1 

·ebowfltt!~ 

Provide <mcarons pivots =§=======5:~ P""vfda tlex batten pivo~ / 

Aroc~or f1axirq d!"~naJs ..../ 

• CIaCONis 

Provide shear ""d / 

/ 

torsional stiffn.ss / c.:::.-----------.-..:::::J> PosiHon lontWeron / I 

_bow joint; In / I 

deployed coneition ..../ / 

Side view. one tull bay 

I 

• Comar ftttlngs I 

1'TovK1. ~ron pivots / 

Arocha f1axinQ V>d I 

non-llexinq d;"~naJs - ' 

Figure 2 - FASTMast Structure 

I 

i 

U 

\.CN6~JJ ~\ 

ELCO\J.J FITT1:N G 

Figure 4 - Flex Batten / Longeron Interface 
.

p 

UF!>E1.. '""'" !>c~C>-.l 

(~"'~~C» 

.i 

.! 

1----:--- L=..a f..cl..:S~,..;. 

(DE?~'{E~) 

Figure 5 - Elbow Fitting Longeron Interface 

o-S""P 

- -

1 i 

I __ I 

c..:."*31' I ;)-:.::-:- I 

I 

r i 

I L.: ~ I ' 

1 o .. !3P lo..:Ep 

T 

"- '!-_::::x:l '? ..-

.-~ So /I ~ I 

I 

Figure 7 - FASTMast Preload Reactions 

307 

---.--------

.~~ . H 

Figure 6 - FASTMast System Loading 

~ 

V 

c~ v 

1....0'/ 

0 ;-

;-

Figure 8 - FASTMast Shear Load Reaction 
v----~~--------------~ 

Figure 9 - FASTMast Fold-up Action 

, "'-, 1_--

C 

\.... 

p 

I 

I 

Figure 10 - FASTMast Failure Modes 

----l:ul.i:.R FAI~L.~F... MODE 

Figure 11 - Longeron Motion During Failure Modes 

308 
r 

I 

I 

Longcron 

Figure 12 - FASTMast Nonlinear Finite Element Model 

Rgure 13 - FASTMast Hinge Joint 

--V 

Figure 14 - Nonlinear Finite Element Model Tip Shear Case Figure 15 - Displacement Transducer Locations for Tip Shear Test 

309 
I 

I 

CONCRETE~ 

BEAM 

y 

LG ';D C=::'.L A ,. .--==> DI ,.-::- r r.or.r' I _ . . . :... ."' I -

, ~ (rTO? ?LAT=: 

~ 

-,-

:qUM T. TIGN 

P 

t 

t 

p 

lC>--.E...;(.

~ =M 

Figure 16 - 3-Bay Tip Shear Test Configuration Figure 17 - Longeron Axial Stiffness Test Configuration 

2 

/ 

1.5 

[ 

~ Ii 

~~ 

~ V 

---:---

III 

j 

"-

~ 

0 

3 0.5 

"-

"-0 

!-o 

0 II 

-0.5 

() 2() 40 6() 80 100 120 140 

APPLIED LOAD (LBS) 

Figure 18 - FASTMastT op Deflection VS. Lateral Load from LeRC Shear Test 

310 
I 

I " 

I 

2 ~------------------------------------------------

,,--. 

Z1 

----- Z 

o 

fU 

W 

.-l 

LL 

W 

o 

0.... 

fo 

-0- TEST DATA 

-{]- RIGID ELBOW JOINT MODEL 

--!;:: - ELBOW JOINT AXIAL STIFFNESS 

-V- ELBOW JOINT AXIAL AND LATERAL STIFFNESS 

, • I I I. I I ' • , , I ' • I 

20 40 60 80 100 120 140 

LATERAL LOAD APPLI ED AT TOP OF MAST (LBS) 

Figure 19 - Test and Analysis Mast Top Deflection vs. Applied Load 

311 

- - -------- -----------_.----------
Mas; I MJteriJ..! I L!ngtil I .).,,;<:~ I Young':) I ~lovem~n[ of i Mo"em~~e of 

kme~( ~lc~u!us I ln~:-j, [.,' lr.e:-: ~ L,: 

I I 1:-' I 1:' .. - I L3il:-;: I L" I 1:';' 

Tap"~ed I Aluminum I 19.750CC1l I 0.3.:3Irxl I 10x!0' I 0.010[00 I O.OIOleD 

Lon2uon 6061-T6 I I 

Straighe I Aluminum I 19.75COCQ I 0.15GCW I 10x10· I 0.C05210 I 0.C{)5210 

Lon2~~o n 6061-T6 I 

Rigid I Aluminum I 19.3 C{)(xO I 0.063900' I 10x10' I 0.0 l7.!GO ' I 0.C{;()'!36 , 

Bane:! 6061-T6 I 0.0361CO' I O.COI ' '0' 0.C003 :6: 

Bane:! I AJuminum I II.1COOOO I 0.051110 I IOxW; I O.CO 1390 I O.CQ 1390 

Tu~ 6061-T6 I 

Flex I FibergJass I 3l.i50000 I 0.103000 I 

8x [O'; I 0.0011[0 I 0.COO59 I 

Batte:! 

Diagonal I Seainless I 36.000000 I 0.002313 I 19xtO· I --- I ---

Seeel 

1. Eleme:H y-a:t.is one:lted suc~ tbJ[ witb ~-<l.'tis !OO~ le~g"Jl of e!eme~t rigoe-h:?r.d .-ule is 5ms;:d 

1. Elemene z:-a.'us vertic:ll [0 coss-section 

3. Wide eod of tapered section 

4. Narrow eod of capered sectioo 

Table 1 - Engineering Properties of Principal FASTMast Components 

Scruc:urol 

I 

y[ate:iaJ I Young's i Poisson 's , 

I .-\rea )[or.:e~t or C0(!~C:':!1t 

Ele:nent I Y[oCuius I R~tio , In:!:-i':!' or 1.1e::n:li I i .~ OilIlSion 

i LB/IN" , ~' I?-,- ) I r-iir-i· r 

Lon!!eron I .~J-6061 i 10XtO· I 0.33000 0.250<Xl 0.0052 ! I 0.1) 

P.e); Batten I Ficwdas.i i 3X IO· I 0.3'3000 , O.lO::OO GOlc.~ l.OX :0" 

Diagon:ll I StainJess I 30XIO· ! O.330c0 ! 0.C02}i , '::-ii.A. IJ.:J 

Seeel I 

Table 2 - Properties for Two-Dimensional Nonlinear Planar Truss Study 

312 

. 

___ J 
I 

I , 

I 

I 

I 

3 {J~'/ Y'I P 

N95- 20534 

AN UPDATE ON THE DEVELOPMENT OF A LINE-FOCUS REFRACTIVE 

CONCENTRATOR ARRAY 

Michael F. Piszczor 

NASA Lewis Research Center 

Cleveland, Ohio 

Mark J. O'Neill 

Entech, Inc. 

DFW Airport, Texas 

and 

Lewis M. Fraas 

JX Crystals 

Issaquah, Washington 

ABSTRACT 

Concentrator arrays offer a number of generic benefits for space (i.e. high array efficiency, protection from 

space radiation effects, minimized plasma interactions, etc.). The line-focus refractive concentrator concept, 

however, also offers two very important advantages: (1) relaxation of precise array tracking requirements to 

only a single axis and (2) low-cost mass production of the lens material. The linear refractive concentrator can 

be designed to provide an essentially flat response over a wide range of longitudinal errors for satellites having 

only single-axis tracking capability. New panel designs emphasize light weight, high stiffness, stowability and 

ease of manufacturing & assembly. This paper will address the current status of the concentrator program with 

speCial emphasis on the design implications, and flexibility, of using a linear refractive concentrator lens as well 

as detail the recent fabrication of prototype hardware. 

INTRODUCTION 

During the past two years, NASA Lewis, ENTECH and JX Crystals have been working on a refractive line-focus 

PV concentrator array for space applications (ref. 1). The linear-focus concentrator array for space is directly 

comparable to ENTECH's current line of terrestrial PV concentrator systems, and benefits from the operational 

and manufacturing knowledge gained by years of experience in this field. This concentrator concept is based 

on the same general principles as the point-focus mini-dome Fresnel lens concentrator (ref. 2,3) and offers 

many of the same advantages of the point-focus system. In addition to the generic benefits offered by PV 

concentrator systems for space (Le. high array efficiency, inherent protection from space radiation effects, 

minimized interactions with the space plasma, etc.), the linear concept offers two very important advantages: 

(1 ) relaxation of precise array tracking requirements to only a single axis and (2) low-cost mass production of 

the concentrator lens material. 

While the program is still pushing to maintain high array performance (Le. high efficiency, low weight & volume, 

etc.) as its primary goal, greater emphasis is being put on manufacturability and overall array and spacecraft 

system cost reductions . Since the last report on this technology, low-cost, roll-to-roll production of the 

concentrator lens material has been demonstrated. New array structure designs are also being developed. 

The new array designs emphasize light weight, high stiffness, low thermal distortions and ease of 

manufacturing & assembly. 

313 
LINEAR CONCENTRATOR LENS DESIGN 

As mentioned previously, the linear concentrator is based upon the same general lens design as the minidome Fresnel lens concentrator. The lens, currently made from silicone, has a curved, smooth outer surface 

with individually tailored linear Fresnel facets along the inner surface. (Refer to Fig. 1). The curvature of the 

outer surface and the Fresnel facets are designed such that the angles at which the light enters and exists the 

lens are equal, producing a condition that maximizes optical efficiency while minimizing the effects of radial 

shape errors. Thus, the lens is very tolerant to shape errors due to manufacturing, assembly or operational 

thermal distortions. The lens can also be easily designed to adjust the desired flux profile on the cell. (A 

double-hump design is currently used). This design flexibility has minimal impact to the overall design of the 

panel and array structure. 

One of the major concerns facing photovoltaic concentrator systems is the precision of sun-pointing required. 

A key advantage of the linear refractive concentrator is the ability to modify the lens/cell design to obtain the 

necessary off-track design profile required for a specific mission without significantly affecting the basic array 

design. Because of the linear nature of the lens design, pointing errors along the longitudinal (length-wise) 

axis of the lens are much more tolerant to off-tracking than errors along the lateral (critical) axis. Errors along 

the critical axis are on the order of a few degrees and are highly dependent upon the lens/cell configuration, 

concentration ratio and the use of optical secondaries. 

Fig. 2 demonstrates what happens when a linear lens is off-pointed along the longitudinal axis for the current 

lens design (40 degree rim angle). In Fig. 2a, light rays from the sun enter the curved lens and are then 

focused on to the cell within a specific flux profile. As the array begins to off-track along the longitudinal axis 

the focal length begins to effectively "shorten" (in a two-dimensional visualization). As this happens, some of 

the outer rays begin to move off the cell. Fig. 2b shows what happens for a 10 degree longitudinal error. At 

this angle the focal "shortening" is small and most of the light still hits the cell. The effective loss on lens 

optical efficiency is minimal. As the off-tracking error is increased (Figs. 2c. & 2d.), the effective "shortening" 

becomes more pronounced and the lens optical efficiency begins to decrease. 

To illustrate the flexibility of the linear refractive concentrator lens, the effects of both lateral and longitudinal 

errors are plotted in Fig. 3 for two different lens designs. Lens optical efficiency refers to the amount of 

sunlight that actually falls on the photovoltaic cell and includes all reflection losses associated with the lens. 

Please note that these plots are for two specific designs and that optical secondaries were not included. (The 

use of optical secondaries will significantly increase tolerance along the lateral (critical) axis and cell 

concentration ratio along the longitudinal axis). The purpose of Fig. 3 is to qualitatively illustrate the variability 

of off-tracking tolerance as a function of the combined lens/cell design. 

Fig. 3a shows a 3-dimensional plot of lens efficiency as a function of tracking error for the current lens design 

(40 degree rim angle). Note the much greater degree of tolerance along the longitudinal axis for a line-focus 

design. Still, the decrease in lens efficiency is fairly symmetric about the "on sun" position. Fig. 3b shows the 

same plot for a lens/cell design that is "single-axis tracking" (Le. can accommodate errors of ± 23.5 degrees 

along the longitudinal axis without significant power loss). Note the difference in the shape of the plots. Also 

note that, unlike certain point-focus concentrator systems, there is still a significant amount of power available 

as the array begins to move off the sun. The "single-axis tracking" deSign could be used to provide an 

essentially flat response over a wide range of longitudinal errors for those satellites having only single-axis 

tracking capability. However, the cost of this added off-tracking capability is usually a decrease in 

concentration ratio (Le. increased cell area). Thus, a true optimization of the lens/cell design is dependent 

upon a number of array, system and mission level factors (Le. cell cost, array cost, radiation damage, single axis 

vs. double-axis tracking, contingency requirements, etc.). The plots in Fig. 3 represent the start of a detailed 

analysis to evaluate and quantify these relationships. 

Fig. 4 illustrates how the design calculations compare to actual measured performance. Fig. 4 shows the 

calculated curve of geometric concentration ratio as a function of sun-pointing error tolerance along the lateral 

axis. As noted previously, to effectively improve the sun-pointing tolerance of the array, the geometric 

314 
concentration ratio on the cell must decrease (Le. cell area must increase). Also plotted on the chart are 

measured data pOints from both the current linear lens design for space and one of ENTECH's production 

level terrestrial concentrators. (The range on the measured values correspond to optical lens efficiencies in 

the range of 90 to 95%.) Note that the measured pOints follow the predicted values quite well, being slightly 

above the predicted curve in all cases. 

PROTOTYPE HARDWARE DEVELOPMENT 

Fig. 5 shows some of the early prototype hardware developed under the linear concentrator program. The 

photograph shows a single linear lens with a cell receiver assembly that uses gallium arsenide/gallium 

antimonide (GaAs/GaSb) tandem cells (ref. 4) and molded optical secondaries. Further optimization and 

fabrication of the cell receiver assembly is currently under way at JX Crystals. 

The lens in Fig. 5 was one of the first lenses produced and was fabricated by hand using a single tool to mold 

the unit. Since that time, low-cost, mass production of the lens material has been demonstrated. Because the 

Fresnel facets run linearly along the inner surface of the lens, the linear lens is very easy to fabricate and lends 

itself to roll-to-roll fabrication techniques. (The lens can be fabricated in a flat form and then curved to the 

-proper shape upon integration with the array structure). Fig. 6 shows a 200 ft. roll of lens material fabricated by 

3M to ENTECH specifications. The lens fabrication process is similar to that used by ENTECH on their large 

terrestrial linear modules. The roll contains five linear lenses side-by-side, which, by proper design of the 

spacing between lenses, could be mounted to the array structure as a single unit. 

Significant progress has also been made on the development of an array structure. Fig. 7 shows different 

views of an array panel that was recently fabricated. The structure was made from carbon composite material to 

minimize weight and the effects of thermal distortions. An important point to note is that the structure 

achieves its stiffness from the honeycomb panel on the back of the structure, while the lenses are held 

accurately in place above the cell plane with a minimal amount of supporting structure. This design allows for a 

number of options with regard to reducing the stowed panel thickness by rotating the lenses down. 

Optimization of the panel structure indicates that a total panel weight of < 2 kg/m2 is readily achievable. 

Estimates for both the panel mass and performance are given in Table I. With a cell operating efficiency of 

around 25% and an optical lens performance of 90%, an array efficiency corresponding to 300 watts/m 2 is 

achieve able. This provides a panel specific power of 150 watts/kg. Based on this, a specific power of> 100 

watts/kg at the array level should be achievable in the near-term. 

SUMMARY 

The line-focus refractive concentrator array is a novel photovoltaic array concept that offers two very important 

advantages compared to some other concentrator concepts: (1) relaxation of precise array tracking 

requirements to only a single axis and (2) low-cost mass production of the lens material. This means that the 

linear refractive concentrator can be designed to provide essentially constant power for satellites using only 

single-axis tracking. Low-cost, roll-to-roll production of the lens material has been demonstrated and new 

panel designs, emphasizing light weight, high stiffness, stowability and ease of manufacturing & assembly, are 

currently being fabricated and tested. 

315 

---- -~ -------' 
REFERENCES 

1. M.F. Piszczor et aL, "A Novel Space Photovottaic Module Using a Linear Fresnel Lens and a Line-Focus 

Tandem Cell Receiver," Proc. of 23rd IEEE Photovoltaic Specialist Conf., Louisville, KY, May 1993, pp. 

1386-1391. 

2. M.F. Piszczor and M.J. O'Neill, "Development of a Dome Fresnel Lens/Gallium Arsenide Photovottaic 

Concentrator for Space Applications," Proc. of 19th IEEE Photovoltaic Specialist Conf., New Orleans, LA, 

May 1987, pp. 479-484. 

3. M.F. Piszczor et aL, "A High-Performance Photovoltaic Concentrator Array: The Mini-Dome Fresnel Lens 

Concentrator with 30% Efficient GaAs/GaSb Tandem Cells," Proc. of 22nd IEEE Photo voltaic Specialist 

Cont., Las Vegas, NV, October 1991, pp. 1485-1490. 

4. L.M. Fraas et aL, "Over 30% Efficient Tandem Gallium Solar Cells for Use With Concentrated Sunlight," 

Optoelectronics-Devices and Technology, 3, 1990, p. 297. 

316 

1 

I 

I 
Mass Estimate of the Line-Focus Fresnel Lens Concentrator Pan el 

Element Thickness Density Part Area/Aperture Area Mass/Aperture Area 

(microns) (w'cucm) (kw'so.m) 

Lens Glass Superstrate 50 2.50 1.20 0.15 

Silicone Fresnel Lens 200 1.00 1.20 0.24 

Comoosite Structure 375 1.80 1.60 1.08 

Cells & Cover Glasses 1250 2.50 0.07 0.22 

Miscell aneous 0.31 

Total 2.00 

Panel Performance Estimate 

Cell Efficiency Lens Efficiency Packinl! Factor Areal Power Specific Power 

lit Operllting Temp. 

(%) (%) (%) (W/sQ.m.) (W/kl!) 

18 90 95 211 105 

20 90 95 234 117 

22 90 95 258 129 

24 90 95 281 141 

26 90 95 305 152 

28 90 95 328 164 

30 90 95 352 176 

Table I. Mass & Performance Estimates for a Linear Refractive Concentrator Panel 

NCOMING 

LIGHT 

FOCAL LINE 

TANDEM CEll 

LtlEAR RECEIV ffi 

\ ..... 

\~~~ \ "/ PRISMSIlEANDIMTERIAL //. 

THICKNESS ClR&t.TLY EXAOOERATED lOR CLARITY 

Fig. 1. Sketch of line-focus refractive concentrator components 

detailing operation and construction of the lens/cell element. 

317 

I 

_ _ _ J 
A. Zero 

Longi t uclinal 

I11cidence 

c. 20 cleg 

Longitudinal 

111 c ide 11 c e 

B. 10 deg 

Longi t udinal 

Incidence 

D. 30 deg 

Longitudinal 

Incidence 

Fig. 2 - Effective 2-Dimensional Visualization of Focal Length 

Shortening Due to Off-Pointing Along the 

Longitudinal Axis 

318 

---------
ftj 

.2 

Ci. 

o >- II) U c: c: G> G> --I .-U 

1 

0.8 

0.6 

"'O:E 0.4 G> w U 

'6 

G> 

D: 

o 

Longitudinal Error (degrees) 

o 

40 

Latera' 

Error 

(degrees) 

Fig. 3a. Predicted optical lens efficiency vs. off-pointing for a linear refractive 

concentrator. (Current lens design without an optical secondary.) 

1 

ftj 

.2 

Ci. 0.8 

0 >- en u 0 . 6 c: c: G> G> 

--I U 

"'0 --.!! UJ 

.2 

"'0 

G> 

D: 

Longitudinal Error 

(degrees) 

o 

Lateral Error 

(degrees) 

Fig. 3b. Predicted optical lens efficiency vs. off-pointing for a linear refractive 

concentrator. (Single-axis tracking lens design without an optical secondary.) 

319 
Calculated and Measured Geometric Concentration Ratio vs. SunPointing Error Tolerance for Production-Version Line-Focus Lens (40 

deg Rim, Designed for +/- 1 deg Tolerance at 15X) with No Secondary 

Optics 

30 

Calculated for Design Image Width of 1.6 nun 

\ 

\ ....... / Two Measured Data Sets for GaAs CeUs 

'" I (90-95% Range of On-Track Perfonnance Shown) -

/ \ 

I ~ .. \, ~ 

.-.. II -............. r-..-~ 

25 

20 

Geometric 15 

Concentration Ratio 

10 

5 r Reference Data for Our Terrestrial Silicon Module I 

o , 

o 0.5 1 1.5 2 2.5 3 3.5 4 

Sun-Pointing ElTOr Tolerance (degrees) 

Fig. 4. Calculated vs. Measured Design Values for Linear Refractive Concentrators 

320 

I 

I 

I 

I 

I 

I 
'. 

Fig. 5. Ear1y prototype of linear refractive concentrator element. 

Photograph shows a single linear lens and a GaAs/GaSb tandem cell 

receiver with optical secondaries. 

Fig. 6. A 200 ft. roll of linear con. lens material. The 5-element wide roll 

was fabricated by 3M using a low-cost, roll-to-roll manufacturing process. 

321 
...... 

Fig. 7a. Prototype linear concentrator panel structure recently 

fabricated. Note that the stiffness of the carbon composite panel is 

achieved from the honeycomb panel on the back of the structure. 

Fig. 7b .. Prototype linear concentrator panel structure with linear lenses 

in place focusing light onto the back panel. Lens elements are a single 

unit cut from the 200 ft. roll of lens material. 

322 

--- ------ -- --.-- - .. - ----- --- ----- -----
59? '19 

N95- 20535 

AN INTRODUCTION TO THE ASTRO EDGE1 SOLAR ARRAY 12/,. 

B.R. Spence and G.W. Marks 

Astro Aerospace Corporation 

Carpinteria, California 

ABSTRACT 

The Astra Edge solar array is a new and innovative low concentrator power generating system 

which has been developed for applications requiring high specific power. high stiffness. low risk. light 

weight. reliability. low stowed volume. negligible thermal snap. and affordability. The basic system is of 

modular construction which utilizes conventional materials and technology. and standard photovoltaic 

solar cells and laydown processes. Mechanisms. restraint/release devices. wiring harnesses. substrates. 

and support structures are designed to be simple. functional. lightweight, and modular. A brief overview 

of the Astro Edge solar array is discussed. 

INTRODUCTION 

The Astra Edge solar array is an innovative low cost. lightweight. compact. low concentrator solar 

array of modular construction which provides significant advancements in performance. reliability and 

affordability of solar array systems. A breadboard demonstration model is shown in Figure 1. 

The Astra Edge solar array takes conventional materials and achieves a Significant advance in 

efficiency by innovative packaging and incorporation of unique deployment systems. Mechanisms are 

designed to be simple and functional. tailored to the strength of lightweight substrates. Simplistic design 

and use of reflectors to decrease the number of solar cells enables an increase in efficiency while 

decreasing the system cost. 

The Astro Edge solar array achieves low cost and low weight by using a lightweight reflector 

system to concentrate energy from the sun. obtaining the required output from 33 percent fewer solar 

cells than standard flat arrays. Net solar concentration is 150 percent (1.5 AMO). assuming a 

conservative 15 percent reflectivity, dimensional inaccuracy and thermal distortion loss. Specific powers 

for a 1000-watt EOL solar array wing of 150 W/kg with near-term multi-junction cells and 122 Wlkg with 

existing GaAs/Ge cells are realistically achievable. 

The deployed Astro Edge solar array forms a "channel" shaped configuration to provide very high 

stiffness allowing for the use of extremely lightweight substrate materials and hinges. Near-frictionless 

rolling tape type hinges and a simple damped spring-driven coordination sequence promote reliable 

deployment. 

A more compact stowed package is achieved as a result of extremely thin substrates which are in 

contact when stowed. The cells are protected for launch by cushioned reflector surfaces. The reflective 

film is mounted to a thin foam cushion on the reflector panels. In the stowed configuration. the reflector 

panels are folded onto the solar cell panels, and the stack is compressed together for launch. The stowed 

configuration forms a compact package which is both secured and cushioned to provide a highly damped 

launch environment. In the deployed condition the reflector surfaces and structural configuration 

thermally decouple extreme temperature differentials to eliminate ''thermal snap" phenomena resulting 

from exiting an eclipse. 

The Astra Edge solar array is applicable to a broad power range and provides a low risk 

alternative for future spacecraft missions requiring high performance. light weight, reliability, low stowed 

volume and affordability. An overview of pertinent components is discussed in the ensuing paragraphs. 

1 Patent applied for. 

323 

~ - ~ .. ~--~------------ -~ ~~--------~~-~ ~ 
DEPLOYMENT 

The deployment sequence is shown in Figure 2. The array is pre loaded against the spacecraft 

sidewall for launch and released wnh a lightweight hot wire cutter. On release, the hinge lines at the 

yoke/orientation drive, the yoke/panel, and the reflector paneVpanel joints start to spring open. The 

motion is controlled by a system of lightweight graphite control rods and damped with a small viscous 

damper at the root hinge. When the reflector panel/panel hinges are fully open, the reflector to cell panel 

hinges will then spring open. This deployment is undamped, but because of the low torque springs and 

the lightweight reflectors, the energy involved is small. The reflector panels will deploy until terminated by 

diagonal tension lines. These lines will accurately determine the angle of the reflector. Preload from the 

hinges will be adequate to maintain the structural configuration through all spacecraft motions. A latch or 

locking mechanism is not required. In fact, latches are not necessary on the reflector panel/panel hinges 

because their geometry will prevent any bending. 

The deployment forces associated with this lightweight solar array are very low. Spring designs 

are such that the array would survive an uncontrolled deployment. Control is required primarily to ensure 

that the array does not impact the spacecraft or any payloads. Control does not have to be precise and 

small errors in relative panel positions are acceptable. 

The coordination linkage, shown in Figure 3, is a series of parallel linkages in which the panels 

form one part of the linkage and an assembly of 0.25-inch graphite composite tubes form the other. 

Because of the allowable inaccuracy, the hinges between the links need not be precision devices. In 

addition, because of the low drive torque involved, they can be of minimum strength. 

CELL SUBSTRATES 

The Astro Edge solar array geometry inherently provides a high degree of stiffness in the 

deployed position due to the reflector and cell substrate panels configured at steep angles to each other. 

By using the high in-plane stiffness of the flat panels, the out-of-plane bending stiffness requirements are 

much lower than typical deployable solar panels. This provides major opportunities for selection of 

lightweight materials and panel designs using thin sandwich construction for reducing volume and 

providing low weight. The solar cell substrates provide a dimensionally stable surface with sufficient 

stiffness and strength to support and prevent damage to the bonded solar cells. In addition, the substrate 

also electrically isolates the cells and provides an effective heat flow path for diSSipation of waste heat. 

The substrates incorporated in the Astro Edge design are 1/8-inch-thick aluminum honeycomb core with 

O.003-inch-thick Kevlar facesheets. The low CTE of the Kevlar laminate, combined with the low moisture 

absorption of the cyanate ester matrix, provides dimensional stability near that of a graphite facesheet at 

lower weight and cost. The Kevlar also provides an inherent electrical barrier for the solar cells without 

the added weight or the concern over pin holes of a Kapton film. The facesheet on the non-cell side of 

the panel incorporates a resin matrix loaded with graphite, resulting in a black Kevlar appearance which 

increases emissivity to 0.9 for higher heat dissipation from the panel. 

REFLECTOR SUBSTRATES 

Reflector substrate materials are identical to the cell substrate materials except for the additional 

mirrored surface and under1ying foam. The back surfaces of the reflector substrates are not carbon 

loaded so that wiring harness can be directly bonded to the substrate. The integration of a 1/16-inch-thick 

polyurethane foam combined with a O.002-inch-thick aluminized Teflon film comprise the remainder of the 

reflector panel. The layer of foam provides compliance for the reflective film and protects cells during 

launch. 

HINGES 

Hinging lightweight panels requires a different design from the traditional hinges and torsion 

springs commonly used. The panel to panel hinges employed are essentially frictionless, as they rely on 

rolling motion, and wrapping and unwrapping of metal tapes. Low torque and lightweight leaf springs, 

incorporated into each hinge, are used as the driving force for deployment. An engineering model of the 

324 

) 
panel to panel hinge is shown in Figure 4. The hinge incorporates a female receptacle for easy assembly 

and changeout that becomes an integrated component of the substrate panels during lay-up. 

The yoke to panel hinges are of the same rolling tape type as on the panels, but for this 

application incorporate a locking feature to maintain deployed position since geometry will not 

automatically lock it. 

The orientation drive to yoke hinge, located at the root of the array, is the most heavily loaded and 

needs to lock out in a backlash-free manner. The baseline hinge is shown in Figure 5. At the root of the 

yoke structure, two journal bearings form a clevis hinge with the orientation drive flange. A leaf spring is 

mounted on the hinge axis to slow the whole deployment. The leaf spring is sized to provide the required 

margins over the damper and bearing friction, and the resistance from the stabilization link. The 

stabilization link is mounted to the orientation flange by a flexible pivot built of S-glass composite. It 

therefore will always have a tendency to move to the deployed position. At the other end of the link, 

rollers roll in the track mounted to the yoke and in the deployed position run over a spring latch which 

latches them against a hard stop eliminating any backlash in the joint. 

WIRING HARNESS 

Wiring harness has been designed such that at all times it assists deployment. Power is carried 

on flat beryllium copper strips attached to the backside of the reflector panels as shown in Figure 6. 

Beryllium copper cusps, preformed to deployed configurations, provide positive torque during deployment 

and bridge power between panels and other hinging points. The wiring harness is made from thin 

beryllium copper spring material, and is preformed along the hinge-lines to provide additional deployment 

force. 

SUPPORT STRUCTURE 

Because the array is an extremely lightweight device, all support structures are not required to be 

very stiff to achieve the common 0.1 Hz deployed system frequency. 

The yoke structure is an extremely lightweight structure composed of high modulus graphite 

tubes with a 0.010-inch wall thickness. This member, as with all other support structures, is sized by the 

minimum practical manufacturable thickness and easily provides more than adequate stiffness. 

STOWAGE SYSTEM 

The stowed design concept, shown in Figure 7, pre loads a sandwiched package of solar panels 

and cushioned reflector panels together against snubbers on the spacecraft sidewall. Shear load transfer 

is accomplished by features on the central snubber and selected external snubbers. 

A preloaded structural load spreader, shown in Figure 8, is attached to the backside of an 

outboard panel which spreads the load from the central tie-down to attachments at the outermost panel 

edges. The attachments are at the feet of the structural load spreader and slide in a radial direction so 

that as load is applied to the hub the feet slide outward until the hub is in contact with the central preload 

stack. Upon release, the structure hub will jump up and clear the panel, but the feet will remain attached 

to the panel. As deployment takes place, the structure remains attached to an outer panel. The extra 

weight at the outboard end of the array has minimal effect because of the inherent stiffness of the 

deployed array, and the fact that the entire structure is at a minimum manufacturable thickness. 

As mentioned previously, the solar array is preloaded against the spacecraft side wall and 

reacted at several snubbing pOints, but directly attached at the orientation drive and at a centrally located 

single restraint/release mechanism. Launch restraint/release is provided by a lightweight thermal cutter 

system as illustrated in Figure 8. The basic components are a Kevlar cord which is used to preload the 

array system and two cutting wires. The cutters are sequenced to give redundancy. Initially, the lower 

wire is heated. If this fails to cause release, then the upper wire is heated. The cutter wires are 

325 
preloaded against the Kevlar tie cord by a light torsional spring, sufficient to prevent vibration on launch, 

but not enough to cause damage to the Kevlar. 

PERFORMANCE SUMMARY 

The Astro Edge solar array can be adapted to a variety of applications ranging from Pegasusclass size to EOS-class size power ranges. Design studies have been performed to evaluate 

performance characteristics of an Astro Edge solar array modified for both a Pegasus-class size and 

EOS-size spacecraft. Table 1 shows a performance summary detailing mission requirements, EOl wing 

power, mass properties (including orientation drive). first deployed mode, first stowed mode, and stowed 

volume. 

A preliminary COSMOS/M structural model of the deployed array is shown in Figure 9. The array 

forms into an inherently stiff configuration when deployed. As a result, the first deployed mode is driven 

primarily at the yoke structure root as shown. Deployed natural frequency can be tailored by simply 

modifying the yoke structure section. 

A preliminary COSMOS/M structural model of the stowed array is shown in Figure 10. Stowed 

first mode is driven primarily by the stowage structure stiffness. Stowed natural frequency can be tailored 

by Simply modifying the stowage structure section or by increasing the number of leg supports. 

A conservative thermal analysis was performed to determine GaAs/Ge cell operating 

temperatures. Figure 11 shows the GaAs/Ge array temperature distribution for a sun synchronous Earthoriented low-Earth orbit. The preliminary analysis indicates a maximum cell operating temperature of 

89.1°C, and an average cell operating temperature of approximately 86°C. Additionally, a worst case 

thermal analysis was performed for an equatorial low-Earth orbit. Figure 12 shows the temperature profile 

for this worst case condition. The analYSis indicates an average cell operating temperature of 94°C. 

Another inherent advantage of the Astro Edge solar array lies in its ability to react large thermal 

differentials in a controlled manner without affecting overall spacecraft control. The reflective surface and 

foam interlayer on the reflector panels effectively thermally decouple extreme temperature differentials to 

eliminate thermal snap resulting from exiting an eclipse. The end result is a dynamically stable array 

during eclipse transition. 

CONCLUSION 

The Astro Edge solar array represents a unique technology applicable to a variety of power 

ranges and provides a low risk alternative for applications requiring high specific power, low weight, low 

stowed volume, reliability and affordability. 

TABLE I.-ASTRO EDGE SOLAR ARRAY PERFORMANCE SUMMARY 

WITH 3.S-Mll-THICK GaAs/Ge SOLAR CEllS 

Parameter Pegasus-Class Size EOS-Class 

Performance Size Performance 

Mission 7 years lEO 7 years lEO 

EOl Power 1364W S017W 

EOl Specific Power 113 W/kg 138 W/kg 

Deployed First Mode (governed 0.32 Hz 0.092 Hz 

by yoke structure stiffness) 

Stowed First Mode (governed by 25.1 Hz 25 Hz 

stowed structure geometry and 

stiffness) 

Stowed Volume -1.2 ft3 -15 ft3 

326 
1. Stowed 

Configuration 

Figure 1. Astra Edge Solar Array Breadboard Demonstration Model. 

2. Solar Panels 

at Partial 

Deployment 

3. Solar Panels 

at Full 

Deployment 

4. Reflector 

Panels at 

Partial 

Deployment 

Figure 2. Astra Edge Deployment Sequence. 

327 

154 1-004-02 

5813FH 
Figure 3. Astra Edge Coordination System. 

Figure 4. Astra Edge Panel to Panel Hinge Assembly. 

328 
/ 

Orientation drive 

mounting flange 

" 

Deployment 

coordination 

linkage 

Yoke structure 

Graphite/ 

epoxy 

tube 

.1fI!L! ... 

Figure 5. Astra Edge Yoke to Orientation Drive Hinge. 

PANEL WIRING CONNECTOR 

Copper-plated substrate 

Solder connections 

Conductor tape 

Adhesive 

External clip 

STOWED 

CONFIGURATION 

CUSps 

eliminate 

resistance 

torque 

Formed shape of conductor tape 

(preformed to deploy configuration) 

The conductor harness is routed across the 

yoke as shown and mounted to a Kapton 

substrate to route to the spacecraft sidewall. 

Figure 6. Astra Edge Wiring Harness. 

329 

Kevlar cable 

5835 

5833 

_~ ___ J 
Preload Structure 

Snubber 

[.IJ] ]]oneycolllh paneb 

Solar cells 

C==:J Fo a III 

\/P"""" 

\ I I I I ."., 

I I I I I I I I I I I 

( I I I I I I I I ~' ' .vW 

\ I I I 1 I I 1 ] 1 1, 

\ I I I I I I I I I I I 

I I I I I I I I I ! I I 

I I I I I I I I 

/ I 

Snuhhe r 

==- -== 

Figure 7. Astra Edge Stowed Package Concept. 

Center tie·down hub 

Tapered channel 

section load 

spreaders 

Shear I""'~"'r--.../ 

provided by cups 

and cones bonded 

I 

I 

} 

I 

" 

5837 

Molded graphite 

structure acts as 

load spreader 

to panels Release Mechanism Power Di sconnect 

Two wires heat up --_.../ 

and melt through a 

Kevlar tie cord Disconnect 

Mechanism 

(See Insel) 

Torsion spnng _____ J 

preloads carbon wires 

against the Kevlar 

' .. : ..... 

-'. ~. 

Figure 8. Astra Edge Solar Array Launch Restraint. 

330 

I 

587 1 
, \ 

•• I ',I I " 

• First deployed mode driven 

primarily by yoke structure 

stiffness at root 

Figure g, Astro Edge Solar Array First 

Deployed Mode. 

36.3 

89.1 

88.4 

85.3 

Aluminized film 

SUN 

83.1 86.8 

82.4 

79.5 

55.5 49.9 

-4.7 -7.3 

-5.1 -7.7 

Kevlar ------'" 

face 

sheets 

~EARTH 

Figure 11 . Astra Edge Solar Array Temperature 

Distribution for Sun Synchronous 

Earth Oriented Orbit. 

5848 

5904 

331 

• First stowed mode driven 

primarily stowage 'spider" 

structure stiffness 

5851 

Figure 10. Astro Edge Solar Array First 

Stowed Mode. 

Temperature (Oc) 

125 r---.,-----,.---~--:--~-,..._-__, 

75 

50 1········;···· . ··;········;···,······ ········;········., .... 

25 

o ······· ····· ········ I' ····· ········ ······· · 

-25 

-~ 1······, ·······\······· .. :, ················;······· 

-75 

-1 00 L.--'-_~---'-___ ~--'-_~_'--~----' 

o ~ ~ W M 1001~1~100100200 

lime (minutes) 

5905 

Figure 12. Astro Edge Solar Array GaAs Cell 

Temperature Profile for Worst Case 

Equatorial LEO. 
,r--

( 

I ' L ' " 

J?;3 

SESSION VI 

NON-SOLAR ENERGY 

CONVERSION 

PRiGIWNI PAGE B NK NOT FtLME 
,-

I 

RARE EARTH GARNET SELECTIVE EMITTER 

Roland A. Lowe 

Kent State University 

Kent, Ohio 

and 

Donald L. Chubb, Serene C. Farmer, and Brian S. Good 

NASA Lewis Research Center 

Cleveland, Ohio 

ABSTRACT 

3 7''1 P-20 

N95- 20536 

Thin film Ho-YAG and Er-YAG emitters with a platinum substrate exhibit high spectral emittance in 

the emission band (E)..-;!0.?5 ,411 5/2 - 4113/2' for Er -YAG and E).. ~ 0.65, 51? - 51s for Ho-YAG) at 

1500K. In addition, low out-of-band spectral emittance, E).. < 0.2, suggest these materials would be 

excellent candidates for high efficiency selective em~ters in thermophotovonaic (TPV) systems operating 

at moderate temperatures (1200-1500K). Spectral emittance measurements of the thin films were made 

(1.2<)"<3.0Ilm) and compared to the theoretical emittances calculated using measured values of the 

spectral extinction coefficient. 

In this paper we present the results for a new class of rare earth ion selective emitters. These 

emitters are thin sections ( < 1 mm) of yttrium aluminum gamet (VAG) single crystal with a rare earth 

substitutional impurity. Selective emitters in the near IR are of special interest for thermophotovoltaic 

(TPV) energy conversion 1,2. The most promising solid selective emitters for use in a TPV system are rare 

earth oxides. Early spectral emittance work3 on rare earth oxides showed strong emission bands in the 

infrared (.9 - 3 microns). However, the emittance outside the emission band was also significant and the 

efficiency of these emitters was low. Recent improvements in efficiency4,5 have been made with emitters 

fabricated from fine (5-1 0llm) rare earth oxide fibers similar to the Welsbach mantle used in gas lanterns. 

However, the rare earth garnet emitters are more rugged than the mantle type emitters. 

A thin film selective emitter6 on a low emissivity substrate such as gold, platinum etc., is rugged 

and easily adapted to a wide variety of thermal sources. The garnet structure and its many subgroups 

have been successfully used as hosts for rare earth ions, introduced as substitutional impurities, in the 

development of solid state laser crystals? Doping, dependent on the particular ion and crystal structure, 

may be as high as 100 at. % (complete substitution of yttrium ion with the rare earth ion). These materials 

have high melting points, 1940 C for YAG (Yttrium Aluminum Gamet), and low emissivity in the near 

infrared making them excellent candidates for a thin film selective emitter. 

As previously stated, the spectral emittance of a rare earth emitter is characterized by one or more 

well defined emission bands. Outside the emission band the emittance(absorptance) is much lower. 

Therefore, it is expected that emission outside the band for a thin film selective emitter will be dominated 

by the emitter substrate. For an efficient emitter (power in the emission band/total emitted power) the 

substrate must have low emittance, Es. Within the emission band the spectral emittance is governed by 

the index of refraction, nf' and the spectral extinction coefficient, a).. (sum of the absorption coefficient, 

a).., and scattering coefficient, (J)..), the emitter temperature, T E' and thickness, d. In reference 6 the 

emitter emittance and efficiency are derived as a function of the optical depth, K).. = a).. d, and the 

dimensionless emission-band energy, Eg/kT E' where Eg is the photon energy at the center of the 

emission band. This analysis shows that maximum efficiency occurs for an optimum K).. and Eg/kT E - 4. 

This paper presents normal spectral emittance, E).., measurements of holmium, (Ho) and erbium 

(Er) doped YAG thin film selective emitters at (1500 K), and compares those results with the theoretical 

spectral emittance. The spectral extinction coefficient, which is required to calculate E).., was also 

335 

GE N OT FILMEIJ 
calculated from transmission and reflectance measurements performed at room temperature. 

Specimens were cut from Czochralski grown crystals and polished on both sides with 1 J..l. diamond 

abrasive. Spectral transmission, T ')..., and reflectance, R')..., measurements were made at room temperature 

with a Perkin Elmer Lambda 19 UVNIS/NIR spectrophotometer and a Nicolet Model 750 Magna-IR 

spectrometer. The spectral extinction coefficient, IX')..., was calculated8 using: 

(1 ) 

Fig. 1 shows the spectral extinction coefficient for 25 at. % Ho-VAG (.32 mm thickness) and 40 at. % ErVAG (.65 mm thickness) specimens within the emission band. Relatively strong absorption is evident at 

the characteristic laser transitions in Er-VAG (4115/2 - 4113/

2 ) and He-VAG ( 517 - 518). Measurements with 

the Nicolet Model 750 Magna-IR spectrometer indicate the absorption is low compared to the emission 

band in the near infrared until approximately J... = 6J..l. where absorption increases greatly and continues 

through the far infrared. 

Quantities such as emissivity and absorptivity are usually thought of as properties of an isothermal 

surface. However, for the thin film rare-earth VAG emitter emission from throughout the film contributes to 

the spectral emittance. Since significant temperature gradients (200 K) exist across the specimens the 

"emitter temperature", used to calculate emittance from the intensity measurements, is not the surface 

temperature but rather the average of the front and back surface temperatures in the center of the sample. 

Temperature measurements, made with type R thermocouples, had an accuracy of +/- 6K. Spectral 

intensity measurements were made with a monochromater and a temperature controlled PbS detector 

calibrated with a 1270 K blackbody reference. Measurement accuracy and repeatability with the 

blackbody standard at 1270 K were typically within 2%. Platinum foil was placed behind the samples as a 

low emittance substrate. The configuration shown in fig. 2 was used to eliminate reflected radiation from 

the furnace interior, a near blackbody radiator, from reaching the detector. Two sources of measurement 

error were identified. Alumina insulation, a nearly perfect diffuse reflector, reflects radiation originating 

from the perimeter of the sample (a), which is at a higher temperature than the viewed area, to the surface 

of the sample in the field of view of the detector. Also, emission from the hot alumina (b) is reflected 

directly from the viewed area to the detector. Comparison of EJ... measurements of platinum foil under 

conditions identical to those required to produce an average emitter temperature of 1500 K 

(furnace@ 1664 K) and data from ref. 9 show the increase in EJ... due to reflected radiation is +0.09 @ 2.0 

J..l.m and +0.125 @ 1.5 J..l.m. This will be the maximum measurement error for materials with high spectral 

reflectance ( -80 %) and low emittance such as platinum. However, in the Ho and Er-VAG specimens, 

reflectance in the emission band is less (20-70% for Ho-VAG). Therefore, an estimate for the maximum 

experimental error in measured values of EJ... in the emission band is +/- 0.1. 

Fig. 3 shows the measured and theoretical6 spectral emittance for different values of the 

scattering albedo, OJ... = crJ.../(aJ... + crJ...)' of 25% He-VAG (.65 mm thickness) and 40% Er-VAG (1.04 mm 

thickness) at 1500 K. The theory assumes a uniform temperature for the film and the theoretical spectral 

emittance is independent of temperature. For theoretical calculations the extinction coefficients shown in 

fig.1 were used. An index of refraction of nf = 1.9 was used for the rare-earth VAG and the platinum 

substrate was assumed to have a constant spectral emittance9 of Es = 0.2. Small changes in nf «0.1) had 

a negligible effect on the theoretical results. In comparing the theoretical and experimental EJ... results, 

several features should be noted. First, results with small scattering (OJ... < 0.5) agree more closely with 

the measurements. Second, the measured emittance outside the emission band is considerably higher 

than the theoretical values. As stated earlier, emission outside of the band is primarily from the platinum 

substrate. Measurements of EJ... of pure platinum foil under identical conditions required to produce an 

average emitter temperature of 1500 K provided a value of 0.28 @ 2.4J..l.m. This value corresponds well 

336 
I 

I 

I 

I 

I 

r 

- - -- - -

with the out of band emission observed in fig. 3a,b. Values of experimental c)" for the platinum substrate 

are greater than ref. 9 due to reflected radiation seen by the detector. Finally, due to absorption in the 

rare-earth YAG, theoretical values for E}.. outside the emission band are less than the substrate emittance, 

ES= 0.2. 

In conclusion, we have measured large emittance in the emission band of Ho-YAG, E}..~.65, and 

Er-YAG, EIF.75, thin film selective emitters. Calculated spectral emittances for small scattering (o}.. < .5} 

are in fairly good agreement with the measured emittances. However, accurate emittance calculations are 

possible only if the temperature gradient across the thin emitter is included. Finally, the large emittance in 

the emission band and low out-of-band emittance suggest that Ho-YAG and Er-YAG thin film selective 

emitters may make it possible to develop an efficient TPV energy conversion system at moderate 

temperatures (1500K). 

The authors would like to acknowledge the assistance of Dr. David Hehemann of Kent State 

University in obtaining the FTIR data and Dr. Milan Kokta of Union Carbide Corporation Crystal Products 

Division for providing the Ho and Er-YAG crystals. 

REFERENCES 

1. R E. Nelson and P. A. lies, ASMEIASES/SOLTECH Joint Solar Energy Conference, (1993) 

2. P. A. /les, Conference Record of 21 st IEEE Photovoltaic Specialists Conference, 420 (1990) 

3. G.E. Guazzoni, Appl. Spectra 2 6, 60 (1972) 

4. RE. Nelson, Proceedings of the 32nd International Power Sources Symposium, 95 (1986) 

5. C.R. Parent and RE. Nelson, Proceedings of the Twenty-First Intersociety Energy Conversion 

Engineering Conference, 1314 (1986) 

6. D. L. Chubb and R A. Lowe, J. Appl Phys. 74 (9), 5687 (1993) 

7. L.G. DeShazer, S.C. Rand, and SA Wechsler, CRC Handbook of Laser Science and Technology, 

Vol. 5, Part 3, "Laser Crystals" , (CRC Press) 1987 

8. R Siegal and J. R. Howell, Thermal Radiation and Heat Transfer, (Hemisphere, Washington, DC) 2nd 

ed. 1981, p. 722 

9. Y.S. Touloukian and C.Y. Ho, Thermophysical Properties of Matter vol. 7 'Thermal Radiative 

Properties Metallic Elements & Alloys", (IFIiPlenum, New York-Washington) 1970, p. 542 (curve 39) 

337 

______ . __ . ____ J 
5 25 

; (a) 

..; 

~ 20 ·u 

<&: 

'8 15 

u 

c 

o 

.<a 10 c 

~ UJ 

(ij 

... 

.... 

o 

cu 

a. 

Vl 

E 

u 

€ 

.... 

c 

cu 

·u 

:E 

cu 

o 

U 

§ 

:;:; 

o 

c 

~ 

UJ 

(ij 

... 

t) 

5 

o 

1700 

50 

40 

30 

20 

10 

(b) 

1800 1900 2000 2100 2200 

Wavelength, A nm 

cu 

a. 

Vl oc:.---==~:::::::Z:::=--"""""-'--..L.......L--L.-~:::::1...""';::~--.J 

1300 1 400 1 500 1 600 1700 

Wavelength, A nm 

FIG. 1 Spectral extinction coefficient calculated from measured 

transmission and reflectance (a) 25% Ho-YAG (b) 40% Er-YAG 

Alumina 

a 

Pinhole 

Furnace - 1 

Interior - T = 

viewed 

b area 

Emitter 

FIG. 2 Schematic representation of the experimental setup 

338 
r 

I 

I 

~ 

w 

(l) 

u 

c 

co 

.... 

.... 

'E 

w 

co 

~ 

.... 

u 

(l) 

Q. 

V> 

co 

E 

~ 

0 

Z 

r< 

w 

(l) 

u 

c 

co 

.... 

.... 

'E 

w 

co 

~ 

.... 

u 

(l) 

Q. 

V> 

(ij 

E 

~ 

0 

z 

- --- -- - - - - -

Experimental 

0.8 No scattering 

QA,=O 

Equal scattering 

0.6 and extinction 

coefficients 

m = 0.5 . . 0.4 . . , 

... -. . . , 

. . _. _ • ... . . . 4O ....... _ ...... ----.-

... 

0.2 

o 

1400 1600 1800 2000 2200 2400 

1 

0.8 

0.6 

0.4 

0.2 

, 

, 

, 

, 

· · , 

· , 

Wavelengtht- nm 

(a) 

Experimental 

. '" / , ' ,,' V No scattering 

0).. = 0 

. . , . 

Equal scattering and 

absorption coefficents 

~ =0.5 

-. . ~ 4O' .. __ .- .. _ ••• __ • ___ ._ ............ 4O ••• 4O 

o ~~~~~~~~~~~~~~--~~~~ 

1200 1400 1600 1800 2000 2200 

Wavelength,A, nm 

(b) 

FIG.3 Theoretical and measured spectral emittance ~) .. with a platinum 

foil substrate (constant E;>... = .2) and index of refraction for YAG (n f = 1.9) 

used for theoretical calculations (a) 25% Ho· YAG .65 mm thick; front 

surface temp. = 1389 K, back surface temp. = 1616K, average emitter 

temp. = 1500K (b) 40% Er·YAG 1.04 mm thick; front surface temp. = 

1398K, back surface temp = 1606K, average emitter temp. =1500K 

339 

__ _ ___ J 
N95-20537 

InGaAs PV DEVICE DEVELOPMENT FOR TPV POWER SYSTEMS 

/' J' 

David M. Wilt, Navid S. Fatemi, and Richard W. Hoffman, Jr. 

Essential Research, Inc. 

Cleveland, Ohio 

Phillip P. Jenkins, David J. Brinker, and David Scheiman 

NYMA, Inc. 

Brook Park, Ohio 

Roland Lowe 

Kent State University 

Kent, Ohio 

and 

Donald Chubb 

NASA Lewis Research Center 

Cleveland, Ohio 

SUMMARY 

Indium Gallium Arsenide (InGaAs) photovoltaic devices have been fabricated with bandgaps 

ranging from 0.75 eV to 0.60 eV on Indium Phosphide (InP) substrates. Reported efficiencies have been 

as high as 11.2% (AMO) forthe lattice matched 0.75 eV devices. The 0.75 eV cell demonstrated 14.8% 

efficiency under a 15000 K blackbody with a prOjected efficiency of 29.3%. The lattice mismatched 

devices (0.66 and 0.60 eV) demonstrated measured efficiencies of 8% and 6% respectively under similar 

conditions. Low long wavelength response and high dark currents are responsible for the poor 

performance of the mismatched devices. Temperature coefficients have been measured and are 

presented for all of the bandgaps tested. 

INTRODUCTION 

Research in thermophotovoltaic (TPV) power systems has persisted for many years, driven by 

high prOjected thermal to electric system efficiencies.(ref. 1,2) Several variants of TPV systems have 

emerged with the principal difference being the method of thermal to radiant energy conversion. In 

blackbody based systems an emitter material is heated to produce broadband radiation. Unfortunately, 

much of the emitted energy is below the bandgap of the photovoltaic cell, therefore these systems must 

include some type of spectrum shaping element. This element must efficiently recycle the low energy 

photons back to the emitter in order to obtain high system efficiencies. 

341 

BL N' T FILM D 
The other method of producing photons from thermal sources is based on selective emitters. 

These materials emit energy in a narrow spectral band when heated, eliminating the need for additional 

spectrum shaping elements. Chubb, et ai, have demonstrated rare earth doped Yttrium Alumina Garnets 

(YAG) crystals with good selective emission properties.(ref. 3) Common to all TPV systems designed for 

operation at moderate temperatures « 15000K) is the need for a low bandgap photovoltaic device. For 

blackbody based systems, the optimum bandgap is dependant upon the operating temperature of the 

emitter and of the cell. For an emitter temperature of 1500oK, Woolf (ref. 2) has calculated that the 

optimum bandgap ranges from 0.52 eV to 0.82 eV depending upon the cell temperature . The optimum 

bandgap for selective emitter based systems depends upon the composition of the selective emitter (ref. 

1) . 

The original work in TPV utilized standard silicon solar cells. The high bandgap of silicon (Eg = 1.1 

eV) limited the systems to very high emitter temperatures (> 20000K) . RadiO-isotope and conventionally 

fueled heat sources operate at much lower temperatures, requiring lower bandgap photovoltaic devices. 

Indium Gallium Arsenide (InxGa1_xAs) is a direct bandgap semiconductor material that has a bandgap 

ranging from 0.35 eV to 1.42 eV depending on the IniGa ratio. In.53Ga.47As solar cells (Eg =0.75 eV), 

with efficiencies of up to 11 .2% (AMO), have been fabricated on lattice matched indium phosphide (InP) 

substrates (ref. 4) . 

EXPERIMENT 

InxGa1_xAs device structures were grown by Organo Metallic Vapor Phase Epitaxy (OMVPE) in a 

horizontal, low pressure reactor designed and constructed at NASA Lewis. The source gases consisted 

of trimethyl gallium, trimethyl indium, arsine (100%), phosphine (100%), diethyl zinc, and silane diluted in 

hydrogen. Typical growth conditions were: 620°C growth temperature, 190 torr reactor pressure, Villi ratio 

of 75, and carrier gas flow rate of 3.5 std. I/min. The InP substrates were zinc doped (p = 4e18 cc-1), 

oriented (100) and used as received from the vendor. A co-flow of arsine and phosphine was used at the 

time of crossover from InP growth to InGaAs growth . The co-flow lasted 10 sec. and was used to protect 

the InP substrate from decomposition until the InGaAs had formed a continuous coverage. The growth 

rate of InP was 6.1 Alsec and the growth rate of InGaAs was 8.1 Alsec. 

Device structures for the 0.75, 0.66 and 0.6 eV InxGa1_xAs devices are shown in figure 1. The 

lattice matched InGaAs device (0 .75 eV) incorporated a very thick (1 .5 Ilm) InP window layer to reduce the 

series resistance . Modeling predicts a very high short circuit current density from this device (4.7 Alcm2) 

under a 15000K blackbody emitter (approximately equivalent to 170x AMO), therefore reduction of 

resistive losses through window layer design and front contact grid design will be very important. Losses 

due to absorption in the thick InP window layer are minimal under a 15000K blackbody. 

342 
The lattice mismatched devices (Eg = 0.66 and 0.6 eV) incorporate step graded buffer layers 

between the InP substrate and the cell structure. These layers attempt to minimize the density of 

threading dislocations in the active device layers. An extensive examination of the effect of the grading 

structure on the performance of lattice mismatched devices is planned. Due to the lattice mismatch 

(0.74% and 1.2%) of the 0.66 and the 0.6 eV material. thin InP window layers were used in these devices. 

Alternate window layer materials based on InAsP and AlinAs are under development to allow the 

incorporation of thick window layers for the reduction of series resistance. 

The devices were processed using standard thermal evaporation and photolithographic 

techniques. The lattice mismatched cells were processed with a higher coverage front grid pattern to 

partially offset the limitations imposed by the high sheet resistance of the devices. Single layer antireflective coatings of Ta20S were roughly matched to the expected illumination source. 

RESULTS 

Figure 2 shows the AMO I-V data for the three different bandgap InGaAs devices without AR 

coatings. The large change in Jsc with bandgap is not directly related to bandgap. as might be thought. 

The 0.75 eV cell has a thick (1 .5 Jlm) InP window layer that dramatically reduces the AMO Jsc• which can be 

seen in the external quantum efficiency (OE) data of figure 3. Dark diode measurements of the devices 

demonstrated that they were all diffusion limited. with diode ideality factors of -1 . The dark current 

showed a large dependence on lattice mismatch as can be seen in table 1 . 

The external OE measurements (figure 3) were taken after Ta20S AR coating deposition. 

Unfortunately, our equipment limits the measurements to 1.9 Jlm. at which point the 0.6 and 0.66 eV cells 

are still operating. The roll off of the mismatched InGaAs devices at the longer wavelengths is expected 

due to the deep absorption depth of the low energy photons and the short minority carrier lifetimes 

expected in the heavily dislocated material. Optimization of base thicknesses, doping levels and lattice 

grading structures should improve the long wavelength response. 

The test devices were mounted on fixtures to facilitate their testing under blackbody and selective 

emitter illumination. The test fixture incorporated 4-wire connections for independent current and voltage 

measurement and a thermocouple mounted under the cell to monitor the operating temperature. An 

electric furnace, used for selective emitter development. was used as a 15000 K blackbody illumination 

source (fig. 4). Its' emissivity had previously been determined to be> 0.95. Calculations indicate that total 

emitted power from the black body should be 26.5 W/cm2, although measurements of the actual emitted 

power were only 3.0 W/cm2 where the cells were mounted. This difference is attributed to the reduction 

in view factor which results from the 3.6 cm separating the cell from the furnace viewport. 

343 

I 

_______ . _____ _ _ J 
The 0.75 eV cell was also measured under the illumination of an Er-YAG selective emitter at 

15000 K (fig. 4). The measured output power from the selective emitter was 1.9 W/cm2 at the cell test 

distance. This value is down from the 5.7 W/cm2 value calculated from the measured selective emitter 

(SE) spectral emmisivity data. Difficulties were experienced in accurately determining the surface 

temperature of the SE and in keeping the entire exposed surface at a uniform temperature. Due to these 

errors we will not be reporting efficiencies for the InGaAs devices under SE illumination. 

The results of the test devices under the blackbody illumination are listed in table 2. As expected, 

the cell efficiency without filters to recycle the sub-bandgap photons is very low. The 0.75eV cell is only 

able to absorb 16.8% of the total incident radiative energy. If the sub-bandgap portion of the spectrum is 

eliminated from the measurement, the efficiency increases to 14.8%. An efficiency of 29.3% was 

predicted for this device. Those predictions assumed the illumination of the cell by a perfect black body 

(emissivity =1 ) at 15000 K with a view factor of 1, and used the measured SR and dark diode characteristics 

of the actual test device. The discrepancy in efficiencies is largely attributable to the low intensity of the 

actual measurement compared to the calculated spectrum. The actual cell generated 277 mNcm2 of 

short circuit current, whereas the integration of the SR with the perfect blackbody spectrum predicted a 

Jse of 4.7 Ncm2. Achieving this optical coupling in actual practice will obviously entail the incorporation of 

optical concentrating elements, given the necessity of separating the emitter from the cell for thermal 

management reasons. Another reason for the large difference in the predicted vs. measured cell 

efficiency was a slight degradation in the cell performance after mounting on the test fixture. The cell had 

a smaller shunt resistance after mounting, leading to a reduction in the fill factor. Additional experience in 

mounting devices should eliminate this problem. Calculated cell efficiencies for the 0.66 and 0.6 eV 

devices is not included due to the incomplete QE data for these devices. 

Cell performance as a function of cell temperature is shown in figures 5-7 and table 3 under 

blackbody illumination. The temperature coefficients of Voe were very constant at -1 .6 mVfOC for all of the 

bandgaps tested . As expected, the Jse increased with increasing cell temperature, due to bandgap 

narrowing. An interesting feature of the lattice mismatched devices is the peak in the Jse at a cell 

temperature of -70°C. We believe that this is caused by increased recombination in the bulk as the 

temperature increases. The effect is more pronounced in the greater lattice mismatched 0.60 eV cell 

compared to the 0.66 eV cell. This indicates that the recombination mechanism may be related to the 

misfit and threading dislocations present in the mismatched InGaAs. 

Figure 8 shows the temperature dependence for the 0.75 eV cell under the Er-YAG selective 

emitter. The small change in Jse with increasing temperature indicates that the SE has very little emission 

outside of the Er related emission band. 

344 
CONCLUSIONS 

Lattice matched InGaAs has been demonstrated to have excellent potential for application in TPV 

power systems. Non-optimized device structures have projected efficiencies approaching 30% under 

15000 K blackbody illumination. Lattice mismatched InGaAs devices offer the ability to ''tune'' the 

photovoltaic device response to correspond to the emission band of the illumination source. Initial results 

indicate that poor long wavelength response and high dark currents need to be addressed before these 

devices are feasible. The effect of buffer layer design on device performance must be examined for lattice 

mismatched devices. We are also planning to examine the effectiveness of hydrogen passivation for 

reducing the deleterious effects of threading dislocations. Design of an actual TPV system will require 

many trade off studies. It may turn out to be be preferable to use an efficient photovoltaic device which is 

not optimally tuned to the emission source rather than a poor performance device that is tuned to the 

source. 

REFERENCES 

1. D. Chubb, R. Lowe and D. Flood, NASA TM 105755 (1992) 

2. L.D. Woolf, Solar Cells 1 9, 19 (1986). 

3 . R.L. Lowe, D.L. Chubb, S.C. Farmer and B.S. Good, "Rare Earth Garnet Selective Emitter", 

Submitted to Appl. Phys. Lett. (1994) 

4. D.M. Witt, N.S. Fatemi, R.W. Hoffman, Jr., et ai, Appl. Phys. Lett. 64 (18), p. 2415 (1994) 

345 
I 

l_ 

0.75 eV InGaAs Cell Structure 

Au/Ge Front Metallization 

Ta2~ AR Coating 

1 .5 Jim n + InP Window 

0.2 Jim n InGaAs Emitter 

2.5 Jim p InGaAs Base 

0.3 Jim p+ InP BSF 

300 Jim p+ InP Substrate 

Au/ln Rear Metallization 

Au/Ge Front Metallization 

T'2 C5 AR Coating 

~~~~imiam~&~iii~~ 0.1 Jim n+ InP Window 

I--------------i 0.27 Jim n InGaAs (0.6 eV /0.66eV) 

2.25 Jim p InGaAs (0.6 eV/0.66eV) 

I-------------i 0.17 Jim p InGaAs (0.65 eV/0.69 eV) 

0.15 Jim P ,lnGaAs (0.70 eV /0.72 eV) 

0.3 Jim p InGaAs (0.75 eV/0.75 eV) 

0.2 Jim p+ InP Buffer 

300 Jim p+ InP Substrate 

Au/ln Rear Metallization 

0.60 eV/0.66eV InGaAs Cell Structure 

Figure 1 - InGaAs TPV cell Structures 

346 
'" -

E 

0 

<t --

.s 

-c: 

Q) 

... 

... 

:J 

() 

>-

0 

c: 

Q) 

'(3 

:;:: 

-

w 

E 

:J 

-c: 

cu 

:J 

0 

Cii 

c: 

... 

Q) 

->< 

w 

50 

40 

30 

20 

10 

0 

0 

1 

0.8 

0.6 

0.4 

0.2 

0 

0.60 eV Cell 

100 200 

Voltage (mV) 

300 

Figure 2 - AMO I-V data for InGaAs solar cells 

without AR coatings. 

..... ....... - " 0.75 eV I',,~ ... ~ , .... 

400 

, 

, 1 0.66 e 

/ 

1 

11,../ 

-' 

0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2 

Wavelength (~m) 

0.75 eV Cell 

Voe = 364.8 mV 

Jse = 27.5 mAlem"2 

FF = 74.2 

AMO 11 = 5.45% 

0.66 eV Cell 

Voe = 229.5 mV 

Jse = 42.4 mAlem"2 

FF = 64.1 

AMO 11 = 4.55% 

0.60 eV Cell 

Voe = 199.1 mV 

Jse = 48.4 mA/em"2 

FF = 62 

AMO 11 = 4.34% 

Figure 3 - External QE measurements of AR Coated InGaAs Devices 

347 

-~ --------- ----- ~~~--- ------ ~--~~-~ 
8andgap(eV) A JOl (A/cm"2) Rs (0) Rsh (U) 

0.75 

0.66 

0.60 

... 

Q) , *'0 5 

~ 

0 

a.. 

Q) 8*'0 4 

> ·iii 

.~ 

E 

6*'0 W 4 - ns §. ... 

-(\I () E 

~~ 4 *'0 " 

(ij 

() 

-.;: 

Q) 

2*'0 4 

.l: 

a. 

.!!! 

E 

Q) 

::I: 0 

1.01 3.6e-8 0.453 3.4e3 

0.99 6.5e-6 0.431 2.5e3 

0.96 2.2e-5 0.387 8.0e2 

Table 1 - Dark diode data for InGaAs devices at 25 °C 

E = 0.60 eV 

9 

, . \ , ~~- Er-YAG SE ~ 

"-"--', 

0 0.5 , 1.5 2 2.5 3 3.5 4 

Wavelength (I-lm) 

Figure 4 - Blackbody and Er-YAG Selective Emitter Spectrum 

at 1500 OK. 

348 
Bandgap (eV) 

0.75 

0.66 

0.60 

Measured Cell 

Efficiency w/o filter 

# 

2.5% 

# 

1.9% 

# 

1.9% 

_ . . __ . 

Measured Cell 

Efficiency assuming 

a perfect filter 

14.8% 

8.0% 

6.0% 

Calculated Cell 

Efficiency using 

Measured SR 

29.3% 

# Cell Temperature = 34°C 

Table 2 - Performance of InGaAs Devices under Blackbody Illumination 

Linear 

Bandgap (1/Jsc)(dJsc/dT) (1/Voc)(dVoc/dT) (1/FF)(dFF/dT) (1/Pmax)(dPmax/dT) Temperature 

(eV) (x e-3/oC) (x e-3/°C) (x e-3/°C) (x e-3/°C) Range (OC) 

0.75 1.99 4.20 2.32 4.67 30-60 

0.66 3.18 7.44 6.39 1.01 30-60 

0.60 3.04 6.97 5.87 9.46 30-70 

Table 3 - Temperature Coefficients for InGaAs Cells Under 15000 K Blackbody Illumination 

349 
0 

....,(1) 

" c:: 

cu 

0 

>0 

"0 

Q) 

.!::! 

cu 

E 

.... 

0 

z 

u.-

u. 

0 

...,VI 

"0 

c:: 

cu 

0 

>0 

"0 

Q) 

.!::! 

cu 

E 

.... 

0 

z 

u: 

u. 

1.1 

Normalized J SC 

1 ....... .. .. - ....... ---' 

- - .. - 0.9 -- , .. --- _ .. ---"-, ---...!'I0rmalized V oc 

0.8 ---....... 

0.7 Fill Factor 

0.6 

0.5 

30 40 50 60 70 80 90 100 110 

1.1 

1 

0.9 

0.8 

0.7 

0.6 

0.5 

0.4 

0.3 

Cell Temperature (0C) 

Figure 5 - 0.75 eV InGaAs Cell Performance vs. Temperature 

Under 15000 K Blackbody Illumination 

""'- ...... Normalized V ....... oc 

-

20 40 60 80 100 

Cell Temperature (0C) 

Figure 6 - 0.66 eV InGaAs Cell Performance vs. Temperature 

Under 15000 K Blackbody Illumination 

350 
r- - - - - - -

..,~ 

"0 

C 

ns 

() 

>0 

-g 

.~ 

ns 

E 

0 

z 

u: 

LL 

() 

..,Ul 

"0 

c 

ns 

() 

>0 

"0 

CI.) 

~ ns 

E 

.... 

0 

Z 

LLLL 

1.1 

1 

0.9 

0.8 

0.7 

0.6 

0.5 

0.4 

0.3 

1.1 

1 

0.9 

0.8 

0.7 

0.6 

0.5 

....... Normalized J sc 

,--....... 

-

.. Fill Factor '---'-, '" "- - ... --

30 40 50 60 70 80 90 100 110 

Cell Temperature (OC) 

Figure 7 - 0.60 eV InGaAs Cell Performance vs. Temperature 

Under 15000 K Blackbody Illumination 

"-

40 

Nprmalized J

sc 

- - -....--'" -

Fill Factor ---- --, 

60 

'-, ----

80 

... _- .... -

100 

Cell Temperature (0C) 

- -

'--

120 

Figure 8 - 0.75 eV InGaAs Cell Performance vs. Temperature 

Under 15000 K Er-YAG Selective Emitter 

351 
- - - - .--- - - - - ~ - -

PULSED LASER ILLUMINATION OF PHOTOVOLTAIC CELLS 

Jane A. Yater 1 

NASA Lewis Research Center 

Cleveland, Ohio 

Roland Lowe 

Kent State University 

Kent, Ohio 

and 

Philip Jenkins and Geoffrey A. Landis 

NYMA, Inc. 

Brook Park, Ohio 

SUMMARY 

39 ?g)~ 

N95-20538 

In future space mISSIons, free electron lasers may be used to illuminate photovoltaic array 

receivers to provide remote power. The induction FEL and the radio-frequency (RF) FEL both 

produce pulsed rather than continuous output. In this work, we investigate cell response to pulsed 

laser light which simlulates the RF FEL format, producing 50 ps pulses at a frequency of 78 MHz. 

A variety of Si, GaAs, GaSb and CdInSe2 (CIS) solar cells are tested at average incident powers 

between 4 m W /cm2 and 425 m W /cm2. The results indicate that if the pulse repetition is high, cell 

efficiencies are only slightly reduced by using a pulsed laser source compared to constant illumination 

at the same wavelength. Because the pulse separation is less than or approximately equal to the 

minority carrier lifetime, the illumination conditions are effectively those of a continous wave laser. 

The time dependence of the voltage and current response of the cells are also measured using a 

sampling oscilloscope equipped with a high frequency voltage probe and current transformer. T he 

frequency response of the cells is weak, with both voltage and current outputs essentially de in 

nature. Comparison with previous experiments shows that the RF FEL pulse format yields much 

more efficient photovoltaic conversion of light than does an induction FEL pulse format. 

INTRODUCTION 

The use of high power lasers has been proposed for beaming power to remote photovoltaic arrays 

in space. Power beaming during eclipses, for instance, would eliminate the need for batteries on 

satellites in Geosynchronous Earth Orbit, thus reducing the mass of the satellite power system (ref. 

1). Night operation of a moon base can also be facilitated through earth-based laser illumination 

of photovoltaic arrays, again simplifying the power system requirements of the moonbase (ref. 2). 

Photovoltaics can have very high efficiencies under monochromatic illumination compared to solar 

light (ref. 3), creating another advantage for use of laser power beaming. Many of the issues involved 

in designing an appropriate laser and optical system have been discussed elsewhere (refs. 4 and 5) 

and will influence the ultimate selection of lasers and cell materials. 

The free electron laser (FEL) is an attractive choice of power source. It can produce megawatts 

of power and can be tuned to wavelengths appropriate for atmospheric transmission as well as 

1 National Research Council Postdoctoral Associ ate. 

353 

NK NOT ftL r: 

---------~---
L_ 

the solar cell requirements. Two FEL designs have been proposed, the induction FEL and the 

radio-frequency (RF) FEL (refs. 6 and 7). Both produce pulses of light with high power rather 

than continuous output. The induction laser produces pulses ranging from 10 to 50 nanoseconds 

wide at frequencies of 20 to 50 kHz. The RF FEL operates at MHz frequencies, producing 10 

to 40 picosecond "micropulses" which are spaced tens of nanoseconds apart. While the average 

laser power reaching the cell must be sufficient to generate the required output power, the peak 

pulse power is hundreds or thousands of times higher than the average level. The response of the 

photovoltaic receiver to the input pulses depends on the minority carrier lifetime of the solar cell 

material (refs. 8 and 9). Carriers created by the incident pulses have a finite lifetime before being 

collected at the junction and thus produce an elongated output signal. When the pulses arrive in 

rapid succession relative to the lifetime, the cells effectively see the input as a continuous source. 

However, if the pulse separation is greater than the minority carrier lifetime, the cell must respond 

to the peak power of each pulse. For Si cells, lifetimes range from 10 to 100 J-Ls for undamaged 

material, while radiation damage can lower the value to 1 J-Ls. Direct bandgap semiconductors such 

as GaAs have a much shorter minority carrier lifetime, in the range of 10 to 100 ns (ref. 10). Hence, 

the ability to convert FEL pulses to power depends on the particular laser format and the cells 

being used. 

Previous experimental studies and 1-D computer simulations have focused on the induction FEL 

format (refs. 11-15). Cell efficiencies are significantly reduced, especially for direct bandgap semiconductors. Issues such as minimizing series resistance and designing cells to avoid LC oscillations 

are challenges that must be met to successfully utilize the induction FEL. However, studies of cell 

efficiency and behavior must also be made using other laser pulse formats. In this work, we investigate the response of conventional PV cells to laser light with the RF FEL format. Using a laser 

with a pulse separation of about 10 ns, we expect the cells to respond to the average illumination 

power. The results are compared with those of a previous experiment where a copper vapor laser 

was used to simulate the induction FEL pulse format (ref. 12). 

EXPERIMENTAL PROCEDURE 

A Coherent Antares mode-locked Nd:YAG laser with 50 ps pulses at a frequency of 78 MHz is 

used to simulate the output of an RF FEL. As illustrated in figure 1, the duty cycle of the laser, 

with pulses separated by 13 ns, is 1:260. The peak power per pulse is therefore 260 times higher 

than the average laser power. In contrast, the copper vapor laser used in the previous inductionformat experiment produced 511 nm pulses which were 38 ns wide and spaced 116 J-Ls apart. The 

corresponding duty cycle of 1:3000 was significantly larger. 

The experimental apparatus is depicted in figure 2. The laser is focused by a microscope 

objective into a 300 J-Lm optical fiber and then collimated upon exiting the fiber. PV cells are 

mounted on an electrically isolated vacuum chuck which moves on a rail aligned normal to the 

optical path. A calibrated power meter, also mounted on the rail, is slid into the laser path to 

measure the time averaged power. Apertures block out all but the portion of the laser beam which 

illuminates the PV cell. The spatial uniformity of the beam over the area of the cells is within 10%. 

The frequency doubled 532 nm wavelength is used to illuminate the collection of solar cells listed 

in Table 1, many of which were tested in the induction FEL experiment (ref. 12). Use of the 532 

nm wavelength facilitates comparison with the data taken using the 511 nm Cu-vapor laser. Si, 

an indirect bandgap semiconductor, and several direct bandgap materials are included in order to 

examine the dependence of cell efficiency on minority carrier lifetime. All are planar cells, except 

for several Si and the GaSb concentrator cells. Since concentrator cells are designed with low series 

354 
resistance in order to respond to high illumination intensities and peak currents, they may yield a 

better efficiency in converting high power laser pulses. 

The cells are tested at average illumination intensities between 4 mW /cm2 and 425 mW /cm2 . 

The measurement circuit is shown in figure 3. The average dc output power (Pout =Iout XVbias) is 

determined by applying a constant dc voltage across the cell with a variable bipolar power supply 

that can sink and source current. The average dc current is measured with a digital ammeter. The 

average power is a measure of the output available in an operational situation. Because of pulse to 

pulse variation in the laser energy, the readings are averaged over several hundred laser pulses. The 

conversion efficiency is calculated at the maximum power point using the relation 

Pout 

T} = Pin A , (1) 

where A is the total cell area, Pin is the average incident laser power and Pout is the output power 

as defined above. The time dependence of the cell voltage and current is measured using a Tektronix 

11802 digital sampling oscilloscope equipped with a 200 MHz inductive current pickup and a 3.5 

GHz high-impedance sampling head. Lead lengths are kept to below 3 cm to minimize induced 

voltage caused by current transients during each laser pulse. 

RESULTS 

The time dependence of the voltage and current under pulsed illumination is one indication of 

the solar cell response to the short pulses. The voltage waveform observed on the oscilloscope shows 

the time evolution of the bias voltage across the cell, which is maintained at a nominally constant 

level through feedback control. However, current transients during the laser pulses can force the 

voltage towards the open circuit value. Figure 4a shows voltage and current waveforms for a Si 

concentrator cell illuminated with Nd:YAG pulses at 425 mW /cm2 . With an applied bias voltage 

of almost 400 m V, the resultant voltage waveform is essentially a dc signal, with a small sawtoothed 

ac component repeating every 13 ns as the laser pulses hit. The corresponding current waveform 

also shows a small response, with 10 rnA current transients also following the laser pulses. Similar 

behavior is seen for all the cells tested. The ac signal is largest at short circuit conditions, and under 

the highest laser intensities. At the maximum power point where cells are generally operated, the 

peak response is negligible, as seen in figure 4a. 

In contrast, figure 4b shows the voltage and current waveforms for the same cell and bias voltage, 

but illuminated with the copper vapor laser pulses at 279 m W /cm2 . With a significantly higher duty 

cycle and pulse separations of over 100 ps, the cell obviously exhibits a strong frequency response. 

The voltage rises in a spike as a laser pulse hits, with a slow decay over tens of microseconds down 

to the dc bias level. The current transients of over half an amp decays equally slowly as carriers 

diffuse to the depletion region. With these voltage and current pulses occurring every 116 jls, the 

cell output can hardly be maintained at a constant dc level. The response to the induction format 

pulses varied considerably from cell to cell, with the most dramatic LC oscillations occuring with 

the GaAs concentrator cells (ref. 12). However, every cell exhibited a strong frequency response and 

a corresponding reduced conversion efficiency. The RF-type pulses produce a very weak frequency 

response and, as the data will show, good efficiencies. 

Figure 5 shows current-voltage curves for a 10 n-cm planar Si cell illuminated with both the RF 

(Nd:YAG) and induction (Cu-vapor) type pulses. While the laser wavelength is comparable, the 

incident intensity is not identical. However, the cell performs better at 170 m W /cm2 under the RFsimulated pulses than at the higher average power of 263 m W /cm2 with the induction pulses. Both 

355 
the fill factor and efficiency are significantly better, while J sc is comparable for both cases. The Si cell 

is able to convert the incoming Nd:YAG pulses more efficiently than the Cu-vapor pulses, as already 

suggested by the ac waveforms discussed above. Even for Nd:YAG pulses at 470 mW /cm2 (peak 

power = 800 suns), where series resistance limiting of the current might cause deterioration of the cell 

performance, the fill factor and efficiency are essentially constant. Comparisons of direct bandgap 

cells illuminated with both pulse formats, though not shown here, are even more striking. The GaAs 

and CIS cells perform well under the Nd:YAG illumination (although the 532 nm wavelength is far 

from optimal, especially for GaSb and CIS), whereas efficiencies are exceedingly low for the Cuvapor pulse experiments. Jsc is several milliamps for the induction case but hundreds of milliamps 

under RF pulse conditions at comparable average intensities. 

Efficiencies are calculated at the maximum power point, indicated by a cross on the I-V curves, 

and are compiled in table 2 for AMO, cw (514 nm), and pulsed illumination conditions using both 

the Nd:YAG (532 nm) and Cu-vapor (511 nm) laser pulses. An Argon ion laser was used in the 

previous experiment to collect cw data (ref. 12). Efficiencies for the Si and GaAs cells tend to be 

a bit higher under monochromatic cw light than the solar spectrum, an effect which would be even 

more noticeable at the optimum wavelength of each semiconductor material. A comparison of results 

from the 532 nm pulses and the 514 nm continuous illumination, both at 170 m W /cm2 , shows that 

the pulsed laser efficiency is slightly lower for the planar cells but higher for the concentrator cells. 

However, while the PV conversion efficiency remains 70% to 99% of the cw value using RF-type 

pulses, the detrimental effect of induction-type pulses is more extreme. Si cells show an additional 

reduction in efficiency, while direct bandgap efficiencies drop to almost zero. 

Typical I-V curves of Si and GaAs cells are shown in figure 6 for 532 nm pulsed illumination at 

425,170 and 41 mW/cm2 . These average power levels correspond to approximately 3.1,1.25 and 

0.3 suns, respectively, while the equivalent peak powers are 810, 325 and 80 suns. The dependence 

of cell efficiency on average laser power is plotted in figure 7 for several different cells, while table 3 

compiles the power dependence for all cells tested. Some variation in efficiency with laser power is 

evident, with a maximum tending to occur at 170 m W /cm2. However, the curves at 425 m W /cm2 

show no sign of current saturation due to series resistance limiting at the highest peak pulse powers 

and the fill factors remain constant. Previous results with the induction formatted laser indicated 

current saturation at the highest laser intensities, where the peak power increased to 6000 suns. 

DISCUSSION 

As noted previously, the efficiencies tabulated in this paper do not represent the peak efficiencies 

expected at laser wavelengths matched to the PV bandgap. The 532 nm light used in this experiment 

is chosen so that previous results can be compared and trends noted. The wavelength of peak 

monochromatic efficiency for undamaged Si is about 950 nm (shorter for damaged material), 850 

nm for GaAs, 1600 nm for GaSb and 1000 nm for CIS. Corrections to the efficiency can be estimated 

by using the equation (ref. 4) 

1J(Apeak) 

1J(A532nm) 

Apeak QE(Apeak) - ~--- X ~--~--~ 

A532nm QE(AS32nm) , (2) 

where QE is the quantum efficiency of the cell at the selected wavelength. Assuming that the 

quantum efficiency is nearly constant with wavelength over the range of interest below the bandgap 

(confirmed by measurements of external quantum yield), the correction factor is approximately the 

ratio of the chosen wavelengths. This wavelength term simply describes the inverse proportionality 

between incident laser power and wavelength. 

356 
I 

I 

For the 532 nm Nd:YAG laser pulses, no substantial difference in efficiency is evident between 

the various materials, with all planar cells performing at 70% to 99% of the cw level. The minority 

carrier lifetime, significantly longer for Si than the other semiconductors, is not a limiting factor in 

the ability of the cells to respond to the short pulses. The I-V curves show no evidence of current 

saturation at the highest pulse intensity, where peak output currents due to individual pulses would 

be 260 times larger than the average current. Consider the limiting equation 

I < Voc / Rseries, (3) 

where the series resistance of the cells was measured previously (ref. 12). If the incident laser pulses 

were spaced sufficiently far apart so that generated carriers were all collected before the next pulse, 

every cell tested should have displayed current saturation at the higher laser powers. Such is not 

the case, indicating approximately continuous wave illumination conditions. The Si concentrator 

cells, designed to respond to higher current densities than planar cells, exhibit a modest increase in 

efficiency under the RF-type laser pulses. Despite temporal stretching of the incident pulse due to 

minority carrier diffusion, carrier concentrations rise above the average value as each pulse arrives. 

The concentrator cells are better able to collect these carriers than are the planar cells, as the results 

confirm. All the PV cells, however, convert the incident laser pulses to nearly dc output power with 

little or no loss compared to cw laser results. 

The time dependent current and voltage waveforms also indicate that the cells see quasi-cw 

illumination conditions. For GaAs, the peak to peak amplitude is never more than 20 m V, with 

corresponding current variations of about 10 rnA. Si cells sometimes exhibit larger voltage amplitudes, up to 125 m V under short circuit conditions, but the waveform at the maximum power point 

decreases to values similar to those indicated for the GaAs. As the bias voltage is increased, the 

load impedence (at the voltage supply) also increases, thereby reducing the current transients. The 

oscilloscope shows essentially dc output with a small ac signal. Analysis of the waveforms is not 

pursued given the small magnitude of the peaks and the noise often obscuring the exact shape. 

Some observations, such as a slight rounding off at the peak of the sawtooth-like voltage signal or 

the non-linear decay of some curves, could give insight into cell design for maximum efficiency under 

these pulses. However, a detailed discussion of the many factors affecting cell response to individual 

pulses can be found in the referenced studies (refs. 11-15), while the main result of this work is that 

conventional PV cells yield almost dc outputs when illuminated with high frequency pulses. 

CONCLUSIONS 

Experimental results indicate that the conversion efficiency of conventional PV cells illuminated with high frequency laser pulses is not reduced Significantly. The 532 nm wavelength of a 

mode-locked Nd:YAG laser is used to simulate the RF FEL pulse format, with the resultant cell 

performance improved compared to previous results using a Cu-vapor laser to simulate the induction FEL format. Direct bandgap cells exhibit the most significant enhancement in cell efficiency 

for incident laser intensities up to 425 mWjcm2. The ac frequency response of the cells to the 

short pulses is weak, while time averaged currents and fill factors are comparable those under cw 

illumination conditions. Because the pulse separation is as short as the minority carrier lifetime, 

the cells respond as if the incident illumination is quasi-continuous in nature. 

357 
REFERENCES 

1. Landis, G.A.: Satellite Eclipse Power by Laser Illumination. Acta Astron., 

229-233. 

2. Landis, G.A.; Bailey, S.G.; Brinker, D.J.; and Flood, D.J.: Photovoltaic Powel 

Acta Astron., vol. 22, 1990, pp. 197-203. 

3. Olsen, L.C; Huber, D.A.; Dunham, G.; Addis, F.W.; and Anheier, N.: High Eft 

matic GaAs Solar Cells. IEEE 22nd Photovoltaic Specialists Converence, Oct. 1 

4. Landis, G.A.: Photovoltaic Receivers for Laser Beamed Power in Space. IEEE ~ 

Specialists Conference, Oct. 1991, pp. 1494-1502. 

5. Weichel, H.: Laser Beam Propagation in the Atmosphere. SPIE Optical Engi 

TT-3, Bellingham, WA, 1990. 

6. De Young, RJ.; Lee, J.H.; Williams, M.D.; Schuster. G.; and Conway, E.J 

Electrically Driven Lasers for Space Power Transmission. NASA Technical M, 

1988. 

7. Burke, R: SELENE Laser Power Beaming Applications Workshop. NASA Mar 

Center, AL, March 24, 1992. 

8. Rose, B.H.: Minority Carrier Lifetime Measurements on Si Solar Cells Using Isc : 

Decay. IEEE Trans. on Electron Devices, vol. ED-31, 1984, pp. 559-565. 

9. Iles, P.A.; Non-solar Photovoltaic Cells. IEEE 21st Photovoltaic Specialists 

1990, pp. 420-425. 

10. Ahrenkiel, RK; Dunlavy, D.J.; and Hanak, T.: Photoluminescence Lifetime in 

Solar Cells, vol. 24, 1988, pp. 339-352. 

11. Anspaugh, B.; Mueller, R; Lowe, R; and Landis, G.A.: Results of IIIuminat 

Cells with Pulsed Laser Beams. JPL Publication 92-95, Nov. 1992. 

12. Lowe, R; Landis, G.A. ; and Jenkins, P.: The Efficiency of Photovoltaic Cells E 

Laser Light. 12th Space Photovoltaic Research and Technology Conference, NAS 

1992, pp. 129-146. 

13. Willow by, D.; Alexander, D.; Edge, T.; and Herren, K: Response of Silicon Sol 

Laser Illumination. 12th Space Photovoltaic Research and Technology Conference, 

Oct. 1992, pp. 147-154. 

14. Landis, G.A.; and Jain, RK: Approaches to Solar Cell Design for Pulsed Laser 

12th Space Photovoltaic Research and Technology Conference, NASA CP 3210, Oc 

166. 

15. Jain, RK; and Landis, G.A.: Modelling of High Efficiency Concentrator Sc 

Laser Pulse. SPIE Laser Power Beaming Conference, vol. 2121, Jan . 1994, pp. 264 

358 

1 

I 

I 

I 
Cell Type Type Material Area (cm 2) 

ASEC #10 Planar Si 4.0 

ASEC 10 'o-cm BSR • • ASEC 0.2 'o-cm • 

MSFC - ATM 

ASEC 10 'o-cm BSR Rad. Damaged · • 

ASEC string Planar 1.25 

Sunpower HECO 250 Concentrator Si 1.44 

ASEC 10'o-cm • 1.0 

ASEC 0.15'o-cm • 

Varian Planar GaAs 4.0 

ASEC MANTEC • • ASEC #2 • GaAs/Ge 

ASEC #16 • • 

BOEING #6701 Concentrator GaSb 0.196 

BOEING A096A Planar CIS 4.0 

Table I. Cells Tested Under Pulsed Laser Illumination Using Nd:YAG Laser 

RF Induction 

AMO cw Pulsed Pulsed 

514 nm 532 nm 511 nm 

A ~ Intensitv. mW/cm2 137 170 170 250 

Cell Type Cell Efficiency, % 

Silicon 

ASEC #10 15.0 13.3 

ASEC 10 'o-cm BSR 11 .0 14.5 10.1 5.6 

ASEC 0.2'o-cm 15.6 19.0 14.5 7.2 

MSFC - ATM 10.4 12.6 10.8 

ASEC 10 'o-cm (rad. damaged) 10.5 13.9 13.4 1.9 

ASEC planar string 11.1 7.5 

Sunoower HECO 250 (conc.) 17.2 19.2 

ASEC 10 'o-cm (conc.) 13.0 13.7 15.3 7.6 

ASEC 0.15 Q-cm (conc.) 15.2 15.3 19.0 12.1 

GaAs 

Varian 17.2 29.0 20.5 0.15 

ASEC MANTEC 16.5 28.3 23.0 

ASEC #2 17.5 24.1 

ASEC #16 18.0 26.0 25.7 

KOPIN Super (conc.) 20.7 26.6 1.3 

GaSb 

BOEING #6701 (conc.) 5.8 1.26 2.9 0.25 

CIS 

BOEING A096A 8.2 5.5 5.3 0.01 

Table II. Cell Efficiencies Under Illumination with AMO, cw Laser, and RF and 

Induction FEL Laser Format 

359 
l_ 

Pulsed Nd:YAG Laser at 532 nm 

(simulating RF FEL format) 

Average Intensity, mW/cm2 425 170 41 4 

Cell Type Cell Efficiency, % 

Silicon 

ASEC #10 12.2 13.3 13.1 9.9 

ASEC 10 O-cm BSR 9.5 10.1 10.3 8.5 

ASEC 0.2 O-cm 12.6 14.5 14.3 10.6 

MSFC - ATM 10.2 10.8 9.8 5.2 

ASEC 10 O-cm (rad. damaged) 12.1 13.4 13.3 9.8 

ASEC planar string 7.5 7.5 6.5 5.1 

Sunpower HECO 250 (conc.) 19.2 19.2 18.0 14.3 

ASEC 10 O-cm (conc.) 13.4 15.3 15.3 11.5 

ASEC 0.150-cm (conc.) 19.0 19.0 18.8 14.9 

GaAs 

Varian 14.0 20.5 17.8 

ASEC MANTEC 21 .7 23.0 21 .7 17.5 

ASEC #2 25.0 24.1 21 .8 15.4 

ASEC #16 26.3 25.7 23.5 

GaSb 

BOEING #6701 (conc.) 3.9 2.9 1.5 1.9 

CIS 

BOEING A096A 3.7 5.3 5.3 3.3 

Table III. Cell Efficiencies at Different Laser Powers for RF FEL Simulated Format 

p 

Peak Power = 260 x Average Power 

~ 

50 ps 

(532nm) 

-+----------~T 

(a) Nd:YAG laser pulses 

p 

(b) Cu-vapor laser pulses 

Figure 1. RF (a) and Induction (b) FEL pulse formats 

Coherent Anteres 

mode-locked 

Nd:YAG laser 

78 MHz Collimating Power 

Lens Meter 

Figure 2. Experimental Apparatus 

360 

Vacuum 

Chuck 

translate 

along rail 

for power 

reading 
... ' . ,. 

laser 

pulses cell 

SO-14 :::;:; 

voltage .:;:::; 

probe 

CT-2 

currentprobe 

Tektronix 11802 

digital sampling oscilloscope 

bias • 

Power Supply 

HP plotter 

Figure 3. Schematic of Measurement Circuit 

500 

>-E 

'--' 

o 

> 

300 

lIa) 

I I I I 

40.-------------------------~ 

40L---~-----L----~--~-----

o 10 20 30 

Time (ns) 

40 50 

500 

----- > E 

'--

(!) 

O'l 

0 

+"' 

(5 

> 

300 

500 

----.. « 

E 

'--' 

....... 

c: 

(l) 

'-'-

J 

U 

500 

0 

Time (us) 

b) 

10 

Figure 4. Voltage and Current Waveforms for Si Concentrator Cell. a) Illumination is 

with 532 nm Nd:YAG Pulses at a Repetition Rate of 13 ns. Average Power 

is 425 mW/cm 2. b) Same Cell Illuminated with 511 nm Copper Vapor Laser 

Pulses Arriving at 116 Ils Intervals. Average Power is 263 mW/cm2 . 

361 

___________________________ J 
1-

600 

500 

1400 

"-" 

+' 

~ 300 L.. 

L.. 

=' o 200 

100 

ASEC 10 O-cm Planar Si Cell 

.. _._.- ... _._.-... _._.- .... _._ .-.... 

470 mW /cm2 )( 

RF Laser Format -, 

II! \ 

170 mW/cm2 

,::-:-. =--7: . -:-:-. -~.~ .. ;:-:. -~-=-.,.., . ..... _=_. _. _ .•. _. _.... l. 

263 mW / cm2 -X ... , \ 

Induction Laser Format \ '. . , Ii! , \ iI O~~~~~~~~~~~~~~~~~W 

0.0 0.1 0.2 0 .3 0.4 0.5 0.6 

Voltage (V) 

FF(470) 0.700 

1)(470) 9.5% 

FF(17o) 0.710 

1)(170) 10.1% 

FF(263) 0.525 

1)(263) 5.6% 

Figure 5. I-V Curves for Planar Si Cell as a Function of Pulse Format and 

Average Laser Power. Maximum Power Points are Indicated by an X. 

250 Sun Power Concentrator Si Cell ASEC GaAs/Ge Planar Cell #2 

.. _.-... _ .... _ ... -..... -.... 200 425 mW/cm2 - .... x, 

i '\, '-" 150 I-- \ 

+' ~ C \ 

FF=0.743 

1)=19.2'; 

500 r--._ .•. _ .... -._ .... _ .... _ .... -..... 425 mW/cm2 .... . 

~400 " \ 

'-" E' \ 

+' 300 ~ C \ 

~ 170 mW/cm2 • 

FF=0.786 

1)=25.0'; 

~ 100 170 mW/cm2 \ <3 •. -.-... -.-.... -.... -.-... -.... ~ ~ FF=0.747 

.... \ i 1)=19.2'; 

:J 200 f+ ._ .• . _ ..... - .- ... _ .... - .- ... ..... %,. \ FF=0.765 

u ' . ~ 1)=24.1'; 

50 It " 41 mW/cm2 

• ~ FF=0.759 

_. 

100 41 mW/cm2 "< 

o "...-.-."...-.-_.- ........... 4- ~~ ·~t 1)=18.0'; o "'-'. --'-" -. -~~-w.~" ~::~;~~~ 

0.0 0.2 0.4 0.6 0.8 0.0 0.2 0.4 0.6 0.8 1.0 1.2 

Voltage (V) Voltage (V) 

Figure 6. I-V Curves for Si Concentrator and GaAs Planar Cell. The Fill 

Factor and Efficiencies Correspond to Each of the 3 Powers. 

25 

'-" ~20 

>. 

u 

~ 15 

'0 

;;: 

W 10 

CIl 

U 

5 

--. .. ----------.. ----_ ..... _- ,,------

GaAs#2 

GaAs 

Mantee 

/ I 

I Si Sun power (cone.) 

~======~::==========~S~i 0.20-cm 

rod. damaged 

Si string 

CIS ;'· -: ~~.· ·. ~~ ::. ·. :-:.::7 .. _._ 

- ' GaSb 

0~0~~~1~0~O~~2~O-0~~3-0~O~~~40~0~~~5~OO~w 

Average Laser Power (mW/cm2

) 

Figure 7 . Efficiency as a Function of Nd:YAG Laser Power for Various Cells. 

362 

\ 

I 

I 

I 

I 

I 
r-

/ 

..3 9 ?<t~ 

N95- 20539 

PIN In(AI) GaAs MULTIJUNCTION LASER POWER CONVERTERS 

Steven Wojtczuk and Themis Parodos 

Spi re Corporation 

Bedford, Massachusetts 

and 

Gilbert Walker 

NASA Langley Research Center 

Hampton, Virginia 

SUMMARY 

Eight In(AI)GaAs PN junctions grown epitaxially on a semi-insulating wafer were monolithically 

integrated in series to boost the -O.4V photovoltage per typical In(AI)GaAs junction to over 3 volts for 

the 1 cm2 laser power converter (LPG) chip. This is the first report of a multijunction LPG for the 1.3 to 

1.5 j..lm wavelength range. This wavelength range is optimum for laser power transmission over lowloss single-mode silica optical fiber, and is also useful with high efficiency 1.315 /lm iodine lasers in 

free-space power transmission. 

Advantages of multijunction LPG designs include the need for less circuitry for power reconditioning and the potential for lower 12R power loss. As an example, these LPC's have a 

responsivity of -1 amp/watt. With a single junction LPC, 100 watts/cm2 incident power would lead to 

about 100 Ncm2 short-circuit current at -O.4V open-circuit voltage. One disadvantage is the large 

current would lead to a large 12R loss which would lower the fill factor so that 40 watts/cm2 output would 

not be obtained. Another is that few circuits are designed to work at 0.4 volts, so DC-DC power 

conversion circuitry would be necessary to raise the voltage to a reasonable level. The multijunction 

LPC being developed in this program is a step toward solving these problems. In the above example, 

an eight-junction LPC would have eight times the voltage, -3V, so that DC-DC power conversion may 

not be needed in many instances. In addition, the multijunction LPC would have 1/8 the current of a 

single-junction LPC, for only 1/64 the 12R loss if the series resistance is the same. 

Working monolithic multijunction laser power converters (LPCs) were made in two different 

compositions of the InxAlyGa1_x_I-s semiconductor alloy, Inos3GClo.47As (0.74 eV) and Ino.sAlo.1GClo.4As 

(0.87 eV). The final 0.8 cm2 LPCs had output voltages of about 3 volts and output currents up to about 

one-half amp. Maximum 1.3 j..lm power conversion efficiencies were -22%. One key advantage of 

multijunction LPCs is that they have higher output voltages, so that less DC-DC power conversion 

circuitry is needed in applications. 

INTRODUCTION 

Laser power converters (LPCs) efficiently transform the optical power of a laser beam into 

electrical power (ref. 1). Myriad applications exist for this technology: 

Powering planetary "rover" vehicles - The original goal of this program was to develop an 

efficient converter of 1.315 /lm light that would be supplied by a high-power solar-pumped iodide laser 

(ref. 2) in orbit around a planet. This laser would be aimed at a "rover" vehicle on the planet's 

surface, so that heavy batteries or other vehicular power sources would not be needed. 

Powering orbital transfer vehicles - Powerful ground-based lasers could deliver more optical 

power to an array of LPCs on a space platform than sunlight could deliver optical power on an array of 

363 
solar cells of similar weight to the LPC array (ref. 3). As an additional advantage, because the 

LPC system is optimized to the single laser wavelength, it will also be more efficient at converting the 

laser light than a solar cell array would be at converting the multiple-wavelength sunlight. 

Safe explosive fuses - Explosives for both commercial and military uses are often detonated 

remotely by an electrical signal sent over wires to a blasting cap attached to the explosives. The wires 

can act as antennae, and a passing radio transmitter (e.g. a CB radio) can induce a voltage which 

prematurely detonates the explosive, with unfortunate consequences. If the wires are replaced by 

optical fiber, the system becomes immune to this hazard. LPCs are used to convert the light signal 

sent over the fiber into electricity to trigger the explosion safely. 

Powering remote sites - The Inos3Gao47As LPCs developed in this Phase II are ideally suited 

for use with low-loss silica optical fiber transmission systems. Such optical fiber has an attenuation 

minimum at 1.5 IJ.m (power lost is only 0.2 decibels per kilometer). Undersea cable telecommunications 

repeater stations power is one possible application. 

Power for telephones - Telephone lines superimpose weak voice signals on a steady 48V DC 

potential. This potential is used to power the phone ringer, etc. When optical fiber comes into the 

home, LPCs will most likely come with it to mimic the present system. A steady high intensity laser 

beam present on the fiber would be superimposed with weaker optical voice signals; the LPG would 

supply power to the phone ringer. 

This program (ref. 4) was undertaken since efficient photovoltaic converters did not exist 

for use with 1.3 IJ.m light. For example, silicon (Si) and gallium arsenide (GaAs) solar cells, the most 

popular types, have semiconductor bandgaps too high to absorb any of the 1.3 IJ.m laser light. The 

0.94 eV photon energy is smaller than the 1.1 eV bandgap of silicon or the 1.4 eV bandgap of gallium 

arsenide and passes through these materials without absorption. The quaternary (four-component 

alloy) compound semiconductor indium aluminum gallium arsenide InxAlyGal_x_yAs was selected as the 

cell epitaxial material because it had three desirable characteristics: 

1} The bandgap was tunable in the range from 0.36 eV to 2.2 eV. Material with any 

bandgap in this range could be obtaining by epitaxially growing the right material 

composition . 

2} A small subset of material compositions, with bandgaps ranging from 0.74 eV to 

1.52 eV, was available having the same crystal lattice constant as an available wafer, 

indium phosphide (InP). This would allow defect-free (i.e. dislocation-free) material to 

be grown with better performance than would otherwise be possible. 

3} The InxAlpa1_

x_yAs system is a (periodic table) group III-Ill-II I-V material. Because there is 

only one group V hydride, arsine, the material composition should theoretically be easier to 

control than, for example, InGaAsP, a group 1I1-III-V-V with much the same bandgap range 

as I nXAlyGa1_x_yAs. When growing InxAlyGal_x_yAs, there can be a great excess of the group V 

arsine gas present; it does not have to be tightly controlled. Since the three group III 

materials (indium, aluminum, gallium) should incorporate into the growing epitaxial film 

similarly, the composition should be easy to adjust. 

LASER POWER CONVERTER EPILAYER STRUCTURE 

Table I shows the epilayer structure for the LPCs. Both Ino53Gao47As and Ino.sAlolGao4As LPCs 

used the same overall structure shown in Table I, the only difference being the InxAlyGal_x_yAs 

composition . The epitaxial layers were grown in a Spire 100S low-pressure metalorganic chemical 

vapor deposition (MOCVD) reactor at a temperature of 690°C, using trimethylindium, triethylgallium, 

364 

L _______________ _ 
trimethylaluminum, and arsine. Dimethylzinc and silane were the P and N type dopants, respectively. 

The 690°C growth temperature was a compromise between better InxAlpa1.

x.f.s materiallifetinie 

(better at higher temperatures due less aluminum bonding with oxygen) and better compositional 

uniformity (better at low temperatures since indium has a low vapor pressure). 

Table I P-on-N laser power converter epilayer structure. 

Layer Material Doping Thickness Comments cm·3 ~m 

P, 1019 > Thickness tradeoff - QE (thin better) vs. 

Emitter In(AI)GaAs (max) 0.3 low series resistance (thick better) 

>lnGaAs is a good contact layer 

> -3 absorption lengths thick; absorbs 

Base In(AI)GaAs N,2x1017 2.5 95% of 1.3 ~m light 

> -4 hole diffusion lengths thick 

B.S. Field InP N, 1019 

0.1 > Minority carrier mirror enhances QE 

Etch Stop (max) > InGaAs mesa etch does not etch InP 

Buried InGaAs N,1019 

3 > Thick, hi-doped to lower 12R loss 

Layer (max) > Bottom contact layer 

Substrate InP Semi- 500 > Fe-doped semi-insulating wafer to 

insulating isolate the LPC multi-junctions 

LASER POWER CONVERTER FABRICATION 

The laser power converters (LPC) we have fabricated are a series connection on a semiinsulating InP wafer of In(AI)GaAs P on N mesa photodiodes. Similar devices for 800 nm use have 

been made on semi-insulating GaAs (ref. 5). The main challenge in multijunction LPC fabrication 

was establishing a process to interconnect the individual junctions. 

The first step in fabrication was isolation of the individual mesas required to form the LPCs. A 

positive photoresist process was used to pattern the mesas. Care was taken to insure the mask was 

aligned to the substrate such that any orientation dependent etches would result in profiles which could 

easily be covered by subsequent thin film depositions. A 3:4:1 H3P04:H20 2 :DI etch, selective to 

In(AI)GaAs, was used to remove the junction layers to the InP etCh-stop. Typical etch rates were 

400Ais. Etching to completion was achieved by observing a color change while the samples were 

immersed in the solution. Continuing for an additional 10s beyond this point insured total removal of 

the In(AI)GaAs junction layers. With the same photoresist in place, a 10s HCL dip removed the InP 

etch stop exposing the W In(AI)GaAs buried layer. To complete isolation of the mesas, the 3:4:1 

H3P04:HP2:DI etch was used to remove the InGaAs buried layer down to the semi-insulating InP 

substrate. Since the etch is selective, a protruding overhang of the InP etch-stop resulted as the 3:4:1 

solution removed the In(AI)GaAs. To insure continuity of the films to follow, we used a 5s HCI dip to 

remove the InP. Initially we were concerned that this final dip would undercut the InGaAs buried layer, 

however, examination of several cleaved cross sections proved that there was no undercut. In fact we 

found a transition between the InGaAs and InP which was favorable for film continuity. 

Once the mesas had been formed, the next step was exposure of the N+ InGaAs buried layer 

to form the back contact. Once again, positive resist was used. Care was taken to insure coverage of 

the mesa edges. The 3:4:1 solution was used to remove the junction layers to the etch stop and an HCI 

dip removed the InP. 

365 ~ C----d ( 
A 2000'\ of Si3N4 was deposited on the wafers by plasma assisted chemical vapor deposition 

(PACVO). The Si3N4 was then patterned to cover one side of the interconnect from the pH top contact 

to the bottom of the via. The BHF used to pattern the Si3N4 undercut the photoresist and reduced the 

width of the final pattern. Therefore to correct for the undercut, we adjusted alignment of the Si3N4 etch 

mask to insure coverage of the step. In addition, the BHF bubbled vigorously as it etched the Si3N4. 

Any bubbles which adhered to the surface acted as masks, leaving some Si3N4 underneath as the film 

etched. To minimize the amount of Si3N4 residue caused by the bubbles, we removed them at the 

midpoint of the etch using flowing 01 water. Upon completion of this process, both the Nand P sides 

of the junctions were ready for metallization. 

Based on supporting experimental data indicating good contact resistance, we chose to use the 

same metal system for ohmic contact to both the Nand P type materials. Image reversal photolithography and lift-off was used to pattern the 3 ~m thick metallization. 

The last step in LPC fabrication was deposition and patterning of the anti-reflection (AR) 

coating . A single-layer, quarter-wave PACVO Si3N4 film tuned for minimum reflectance at 1.315 ~m 

was deposited on the wafers. A final photolithographic step removed the Si3N4 from the bonding 

regions. Figure 1 shows a top view of a completed two-inch wafer. Figure 2a shows a single 

multijunction LPC and Figure 2b shows one of the tp.st sites. Figures 3a shows details of the grid lines 

too fine to see in Figure 2a, and Figure 3b shows details of the interconnect. 

Figure 1 Two-inch InP wafer with ten LPCs and two test sites. These square devices are 

designed for uniform illumination over the photoarea from a distant laser. LPCs for 

optical fibers or for free-space transmission with small diameter laser beams should 

have pie-shaped circular photoareas. 

366 

----- --- -------- - - - 
2a) 

2b) 

Figure 2 

Buried Loyer 

Contact 

Resistance 

Test 

Junctions 

Interconnects 

See 

'------'=-- Fig. 

30 

V') 

::> 

cil 

~ o 

(J) 

Q) 

Z 

Top Loyer 

Contact 

Resistance 

Test 

a) Multijunction 1 cni LPG (0.8 cni photoarea); b) test site containing large singlejunction photodiode for QE measurement, contact resistance test patterns for both the 

top (P) and bottom (N) contact layers, and an insulator test pattern. 

367 

_________ J 
3a) 

3b) 

Figure 3 

p+ InGoAs 

a) SEM detail of top contact 5 J.1m gridlines (100 J.1m centers) too fine to see in Fig. 2a. 

b) SEM detail of 100 J.1m wide metal interconnect from bottom N' In(Al)GaAs layer of 

one junction over a silicon-nitride-protected mesa edge onto the upper P InGaAs layer 

of the next junction. The number of junctions per cm for the LPG is limited by this 

"dead" interconnect area to about 10 per cm (100 J.1m dead to 900 J.1m active width). 

368 
r- 

._- - --

MEASUREMENTS 

Figure 4 shows measured absolute quantum efficiencies of Ino.s3Ga0.47As and Ino.sAlo, Gao 4As 

single junction test diodes. 

Figure 4 

100 04 1007 

InGaAs 

>- (.) 

Z 80 

III 

c::; 

ii:: 

.. 

III 

~ 60 ~ 

~ 

z 

CI 1.315J..1m 

:::I 

G 40 ..I 

CI 

Z 

II: 

III 

~ 

Z 20 -

1000 1100 1200 1300 1400 1500 

WAVELENGTH (nm) 

Quantum efficiency of 0.74 eV InGaAs and 0.85 eV InAIGaAs test diodes. The InGaAs 

results are typical, but InAIGaAs results vary considerably depending on latticemismatch and MOCVD reactor growth conditions; the InAIGaAs device here is a from a 

better MOCVD growth run. 

In general, the Ino.53Ga0.47As devices performed better than the In(AI)GaAs LPCs, even though 

the Ino.53Ga0.47As LPCs bandgap was not as favorable for this application. The quaternary In(AI)GaAs 

devices are harder to lattice-match during epitaxial growth repeatably than the ternary InGaAs devices. 

In addition, the aluminum in the InAIGaAs devices scavenges oxygen very readily. In general, quantum 

efficiencies were considerably lower for In(AI)GaAs LPCs than for InGaAs LPCs. The dark currents for 

In(AI)GaAs LPCs were often better than the InGaAs devices, presumably because of the higher 

bandgap, and the photovoltages were as good or slightly better than InGaAs devices ; however, the 

lower photocurrent and quantum efficiency of the In(AI)GaAs devices always meant the InGaAs devices 

.were better performers. 

Figure 5 below shows two illuminated I-V curves of an eight junction Ino.53Ga0.47As LPC. An 

equivalent 1.3 /lm single wavelength power density is shown for each curve. The measurements were 

actually taken with a solar simulator (concentrated white light). We now explain how we arrived at an 

equivalent 1.3 /lm light power density with the simulator. The devices were illuminated and driven to an 

arbitrary photocurrent. The 1.3 /lm quantum efficiency measured previously (similar to Figure 4) from 

the test site single junction diode is used to determine how much 1.3 /lm light would need to be 

incident on the LPC to produce the simulator photocurrent: 

369 

________ J 
Equivalent 1.3 ~ power ~ * (simulator photocu"ent) qATj 

, . 

where h is Planck's constant (J/s), c is light's velocity (m/s), q is the electron charge (G), A. is 1.3 x 10.

6 

m, and 11 is the 1.3 )lm quantum efficiency from the test site diode for the wafer. The above formula is 

a simple re-arrangement of the standard formula for the photocurrent from a photodetector. The dead 

interconnect area (10%) and grid line shadow loss (5%) of the LPG, which is not present or part of the 

test site diode, is included in this calculation. This is the power density indicated on the graph. The 

open-circuit voltage and fill-factor are exactly the same for either white light or single 1.3 )lm 

wavelength illumination as long as the total photocurrent is exactly the same, as it is by definition here. 

Figure 5 

500 0.: 'CoG 

450 4.1 watts~ 5745-2948-1 

400 

~ 350 Voc: 2.98 

g 

... 300 FF: 51.1% Eft: 16% 

z 

11/ 

ct 250 ct 

:::I 

(.) 

0 200 ... 

0 0.45 Watts ~ 150 

a. Vae: 3.02 V 

100 

FF: 70.1% Eft: 22% 

50 

0 

0.0 05 10 15 20 25 30 

PHOTOVOLTAGE 1'1) 

1-V curves of an eight-junction Ino53Gao4r4s laser power converter under two 1.3 pm 

equivalent illumination levels. At higher photocurrents the device becomes series 

resistance limited 

CONCLUSIONS 

We have demonstrated working monolithic multijunction (eight-junction) InxAlpa,.x.I-s LPGs 

with 1.3 )lm power efficiencies up to 22%. Although the (0.85 eV) Ino sAlo ,Gao4As LPGs should have 

been superior to the (0.74 eV) Inos3Gaa 47As LPGs for converting 1.315 j..I.m light (0.94 eV). additional 

material dislocation defects and oxygen recombination sites limit the current performance of the more 

complex InosAlo.,Gao4As LPGs. The simpler Ino.s3G3047As LPGs in all cases had better power 

conversion efficiencies . LPGs of both types are currently limited by series resistance to incident optical 

power densities below 10 W/cm2

. Some additional redesign of the contact grid and emitter and buried 

layers may result in some improvement of the series resistance. However, to impro'/e the operating 

power density by an order of magnitude, more junctions are needed to lower the pl··..,tocurrent and 12R 

loss. Additional development of photolithographic processing of non-planar devl;E' IS needed in order 

to successfully pattern the photoresist in the small trenches between LPG junct J' 

370 
--- -- -- - - - -

REFERENCES 

1. G.H. Walker and J.H. Heinbockel, "Mathematical Modeling of a Photovoltaic-Laser Energy 

Converter for Iodine Laser Radiation," NASA Tech. Memo. 100482. 

2. R.J. De Young, G.H. Walker, M.D. Williams, G.L. Schuster, and E.J. Conway, "Preliminary 

Design and Cost of a 1-Megawatt Solar Pumped Iodide Laser Space-to-Space Transmission 

Station," NASA Tech. Memo. 4002. 

3. GA Landis, "Photovoltaic Receivers for Laser Beamed Power in Space," Prac. of 22nd IEEE 

PVSC, pp. 1494-1502, (1991). 

4. S. WOjtczuk, S. Vernon, and E. Gagnon, "lnGaAs Concentrator Cells for Laser Power 

Converters and Tandem Cells," Proc. afthe 12th NASA SPRAT, NASA Cant. Pub. 3210, pp. 

119-128, (1992). 

5. M.B. Spitzer, R.W. McClelland, B.D. Dingle, J.E. Dingle, D.S. Hill, and B.H. Rose, "Monolithic 

Series-Connected GaAs Converter Development," Proc. of 22nd IEEE PVSC, pp. 142-146, 

(1991 ). 

371 
39?r3/ 

N95- 20540 

HIGH EFFICIENCY GaP POWER CONVERSION FOR BETAVOLTAIC APPLICATIONS1 

12 r" 

Paul E. Sims, Louis C. DiNetta, and Allen M. Barnett 

AstroPower, Inc. 

Newark, Delaware 

SUMMARY 

AstroPower is developing a gallium phosphide (GaP) based energy converter optimized for radio 

luminescent light-based power supplies. A "two-step" or "indirect" process is used where a phosphor is 

excited by radioactive decay products to produce light that is then converted to electricity by a photovoltaic 

energy converter. This indirect conversion of ~-radiation to electrical energy can be realized by applying 

recent developments in tritium based radio luminescent (RL) light sources in combination with the high 

conversion efficiencies that can be achieved under low illumination with low leakage, gallium phosphide 

based devices. This tritium to light approach is inherently safer than battery designs that incorporate high 

activity radionuclides because the beta particles emitted by tritium are of low average energy and are 

easily stopped by a thin layer of glass. GaP layers were grown by liquid phase epitaxy and pin junction 

devices were fabricated and characterized for low light intensity power conversion. AstroPower has 

demonstrated the feasibility of the GaP based energy converter with the following key results: 23.54% 

conversion efficiency under 968 flW/cm2440 nm blue light, 14.59% conversion efficiency for 2.85 flW/cm2 

440 nm blue light, and fabrication of a working 5 V array. We have also determined that at least 

20 flW/cm2 optical power is available for betavoltaic power systems. Successful development of this 

device is an enabling technology for low volume, safe, high voltage, milliwatt power supplies with service 

lifetimes in excess of 12 years. One potential application for the RL-power supply system concept is 

illustrated in Figure 1. 

Light 

~ -rOdiotion& j~;- ~ • \ \ \ ' \ . ;:-=. . ~:\: /' 

)) / \.. J:_' ,/,:~:-

if II ~ j-~ ::-

Tritium Phosphor 

Figure 1. Two step beta voltaic conversion. 

Instrumentation 

Power 

Conversion 

~ Burst 

Power 

There are three major product sectors that the development of betavoltaic technology enables: 

1) 10 to 20 year milliwatt power sources for medical implants, remote instrument packages, deep space 

probes, or security devices 2) 10 to 20 year nanowatt "static" voltage sources for on circuit board (or on 

chip) illumination and power cogeneration , and 3) integrated illumination and power generation for portable 

computers and embedded roadside/road sensors for the Intelligent Vehicle Highway System. 

1 Work funded by National Science Foundation Contract 111-9261713 

1;1 V 

373 

PR6Ga91NQ PAGE BLANK NOT FILMED 
INTRODUCTION 

The approach of this research is to utilize the tritium fueled two step method to develop cost 

effective betavoltaic-batteries W-batteries). The betavoltaic effect was discovered by Rappaport in 1953 

(ref. 1). Soon after, the Elgin-Kidde (ref. 2) two-step, 5 year atomic battery was developed. This was 

based on 147Pm/ZnS/Si. Conversion efficiencies were low, and the subsequent development of Li based 

batteries quickly made the Elgin-Kidde cell obsolete. Olsen (ref. 3, ref. 4) has reported on the conversion 

efficiency of direct betavoltaic power supplies. A review of this technology reveals that there are three 

major limitations to the direct conversion approach . 

Direct Conversion 

• the activity and range of the beta emitter 

must be coupled to the diffusion length of 

the semiconductor material 

• the power flux produced by a beta emitter 

cannot be concentrated 

• the effective ionization energy of the 

converter material limits the efficiency of 

the device 

Indirect Conversion 

• the activity and range of the beta source is 

coupled to a phosphor which is chosen so 

that the light emitted is optimal for 

conversion by the semiconductor material. 

• the use of down converting phosphors 

allows for a volumetric concentration of the 

beta energy in the form of light flux . 

• energy loss is transferred to the phosphor 

which has experimentally demonstrated 

radioluminescent conversion efficiencies 

from 10 to 30%, depending on phosphor 

type and material quality. 

Light emission has been accomplished with an areal power flux greater than that which has been 

considered to be the maximum possible for tritium gas (ref. 5). The available power density from the 

tritium/phosphor light is at least 23 IlW/cm2. Device modeling results indicate that the light to electricity 

converter of a ~-battery system can have a 20 to 30% conversion efficiency. The modeled efficiency for 

identical devices with indirect and direct conversion configurations are displayed in Figure 2. The 

efficiency of the direct conversion configuration is much less than the conversion efficiency of the "indirect" 

process. 

40 

.' . 35 Best case indirect efficiency ..[ J' • 

•• ' .' _ ••• - •• J , ;1 

. . ' 

>- g 30 

Q) 

~ 25 

w~ 

... ... .• . • - • - Me'a~~red efficiency at 440 nm 

c <J2. 20 o~ 

'(fj 

Best case direct efficiency 

W 15 

> 

§ 10 

o 

5 

o 

1.00E-09 

• -

Maximum Ni-63 energy flux 

1.00E-08 1.00E-07 1.00E-06 

Energy flux 

(W/cm2) 

Figure 2. Best case conversion efficiency of GaP. 

374 

ZnS:Ag light intensities 

• • • 

• 

Maximum tritium energy flux 

1.00E-05 1.00E-04 1.00E-03 
EXPERIMENTAL RESULTS 

Blue Phosphors 

The absorption edge for gallium phosphide extends from 549 nm to 520 nm, consequently only 

the higher photon energy blue and UV phosphors are of interest. Blue phosphors consist of a high 

bandgap semiconducting or semi-insulating material doped with mid-bandgap luminescent center 

impurities. These impurities provide midband energy levels for the radiative recombination of a carrier that 

has been excited by impact ionization to an energy state within the conduction band. In principle, 

phosphors can be highly efficient because the light is emitted within a non-absorbing lattice. In practice, 

phosphor efficiencies are lower, mainly due to non-radiative recombination mechanisms dependent on the 

material quality and material processing. It is important to note that phosphor generated light cannot be 

intensified by simply increaSing the thickness of the phosphor layer. Optical effects such as scattering, 

absorption, and light trapping will determine the optimum thickness for maximum brightness of a phosphor 

at a given level of excitation. 

For this application, power generation efficiency depends on beta-to-optical and 

optical-to-electrical conversion efficiencies, therefore, a literature search for phosphor conversion 

efficiency was conducted. The most comprehensive source of this type of data has been found in Sril 

(ref. 5). Table 1 gives a summary of the available phosphors. 

Table 1. Blue Phosphors 

Phosphor Emission Bandwidth e--beam UV 

(nm) (fwhm) Quantum Efficiency· Quantum Efficiency·· 

Ca3 (P04)2:TI 300 40 nm TlR=0.2% TlR = 49%, QE = 56% 

BaSi205:Pb 351 90 nm TlR=4% TlR = 55%, QE = 75% 

SrB40TEu 371 

Sr2P20 TEu 420 

CaW04:Pb 446 120 nm TlR= 3% TIL = 5 ImNV TlR = 42%, QE = 75% TIL = 65 1mN.J 

Sr5CI(P04)3: Eu 447 

BaMg2AI160 2TEu 450 

ZnS:Ag 450 55 nm TlR = 21% TIL = 22 ImNV 

Sr2P20 TSn 460 

MgW04 473 160 nm TlR = 2.5% TIL = 71mNV 

ZnS:Tm 477 10 nm 

Ca5F(P04)3:Sb 482 

• e -beam effiCiency IS determined at 20 keV. 

•• UV efficiency is determined at 260 nm. 

TlR - radiant effiCiency 

TIL = luminous efficiency 

From these data, it appears that silver doped zinc sulfide is a superior phosphor for this 

application. The majority, 99%, of the light emission from ZnS:Ag is above the bandgap of gallium 

phosphide (2.26 eV) and 75% of the light emission is in the spectral region where GaP has a high spectral 

response. Silver doped zinc sulfide also has demonstrated a high radiative conversion efficiency when 

excited by 20 keV mono-energetic electrons (ref. 6) and the highest conversion effiCiency (20.1 %) of the 

phosphors found in the literature search. Since the phosphor must be stable for 10 to 20 years under 

betavoltaic irradiation, lifetime studies under high energy electron irradiation need to be performed for the 

appropriate phosphors. 

375 
Phosphor packaging 

There are currently four options for packaging light generating phosphors for the p-battery. These 

are: the commercially available tritium/phosphor tube; microspheres; aerogels; and, tritirated polymers. 

The standard commercial tritium/phosphor light is basically a fluorescent light bulb where the UV 

excitation by an ionized gas is replaced by beta excitation from tritium gas. This technology is compatible 

with all of the commercially available powdered phosphors. "The maximum light power flux from a 

standard tritium gas tube RL light has been experimentally shown to be limited to - 2·3 flW/cm2 

{1 Footlambert (fL) @ 520 nanometers} because of beta self-absorption in the gas"(ref. 5). 

The second possible light source is "Self-Luminous Microspheres" (ref. 7). Microspheres consist 

of a phosphor and T2 gas enclosed in a hollow glass sphere. This is similar to the standard commercial 

package with some important differences. The spheres are very small, 0.1 mm in diameter, and each 

contains less than 3.6 x 10-5 cm3 of T 2 at STP. The possible light intensity is estimated at 1 to 

10 ft.-Iamberts. Since the glass sphere completely absorbs beta emission, no external shielding or high 

pressure containment is required for a battery assembly. The use of microspheres solves the design 

issues of safety and containment of the tritium gas, and is compatible with any phosphor. In principle, this 

type of light source can be concentrated by suspending the glass balls in a transparent matrix. However, 

there is currently no experimental evidence of the degree of concentration that can be achieved. Some of 

the items that must be considered are that the glass used to encapsulate the tritium and phosphor will 

have some absorption and that the scattering mechanisms that limit the brightness of commercial 

T2/phosphor lights still applies. 

The third possible light source is the use of aerogels. Phosphors may be suspended in a 

transparent matrix such as a silica aerogel. The advantage to this type of approach is that volumetric light 

concentration can be achieved since the aerogel is an open pore structure and tritium completely 

infiltrates the matrix. Preliminary experiments at Sandia National Laboratories (SNL) have determined that 

this type of light source can achieve an intensity of at least 23 flW/cm2 (ref. 8). This technology is 

promising and seems to have the best potential for producing a high luminous flux. 

The fourth possible technology is a completely organic, optically clear polymeric matrix 

demonstrated by Renschler et al. (ref. 9) at SNL in 1989, and Naumann (ref. 10) at E.F. Johnson in 1991 . 

This technology is based on an all-organic, clear system where tritium is covalently bound within a 

transparent polymeric matrix containing a set of soluble organic scintillant dyes. All of the components are 

distributed on a molecular scale and reside within angstroms of each other for efficient conversion. 

However, research by Webb (ref. 7) indicates that there are inherent problems with this technology; 

exposure to its own radiation causes the tritirated polymer to lose gas. Therefore, the tritirated gas 

compounds readily diffuse through the polymer leading to radiation hazards and degradation of the 

transparency of the polymer. Walko (ref. 8) comments that the overall conversion efficiency of this 

technology is less than 5%. American Atomics, Inc., Self Powered Lighting, Inc., the Oak Ridge National 

Laboratories, and the 3M company have investigated self-luminous signs and self-Iumination based on the 

tritirated polymer. 

The only technology that is immediately available is the standard commercial package. The two 

more promising technologies, the microsphere concept and the aerogel concept, are both in 

developmental stages. The microsphere is being developed by Encapsulight, Inc. The aerogel concept 

has been evaluated by SNL and some preliminary investigations have been done at AstroPower using 

aerogel material supplied by CF-Technologies, Inc. The organic matrix approach does not seem to be 

interesting for this application at the current state of development. 

376 

1 

I 

I 

I 

I 

I 
I 

I 

I 

I 

I 

I 

~Modeling 

In order to develop a cost-effective ~-battery system, a conversion efficiency of 30% at 

20 IlW/cm2 is desired. The modeling results indicate that this is possible. A 30% efficient device imposes 

stringent requirements on both the device design and on the quality and reproducibility of the growth and 

fabrication process. The device must both collect the maximum amount of current and generate power 

with minimal losses. The pertinent factors are the shunt resistance of the diode (Rsh), the depletion region 

recombination reverse saturation current (J02), the external quantum efficiency of the device (EQE), and 

the photodiode's fill factor (FF). Figure 3 displays the effects that each of these parameters has on the 

device conversion efficiency. 

30 30 

0) Effect of J. b) Effect of EQE 

25 25 

~ 

g 20 20 

J. = 10'" Nem' EQE = 70'% 

.Q1 

~ LU 15 15 c EQE = 50% .Q 

I!! 

~ 10 10 § 

U 

5 5 

a 0' 

10- 10~ 10" lD" 10~ lD" 1 0" 10" 10" lD" 10" 10' 10" lD"' 

30 30 

e) Effect of R", dJExpertn1entolDcrto 

25 25 a..Bll~2 01 

Jo - 2xlO··

6 A/cm' 

~ EQE - 56% 

R", = 2x1 rt n ern' 

> 20 20 FF - 0.83 g 

.Q1 

~ LU 15 R",= lO" Oem' 15 c ( .Q 

I!! 

~ 10 10 c 

0 

U 

5 

R", = 10'Oem' R", = 1 r:f 0 em' 

5 " -8115-05-02 

Jo - lxl0'" Ncm' 

EQE - 52 % 

R. - 7.5xl 0' n em 

a a FF - 0.75 

lD- 10~ 10" 10" 10' 10· 10"' 10" 10· 10" 10· la' 10· 10" 

lntensily (W/em') Intensily (W/em') 

Figure 3. Effects of model parameters on predicted device conversion efficiency. 

At a given light intensity conversion efficiency is linearly proportional to the external quantum 

efficiency (EQE) of the device. The fill factor also affects conversion efficiency linearly. The EQE is 

affected by surface recombination, the diffusion length of the emitter and base material, and the depth of 

the junction. For this application, the fill factor is mainly affected by the quality of the contacts at low 

current levels. Reducing J 02 increases the voltage of the device at a given light intensity. The efficiency 

increase is less, it is log-linear (i.e. <1>·ln<1» with the flux. The reverse saturation current depends on the 

carrier concentration of the material and the diffusion length in the material. The shunt resistance of the 

diode determines a threshold intensity for efficient conversion. The shunt resistance of the device seems 

to be mainly determined by the isolation technique utilized, and the depth of the junction. 

377 
The results of this modelling were used to determine the device parameters necessary to achieve 

30% conversion efficiency for the "indirect conversion" {j-battery. These are summarized in Table 2. For 

reference, parameters corresponding to the middle curves in the Figure 3 graphs ae also shown. 

Table 2. Device Modeling Results 

Device Parameter Middle Curves 30% Conversion Efficiency 

Reverse Saturation Current 10-

15 Afcm2 10-18 Afcm2 

External Quantum Efficiency 50% 80% 

Shunt Resistance 109 Q_cm2 109 Q_cm2 

Fill Factor @ 10-6 watts/cm2 70% 80% 

Device Results 

Gallium phosphide layers were grown by liquid phase epitaxy and pIn junction devices were 

fabricated and characterized for low light intensity power conversion . AstroPower has demonstrated GaP 

based energy conversion efficiency of 440 nm blue light of 23.54% at 968 IlW/cm2 and 14.59% at 

2.85 IlW/cm2 

Representative quantum efficiency curves of devices produced during this program are displayed 

in Figure 4. The external quantum efficiency (EQE) of GaP devices can be increased up to 20% by the 

application of an optimized AR coating, resulting in 70 to 80% EQE in the 380 to 480 nm spectral range. 

70 

>. 60 0 

c: 

ID 

.(3 50 lE 

w 

E 40 :l ,-. 

c~ ro- 30 :l 

a 

til 20 c: 

LID 

X 

w 10 

0 

380 

_ ..... 

-- ...... / ...... /", ...... , 

/ ' 

/ / 

/ " 

","'/ ". . ". ". 

............... '"." 

, ...... " , 

"\. "\. 

" , "\., 

" " 

" , ,"\. 

" '~ 

~ 

'). 

400 420 440 460 480 500 520 

Wavelength 

(nm) 

"-

" , 

~ 

540 

Figure 4. Quantum efficiency of some devices produced during the research program (no AR). 

The J-V curves of device 8115-05 illuminated by different wavelengths of light are displayed in 

Figure 5. This device has good shunt resistance, but the reverse saturation current is higher so that the 

photogenerated voltages are lower. Note that the conversion efficiency of 2.85 IlW/cm2 440 nm light is 

14.59%. This color and intensity of light is barely visible to the naked eye. 

378 

I 

I 

-.J 
I 

I 

0.00 

O.OOE+OO 

0.20 

Voltage (volts) 

0.40 0.60 0.80 1.00 

Figure 7. Conversion efficiency of device 8115-05 for different wavelengths and intensities of 

monochromatic light. 

1.20 

At the low operating current levels for this type of device, a high series resistance can be 

disregarded, although the contacts must be ohmic. However, high shunt resistance is critical for low 

energy conversion efficiency. 

A window layer was 

incorporated in the device design to 

passivate the surface of the device. The 

growth of a thin layer of AIGaP on the 

surface of the diode has also been found 

to significantly reduce surface 

recombination. This can be attributed to 

a combination of band bending effects 

and the possible formation of a 

passivating oxide. 

A measurement of Voe vs . Jse, 

Figure 6, at varying light intensities is 

used to determine J02 and the shunt 

resistance of the device. This technique 

eliminates the effects of series 

resistance on the J-V characteristic and 

permits a straight forward determination 

of the diode ideality factor nand 

saturation current density Jo. In 

general, such a plot will exhibit several 

distinct regions with characteristic values 

of nand J. Under low forward bias, 

n "" 2 and Jo is the depletion region 

1.00E-02 

1.00E-04 

N 

lil 5 1.00E-06 

""")~ 

1.00E-08 

1.00E-10 

o 

GaAs 

/ 

.' 

I 

I' 

/ 

/ 

f 

I 

I 

I 

I 

700 kohm-cm2 " 

I 

I 

/ 

I , 

~ 

I 

I 

, 

I 

I 

I 

I 

I 

I 

I 

I GaP 

750 Mohm-cm2 

0.5 

Voc 

(volts) 

1.5 

Figure 6. log Jsc versus Voc for GaAs and GaP devices. 

379 

2 
recombination current. Under high forward bias, n ~ 1 and Jo is due to the recombination of injected 

minority carriers in the quasi-neutral region of the base layer. High bandgap materials rarely display 

n = 1 behavior. Figure 6 displays log Jsc vs. Voc for a GaAs solar cell and a GaP energy converter. The 

light intensity varies from - 1 'IlW/cm2 to 100 mW/cm2. There is no evidence of n = 1 behavior for either 

device. Rshunt is extracted from Ohm's Law. J02 is determined by the y-intercept. At room temperature, 

the n = 2 slope is 120 mY/decade of current. 

The diffusion length of the base material is estimated from the absorption edge portion of the 

quantum efficiency curve. Due to some details in the experimental process, the values obtained are 

"characteristic" of the true value of the diffusion length. A proprietary growth enhancement has been 

found to both increase the short wavelength current generation of the device by promoting good window 

layer formation and to increase the long wavelength response of the device by increasing the diffusion 

length of the material. The "diffusion lengths" of the enhanced growth devices measured during this 

program are twice those of non-enhanced growths. 

A twelve cell array was fabricated using GaP devices grown by the non-enhanced technique. 

This array consisted of devices in a two-string configuration. The cells were current matched at an 

operating point of 0.8 volts while illuminated by an 8.24 'IlW/cm2 ZnS:Ag light source. The fabricated array 

had a better performance than the predicted values from the individual cell test data due to a current boost 

from the reduced shading for wire bonded interconnects compared to the probe station. The details of the 

current matching data are in Table 2. Figure 7 displays the J-V and P-V curve of the array illuminated by 

the phosphor light source. Since this array was fabricated from devices without growth enhancement, it is 

expected that an array with at least double the efficiency could be fabricated. Two devices (6115-50-01 

and 6-115-05-02), from enhanced growth runs, that were not incorporated into the array are included to 

show the efficiency achievable with this material at its present state of development. Array efficiencies 

above 20% should be achievable. For reference a GaAs cell is also included. 

Table 3. Cell data for array elements illuminated by == 8.24 'IlWlcm2 ZnS:Ag light 

Array Voc Isc V

M Im..{1 I at 0.8V 11 

Position Cell (V) (nA) (n ) (nA) FF (%) 

A1 8114-14-05 0.99 51 0.8 45 45 0.72 4.37 

A2 8114-10-05 1.03 54 0.8 48 48 0.69 4.66 

A3 8114-10-07 0.95 52 0.8 48 48 0.78 4.66 

A4 8114-12-05 1.00 51 0.9 44 49 0.77 4.81 

A5 8114-10-04 1.03 51 0.8 50 50 0.76 4.86 

A6 8114-14-01 1.04 53 0.8 50 50 0.73 4.86 

81 8114-12-04 1.06 52 0.9 46 51 0.75 5.03 

82 8114-12-02 1.01 52 0.9 47 52 0.80 5.14 

83 8114-12-01 1.12 53 0.9 53 52 0.76 5.46 

84 8114-14-02 1.14 54 1.0 50 53 0.81 6.07 

85 8114-10-01 1.08 55 0.9 52 54 0.79 5.68 

86 8114-10-03 1.16 55 1.0 51 55 0.80 6.19 

Array 6.27 145 5.25 116 

121 @ 

0.56 5.52 4.8 V 

8115-02-01 1.19 106 0.9 100 104 0.71 10.93 

8115-05-02 1.11 111 0.9 108 109 0.79 11 .8 

GaAs cell 0.24 251 0.2 159 0.52 3.86 

• Current boost is due to the elimination of probe shading after array fabrication. 

380 
II 

I 

l 

I 

I 

O.OOE+OO 

N -2.00E-O? E 

0 J-V 

~ 

-N 

E 

0 

~ -4.00E-O? -V 

-G.OOE-O? 

o 2 3 4 5 6 ? 

Array Voltage (V) 

Figure 8. Prototype array illuminated by the ZnS:Ag phosphor. 

CONCLUSIONS 

Completion of the Phase I program resulted in a prototype 5 volt array based on a non-optimized 

photodiode structure for the indirect conversion of tritium beta decay to electrical power. Future programs 

will concentrate on f1-battery system design, optimization of the photodiode energy converter, and the 

manufacturing solutions required for the economic production of the large semiconductor material 

quantities needed for the ~-battery . 

The ~-battery design encompasses unique advantages. For low power applications, the proposed 

beta battery has six times less volume than LiS02 batteries. An optimized f1-battery requires 60 cm3 per 

mW while a lithium battery requires 360 cm3 per mW (ref. 8). These batteries can be used in applications 

where an instrument package is to be left unattended for years in a remote location (such as at the bottom 

of the ocean or embedded in a roadway for the Intelligent Vehicle Highway System) with an intermittent 

telemetry stream or a low power housekeeping load. This type of battery will also prove useful in deep 

space probe power systems. The beta battery also has widespread commercial potential as an on-board 

power supply to maintain non-volatile memory. 

REFERENCES 

1. Rappaport, "THE ELECTRON-VOLTAIC EFFECT IN P-N JUNCTIONS INDUCED BY BETA-PARTICLE 

BOMBARDMENT," PHYSICAL REVIEW, 93, 1954, PP. 246. 

2. Elgin-Kidde ATOMIC BATTERY, "MINIATURE ATOMIC POWERED BATTERY", RADIO AND TV NEWS, V.57, 

PAGE 160, MAY 1957. 

3. Olsen, "BETAVOLTAIC ENERGY CONVERSION", ENERGY CONVERSION, 13, 117 1973. 

381 
4 . Olsen, "REVIEW OF BETAVOL TAlC ENERGY CONVERSION", PROCEEDINGS OF THE XII SPACE PHOTOVOL TAlC 

RESEARCH AND TECHNOLOGY CONFERENCE (SPRAT XII), NASA LEWIS, OCT. 1992. 

5. Walko, R.C. Lincoln, W .E . Baca, S .H . Goods, AND G .H. Negley, "TRITIUM-FuELED BETACELLS", ~ 

1991 . 

6. Bril, "ABSOLUTE EFFICIENCIES OF PHOSPHORS WITH ULTRAVIOLET AND CATHODE-RAY EXCITATION", 

LUMINESCENCE OF ORGANIC AND INORGANIC MATERIALS, H .P . KAUMANN AND G. MARMOR SPRUCH ED., 

JOHN WILEY AND SONS, 1962. 

7. Webb, "SAFE AND EFFICIENT SELF-LUMINOUS MICROSPHERES", U .S. PATENT 4,677,008, JUNE 30, 1987. 

8. Walko, C .S Ashley, C.J . Brinker, S.T. Reed , C.L. Renschler, T.J . Shepodd , R.E. Ellefson, J.T. Gill, 

AND L.E. Leonard , "ELECTRONIC AND PHOTONIC POWER ApPLICATIONS", RADIOLUMINESCENT SPECIALISTS 

CONFERENCE, ANNAPOLIS, MD, 1990. 

9. Renschler, R.L. Clough, AND T.J . Shepodd, "DEMONSTRATION OF COMPLETELY ORGANIC, OPTICALLY 

CLEAR RADIOLUMINESCENT LIGHT", JOURNAL OF ApPLIED PHYSICS, 66 (9), NOVEMBER 1987. 

10. Naumann, "COLD LIGHT BECOMES LONG-LIFE ELECTRIC GENERATOR", SENSOR, DEC. 1990. 

382 
I 

I 

I 

SESSION VII 

CELL/MATERIAL PROPERTIES 

377Z 

PR£CIf.)f PAGE ... ANK OT FtLME'-

-- --- -- -------- --- --
r ----

~C;~y~­

N95- 20541 

REVIEW OF SOLAR CELL TEMPERATURE COEFFICIENTS FOR SPACE 

Geoffrey A. Landis 

NYMA, Inc. 

Brook Park, Ohio 

INTRODUCTION 

Energy conversion efficiency is an important parameter for solar cells, and well reported in the 

literature. However, solar cells heat up when in sunlight, and the efficiency decreases. The 

temperature coefficient of the conversion efficiency is thus also extremely important, especially in 

mission modeling, but is much less well reported . It is of value to have a table which compiles into a 

single document values of temperature coefficient reported in the literature. 

In addition to modeling performance of solar cells in Earth orbit, where operating temperatures may 

range from about 20°C to as high as 85° C, it is of interest to model solar cells for several other recently 

proposed missions. These include use for the surface of Mars, for solar electric propulsion missions 

that may range from Venus to the Asteroid belt, and for laser-photovoltaic power that may involve laser 

intensities equivalent several suns. For all of these applications, variations in operating temperature 

away from the nominal test conditions result in a significant changes in operating performance. 

In general the efficiency change with temperature is non-linear, however, in the range from 

negative 100 °C through room temperature to a few hundred degrees C, efficiency is usually quite well 

modeled as a linear function of temperature (except for a few unusual cell types, such as amorphous • silicon, and for extremely low bandgap cells, such as InGaAs). Typical curves of efficiency versus 

temperature are shown in figure 1, from Reference [9]. 

TEMPERATURE COEFFICIENTS 

This is a compilation of data reported in the literature on a variety of cell ·types. Not all literature 

values were reported. Some of the literature is ambiguous (for example, not listing whether reported 

values are normalized), or lacks required information (for example, reporting power variation in mWfoC 

without reporting cell power, or reporting Isc variation without listing cell area). This compilation is also 

biased toward more recent cell types and cell types currently in production, and data on several old cell 

designs no longer in use has been left out. 

These parameters are reported in terms of the normalized coefficients, that is, the fractional change 

in value per degree Celsius, 1fT] dT]/dT. Clearly, the same coefficients apply to both efficiency and 

power, 1fT] dT]/dT= 1fP dPfdT. A negative coefficient indicates a decrease in efficiency as temperature 

rises . Thus, given the efficiency at the nominal measurement temperature To, the efficiency at Tis: 

T](T) = T](To) [l+(l/T] dT]/dT)(To -T)] (1) 

The temperature coefficient can be resolved into the sum of the variations of the open circuit 

voltage,Voc, the short circuit current, Jsc' and the fill factor, FF: 

liT] dT]/dT = l/Voc dVoc /dT + lfJsc dJsc/dT + l/FF dFF IdT (2) 

The Voc variation contributes the majority of the change in efficiency. The V oc and FF variations can 

be found in the references and are not tabulated here. The short circuit current temperature coefficient, 

1/Jsc dJsddT (= 111sc dlsddT) is reported, for reasons discussed below. 

385 
GAAs, n + p 

45 36 1.8 90 

FF 

N 35 N 28 :1: :1: 1.4 70 u u 

1 "- >- :x: 

:1: ~ 

, 

, u , 

, x 0 LL 

U ..ex: >- LL 

U') 

25 a_X:: 20 1.0 50 

15 12 L---_--'-___ .-.L.. ___ ..L....... __ --.-J 0.6 30 

120 200 280 360 440 

TEMPERATURE, K 

InP, n+p 

45 36 FF 2.0 90 

N 35 N 28 :1: x: 1.4 70 u u 

1 "- >- :x: -PMAX ~ 

:r , 

, u , , LL X 0 U 

a...~ 20 

>- LL 

U') 

25 1.0 50 

15 12 L..-__ ----.JL--___ L-__ --.-lL-..---.I 

120 200 280 360 0.6 30 

TEMPERATURE, K 

Figure 1 

Variation of maximum power (Pmax) and Voc , Isc, and fill factor with temperature for gallium arsenide 

solar cell (top) and indium phosphide solar cell (bottom) (data from Weinberg et al. [9]. used with 

permission of the author). 

386 
The wide variability in the quoted values of the Jsc temperature coefficient for cells of the same 

general type deserves discussion. The values differ not only from each other, but from the actual value 

measured under space conditions. This can be seen, for example, in a comparison of temperature 

coefficients measured using simulated sunlight compared with measurements made in space on the 

NTS-1 mission [table 4-12 of reference 32, reproduced as figure 2]. The temperature coefficients of 

voltage match to well within the error bars of the measurement. Measured temperature coefficients of 

Jsc' on the other hand, are incorrect by an average of 340 percent. 

Variation of short circuit current with temperature is primarily due to the change in bandgap energy 

with temperature. As the cell heats up, the bandgap decreases, and hence the cell responds further 

into the infrared portion of the spectrum. Hence, the Jsc variation term is roughly proportional to the 

incident spectral intensity at wavelengths near the band edge. Solar cells are not typically tested under 

actual sun illumination, however, but under a solar simulator, often a Xenon lamp. While a Xenon lamp 

has a spectrum that approximates that of the solar spectrum on the average, the intensity does not 

duplicate the solar spectrum in detail [31] . This is shown in figure 3. In particular, the spectral intensity 

near the semiconductor band edge (the range from about 800 to 1100 nm for silicon and GaAs cells) is 

significantly different from that of the sun, and in general different simulators (even of the same type) will 

have differences in the detailed structure. Thus, the variability of Jsc temperature coefficient is due to 

variations in the solar simulator, and not differences in the cells. 

Measurements of Jsc temperature coefficient made with simulated sunlight cannot be trusted. 

Fortunately, the Jsc variation accounts for only about 10-20 percent of the efficiency variation. For 

greatest accuracy, it is suggested that the measured 1/Jsc d JsddT term should be subtracted from the 

normalized power temperature coefficient to cancel this variation, and a calculated value appropriate to 

the cell material should be substituted. For Si and GaAs cells, use of values from flight experiments 

(last lines of tables 1 and 2) are suggested. 

Tables 1-2 give the compiled values of temperature coefficients from the solar cell literature, along 

with the temperature range of the measurement and the cell efficiency when listed, for silicon and GaAs 

space solar cells. The first three values in the list show current production cells for space. Table 3 

compiles temperature coefficient data for various emerging materials not yet being used for space 

power. Table 4 shows values for low-bandgap TPV cells under 15000 blackbody illumination. Note that 

since 15000 blackbody radiation contains considerably higher amounts of infrared than the solar 

spectrum, the Jsc temperature coefficient is much higher than under solar spectrum illumination. 

As expected, the temperature coefficients varies with the bandgap of the material, with the highest 

temperature dependence shown by the materials with lowest bandgaps. For comparison, table 5 

shows the theoretical values of temperature coefficient for idealized GaAs, Si, and Ge cells [24] . The 

efficiencies of these idealized cells are higher than that of those achieved today, and hence the 

temperature coefficients are slightly lower. However, the theoretical values for efficiency coefficient 

agree rather well with the measured values for the GaAs and the Ge cells, as well as for the best of the 

silicon cells. For reasons discussed above, the short-circuit current coefficients do not agree very well. 

The emphasis here is for space operation (Air Mass Zero spectrum). However, since the Voc and 

FF coefficients are not dependent on spectrum, most of this data is also usable for terrestrial 

calculations. 

Reference [24] discusses the theoretical basis for the variation of performance with temperature. 

The largest term in the temperature dependence is the voltage term, which is: 

dVoc /dT = (Voc -Eg/ q)/T - 3k/q - d(Eg/q)/dT + kT/q(l/Jsc dJsc/ dT) (3) 

387 
~(mVrC) 

dV dI 

d;c (mArC/4 cm2

) 

dT 

Ground Space Ground Space 

Exp. No. Measurement Measurement Measurement Measurement 

1 2.114 2.139 ± .178 .045 .197 ± .064 

2 2.082 2.144 ± .113 .058 .213 ± .137 

4 1.941 1.871 ± .149 .096 .271 ± .093 

5 2.191 2.098±.117 .046 .139±.047 

6 2.082 1.989 ± .140 .058 .170 ± .041 

7 1.973 2.089 ± .149 .076 .231 ± .055 

8 2.082 2.098 ± .127 .058 .225 ± .061 

Figure 2 

Measure of voltage and current temperature coefficients for 8 silicon solar cell experiments flown on 

NTS-1 satellite, comparing measurements made on the ground with those made in space [from ref. 32]. 

E 

c 

........ 

'" E 

() 

........ 

~ 

~ 

Q) 

() 

c 

ctl 

"0 

ctl 

.... 

.... 

""iii 

.... 

U 

Q) 

a. 

CI) 

300 

200 

100 

Band edge 

for GaAs 

I 

Air Mass 0 Spectrum 

Band edge 

for silicon 

Spectrolab X-25L 

O~·--~·~--~I----~--~I--~~--~I----~--~I--~~--~I--~ 

200 400 600 800 1000 1200 

Wavelength (nm) 

SOURCE: Seaman et al. 1980 

Figure 3 

Spectral comparison of Spectrolab X-25 Xenon-Arc Solar Simulator with AMO solar spectrum 

388 

L 

__ -- .- ------- ------ ---

I 

\ 

I 

I 

\ 

\ 

I 

\ 
rwhere Eg/q is the bandgap in volts, T the temperature, and k/q the thermal voltage, equal to 0.086 

mV/oC. Parameters to calculate the bandgap change with voltage for GaAs, Si, and Ge are given in 

[24], and for InP in [16]. 

The largest contribution comes from the first term, proportional to the difference between the 

bandgap and the open circuit voltage. Temperature coefficient thus decreases nearly linearly as 

voltage increases. Since higher efficiency cells typically have higher open circuit voltages, higher 

efficiency cells tend to have lower temperature coefficients than low efficiency cells of the same 

material. This can be clearly seen in table 1. 

Temperature coefficient is rarely considered as a design parameter for solar cells. From the 

standpoint of temperature coefficient, increasing open circuit voltage, even at the expense of decreases 

in other cell parameters (for example, by increasing base doping of the cell) may result in higher power 

under actual space operating conditions. 

Since open circuit voltage increases logarithmically with short circuit current, temperature 

coefficient decreases as the log of the intensity. This is shown in the graph of temperature coefficient 

versus intensity for GaAs solar cells [from Swartz and Hart, [11]). For cells with low bandgap, such as 

Ge, the fractional rise in voltage is higher, and thus the decrease in the coefficient with intensity larger. 

Since temperature coefficient is a function of illumination level, these tables do not report temperature 

coefficients for concentrator solar cells [33-36]. 

At low temperature, the linear approximation is less valid. The increase in efficiency as 

temperature decreases tends to level off, reaching a plateau typically typically around -80°C for silicon 

cells [see figures 4-57 through 4-60 of reference 32]. At low illumination intensity and low temperature 

("LILT' conditions), some solar cells are subject to additional degradation in performance. References 

[39], [40], and [41] discuss operation of silicon solar cells at low temperatures and low intensity. 

OPERATING TEMPERATURE AND EFFICIENCY 

Operating temperature can be calculated from equating the power incident on the array, Pin, with 

the power produced plus that radiated away, Pout. Here 

Pin = <X.solar P sun +albedo P albedo + <x'thermal P thermal (4) 

and 

Pout = 1](T)P sun + Erront crr4 +Ey.ear crr4 (5) 

where T is in degrees Kelvin and cr is the Stefan-Boltzmann constant, 5.67'10-8 W/m2-oK4 . For the case 

of a laser-illuminated array there are additional terms corresponding to the laser incident power and the 

laser conversion efficiency [47]. Equating these and inserting 1] as a linear function of T results in a 

fourth-order equation, which is typically solved by Newton's method. 

Note that in calculating temperatures in low Earth orbit, the contribution of sunlight reflected from 

the Earth ("albedo") and thermal infrared radiated from the Earth must be accounted for. The Earth's 

albedo varies with cloud cover and season. Average values for P albedo are quoted as "up to 30%" [31], 

"about 0.3" [28], and 0.35 [1]. Note that albedo increases significantly at high latitudes in winter due to 

snow cover; this can be important for polar orbits. The worst (highest temperature) case is at orbital 

noon, when the albedo illumination is directly on the back of the array; for this reason, the rear-surface 

alpha is an important parameter and is listed in table 6 for the cases where it has been reported. 

Thermal radiation also varies with position over the Earth and time of year. Typical thermal 

estimates use a thermal load in low Earth orbit (LEO) of about Pthermal=0.1 7 solar intensity [28,31], with 

389 

____________ J 
L __ 

Figure 4 

0 .0030 

0.0028 

0.0026 

E-

<1 

--. 0.0024 ~ 

g 

0.0022 

0.0020 

. 1 

GaAs planar cell / 

Theory: /' 

Decrease in temperature coefficient proportiona 

to voltage increase of25 mY In(lsc) 

* Jsc variation removed from temp coef. 

10 

Intensity (AMO Suns) 

100 1000 

Normalized temperature coefficient (with Isc variation subtracted) of GaAs solar cells at various incident 

solar intensities. Line shows the expected variation if the only factor in temperature coeffient is the 

variation of Voc. 

390 

I 
I 

I 

I 

the spectrum of a 288 oK blackbody [31] . The absorption constant for thermal radiation uthermal is the 

same as the thermal emissivity, E, rather than the solar absorptivity a.. Higher orbits reduce albedo and 

thermal loading proportionately to the solid angle subtended by the Earth. Hence (as can be seen in 

table 6) operating temperatures tend to be lower for GEO orbit (40,000 km altitude) than in LEO (800 km 

altitude). 

The solar absorptivity, a., is a characteristic of the solar cell, subject to some modification by 

ultraviolet and/or infrared rejection filters on the coverglass. To reduce the solar absorptance, silicon 

cells may have back-surface reflectors ("8SR") to reflect unabsorbed infrared radiation back to space, or 

even gridded back contacts to transmit unused infrared directly through the array. Cells with textured 

surfaces have higher absorptance than cells with planar surfaces, and hence higher operating 

temperature. In the infrared, the glass cover on the solar array is typically opaque, and hence the 

thermal emittance E is characteristic of the glass and independent of the cell type. 

Table 6 gives some quoted values of the thermal parameters, a and E, and also shows the 

calculated equilibrium operating temperatures in orbit for several cells. 

Future cells may have advanced covers which more efficiently reflect undesired IR and UV 

radiation [43,44]. This can reduce the solar alpha for silicon cells to 0.75 for cells without back-surface 

reflectors or gridded back contacts [43], and to 0.72 to 0.79 [43, 44] for GaAs/Ge cells. 

TEMPERATURE COEFFICIENTS FOR NON-SOLAR SPECTRA 

There has recently been some interest in use of photovoltaic cells for converting laser radiation 

[45,46]. Equation 2 shows the components of the variation of efficiency with temperature. Of the three 

terms, only the Jsc term should depend on spectrum. Hence, the normalized temperature coefficient 

un der laser illumination can be calculated from the efficiency and Jsc tenns: 

[Ill) dl)/dTllaser = [Ill) dl)/dTlsola r- [l/Jsc dJsc/dT lsola r+ [lIJsc dJsc/dThaser (5) 

Since the Jsc term typically contributes 10-20% of the efficiency variation with temperature, to 

single-digit accuracy the normalized efficiency temperature coefficient should be roughly the same for 

laser or solar illumination [47J. The V DC term will have a slight dependence on spectrum because the 

laser wavelength may be chosen to be close to the cell spectral response maximum, and hence the 

current output for a given intensity input will be higher, resulting in a logarithmic increase in voltage and 

a slight decrease in temperature coefficient. 

Under solar illumination the short circuit current increases with temperature. Under monochromatic 

illumination this mayor may not be true. There are two regions of operation, depending on the 

wavelength of the laser and the shape of the cell spectral response (in AIW). The two regions are: (1) 

wavelength shorter than the spectral response peak, or (2) wavelength longer than the spectral 

response peak. 

For wavelengths shorter than the spectral response peak, the semiconductor absorbs essentially 

all of the incident light, and this does not change as the bandgap changes with temperature. Changes 

in short circuit current are only due to changes in quantum efficiency, which should be close to unity for 

high efficiency cells. Since the number of incident photons is not dependent on the bandgap, the 

current is nearly independent of temperature, and the temperature coefficient of Jsc is near zero. 

Hence, for this region of operation, the monochromatic temperature coefficient can be approximated by 

subtracting the (solar) normalized Jsc temperature coefficient from the normalized efficiency coefficient. 

For a cell operating at wavelengths longer than the peak of the spectral response this will not be 

391 
true. At long wavelengths the laser light is only weakly absorbed, and hence a small change in the 

bandgap will result in a large change in absorption. For such cells the J sc will increase with 

temperature. This effect may be especially important in indirect bandgap materials, such as a silicon 

cell operated at 1 .0611. For such regimes of operation the change of absorption with temperature will be 

very significant, and the cell may even increase in efficiency with temperature. 

The theoretical monochromatic J sc temperature coefficient is: 

[l/Jsc dJsc/dThaser = A/Eg (dQE/dA)(dEg/dT) (6) 

where QE is the cell quantum efficiency at the laser wavelength A. The variation in bandgap energy 

with temperature for most materials, dEg/dT, can be found in the literature [16,24]. 

Solar cells can be used for spectra other than solar or laser. Wilt et al [47] discusses the 

temperature coefficient of 0.6 to 0.7 eV InGaAs cells under 1500 °C blackbody radiation; these results 

are shown in table 4. Note that, for these cells, the Jsc component of the temperature coefficients is a 

much larger component of the total. This is because of the low bandgap of the cell and the fact that the 

1500° C blackbody has a large amount of its radiation in the infrared. The fact that these cells are 

operated under high intensity also means that the V oc component of the temperature coefficient is 

somewhat reduced by the logarithmic dependence of V oc. 

For blackbody spectra, as for laser spectra, the temperature coefficient can be computed from the 

solar temperature coefficient by subtracting out the measured Jsc coefficient and replacing it with the Jsc 

coefficient for the spectrum desired. 

CONCLUSIONS 

The variation of solar conversion efficiency with temperature has been reviewed. The efficiency is 

assumed to be linear with temperature. This is correct for temperatures near 25° C for most cell types, 

but the behavior is nonlinear at extremely high and low temperatures. Typically the increase of 

efficiency with reduction of temperature flattens out below- 200 OK. Also, at extremely high 

temperatures, the efficiency does not go to negative values, but levels off near zero. References 32 and 

33 have data on (1967 vintage) silicon cells to temperatures down to 82 OK. The Design Handbook has 

data on later vintage silicon cells at low intensity and low temperatures [32] . 

392 

___ J 
Table 1: Silicon Celi Temperature Coefficients 

The first cell listed, SSF gridded back, is the large-area silicon cell developed for the space 

station Freedom project. This, and the Applied Solar Energy Corporation (ASEC) Back 

Surface Field/Reflector (BSFR) and ASEC Back Surface Field (BSF) cells, may be taken 

typical of silicon cells currently used in flight, however, they are not significantly different in 

performance from the other silicon cells listed, with the exception of the last two. 

The University of New South Wales (UNSW) cells are recently developed laboratory cells 

with improved open-circuit voltage and high efficiency. Cells of this design are not yet 

qualified for use in space. 

normalized tern!!. coeffiCients 

Cell Temp T] liT] dT]/dT l!Jsc dJsc /dT Ref. 

~ ~ (28°C) (xlO-3°C-l) (xlO-3°C-l) 

SSF grid back 0 -95 .13 -4.5 [1) 

ASECBSFR 28-60 . 135-. 148 -4.60 [2] 

ASECBSR 28-60 .125-. 134 -4.45 [2] 

2n High E. 0 - 140 .122 -4.62 +0.74 [3] 

IOn Helios 0 - 140 .116 -4.72 +0.59 [3] 

Ion BSF 0 -75 .140 -5.0 [4] 

2n K5 0 -75 .136 -5.0 [4] 

"Violet" 0 - 120 .14 -4.2 +0.65 [5] 

CNR (text.) 10-70 . 148 -4 .35 +0.34 [5] 

AEG IOn BSR .122 -3.79 +1.14 [6] 

UNSW MINP 5 -60 .1870 -3.443 +0.650 [7] 

UNSW PESC 5 -60 .1907 -3.202 +0.650 [7] 

NTS- 1 space measurement (avg. 7 types) +1.12 [32] 

Table 2: GaAs Cell Temperature Coefficients 

The ASEC GaAs/Ge cells listed on the first two lines can be taken as typical of cells that 

are currently flown in space. 

Cell Temp 

~ -..ITL 

ASEC GaAs/Ge 20- 120 

ASEC GaAs/Ge28 -60 

ASEC IGaAs 28-60 

Spire IGaAs 10-80 

Varian 60 

EEV LPE 

Hughes LPE 0-415 

Fr ench LPE 40-200 

LPE 25-350 

DH 15-80 

Hughes pin 0 -80 

LL nip homo. 0 -80 

Japan LPE p/nO- IOO 

Varian 10-80 

normalized tern!! . coefficients 

T] 

(28°C) 

liT] dT]/dT l!Jsc dJsc/dT Ref. 

(x 1 0-3°C-l) (x 1 0-3°C-l) 

.174 

.18-.185 

.18-.185 

.195 

.174 

. 157 

.212 

.156 

. 164 

.166 

. 175 

-1.60 +0.830 

-2.23 +0.56 

-2 .32 +0.56 

- 1.47 +0.246 

-2.2 +0.6 

- 1.90 +0.92 

-2.09 +1.1 

-2.53 

-2.64 

-2.74 

-2.0 

-2.6 

-2. 10 

+0.37 

+0.30 

+0.714 

+0.71 

+0 .63 

(high altit ude test) +0.508 

393 

[8) 

(2) 

[2) 

[38) 

[9] 

[10] 

[11 ) 

(12) 

(13) 

(14) 

[15] 

[15) 

[37] 

[31) 
Table 3: Temperature Coefficients from Cells of Other Materials 

nonnalized tem12 . coefficients 

Cell Temp l) 1/l) dl)/dT l/Jsc dJsc/dT Ref. 

~ ~ (28°Cl (x 1 0-3°C-l) (x 10-3°C-l) 

1.7eVAlGaAs 

AlGaAs 15- 80 - 1.55 +0.95 [ 14) 

1.93 eV AlGaAs 

Varian 25-96 (high altitude t est) +0.761 (31) 

InP 

diffused (best) 60 .132 -3 .45 +0.767 [ 16) 

diffused (wrst) 60 .103 -3 .56 +0.966 (16) 

RPI 60 .136 -2.8 +0.8 [9] 

MO-CVD 0 - 150 .195 - 1.59 +0.890 [17) 

ITO/InP 15-80 -3 .80 +0.515 [ 14) 

Ge 

RTI 20- 80 .090 - 10. 1 +0.617 (18) 

CuInSe2 

Boeing CIS -40-80 .087 -6.52 [ 19) 

ISET CIS -40-80 .088 -6.03 (19) 

Boeing CIS 25 .08 -5.26 +0.43 [20] 

CIS/CdZnS 15-80 -5.87 +0.260 (14) 

CIS/CdS -6 .880 +0.057 (21) 

Amorphous Si allous Reported a-si cells have nonlinear response with 

temperature. Data reported here is for the region listed. 

(singLe junction) 

Solarex 0 - 40 

ECD a-si:H,F 15-80 

ECD aSiGe:H,F 15-80 

*nonlinear 

(two junction) 

.066 - 1.11* +0.74 (19) 

-0 .98 to - 1.97* 0.83 to 0 .95* 114) 

- 1.02 to - 1.97* 1.01 to 1.36* (14) 

Fuji 22- 60 .089-.10 2 .0 [42) 

(22) 

[22) 

(23) 

GaAs/Ge Tandem 

low temp.* 35- 100 .194 

high temp.* 100-180 - .18 

Spire 25- 80 .189 

[see also ref. 38) 

·nonlinear 

-2 .85 

-2 .0 

- 1.54 

394 

+1.02 

+1 .02 

+0.94 
Table 4: Temperature Coefficients For Theromphotovoltaic (TPV) Cells under 

1500°C blackbody radiation 

Note that Jsc temperature coefficient will be much higher under 1500° blackbody radiation 

than under the solar illumination. 

normalized temp. coefficients 

Cell Temp l iT] dT]/dT I/Jsc dJsc/dT Ref. 

~ ~ (xl0-3°C-l) (xl0- -1 ) 

0 .6eV InGaAs 30-70 -9.46 +3.04 (48) 

0 .66eV InGaAs 25-60 -1 0 .1 2 +3 .18 (48) 

0 .75eV InGaAs 30- 60 -4.67 +2.00 (48) 

note: the J sc temperature coefficients for these cells are highly non-linear above 60°C. 

Table 5: Calculated Values of Temperature Coefficients 

These val ues are for theoretical cells with perfonnance at or near the theoretical limit. Numbers are not 

representative of actual cells in use today. 

nonnalized temI!. coefficients 

Cell Temp T] liT] dT]/d T I/Jsc dJsc/dT Ref. 

Material ~ (28°C) (xl0-3°C-l) (Kl.Q-3 -1 ) 

GaAs (calc.) 27 .277 -2.40 +0.34 (24) 

Si (calc.) 27 .247 -3 .27 +0.293 (24) 

Ge (calc.) 27 .106 -9 .53 +0 . 125 (24) 

395 

------- - --. -- - -
Table 6: Other Thermal Parameters 

.c&.ll Front Rear Olleratin~ Orbit Ref. 

~ -L ~ -L temll . (0C) 

Silicon Cells/Arrays 

8x8 (SSF) 0 .62 0.85 0 .25 0 .85 60 (Tl= 13.5) LEO [1 ) 

ASEC grd back 0 .65 (Tl= 14.2%) [2] 

ASECBSR 0.68 (Tl= 12 .5- 13.4%) [2) 

ASEC BSFR 0.72-0.74 (Tl= 13.5- 14.8%) [2) 

Si (HST) 0 .75(cell) 0 .83 0 .54 0 .70 50-61 LEO [25) 

(BSFR SI) 0 .54(array) 

Silicon 0 .80-0 .840.81-0.84 0 .81 [26) 

Si K4, K6 0 .63 23 (Tl=14% ) LEO [4 ) 

(planar SI. grtdded back) 

Si K5, K7 0 .81 43 (Tl=14% ) LEO [4 ) 

(textured Sl . grtdded back) 

SiBSFR 0 .757 0 .83 0 .54 0 .898 71 (Tl=8%) LEO 

4 4 (Tl=8%) GEO [27) 

Si bifacial 0.66 0 .83 0 .57 0 .78 57 (Tl=9 % ) LEO [27) 

Si BSF 0.82 0.82 0 .86 42 (Tl= 14.2%) (28) 

Thin Si 0.72 0.86 0 .86 27 (Tl=13. 5%) [28) 

IR refl . Si 0 .69 0 .86 0 .86 22 (Tl= 14.4%) [28) 

IR transp. Si 0 .67 0 .86 0 .86 21 (Tl=13 .3%) [2 8 ) 

ASEC BST 0 .65 0.88 [2 9 ) 

Sharp BSFR Si 0 .75 0 .82 0 .80 37 (Tl= 14. 3%) [30) 

textured cell 0 .93 0 .77 [5) 

(textured Sl. no BSR) 

textured cell 0 .905 0 .75 [5) 

(above. with conductive coated covergiass) 

GaAs Cells 

ASEC /Ge 0.870 (Tl=18 - 18.5%) [2) 

ASEC/GaAs 0 .83 (Tl= 18- 18.5%) [2) 

thin GaAs 0.82 0 .86 0 .86 36 (Tl= 16.0%) LEO [281 

Sharp GaAs 0 .86 0 .82 0 .80 45 (Tl= 18.0 %) [301 

GaAs/Ge 0.88 0 .80 0 .20 0 .85 84 polar 5600nm {221 

GaAs 0 .84 0 .83 0.540.898 72 (Tl=9 % ) LEO 

• 48 (ij=9% ) GEO [271 

396 
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398 
[37] N. Takata, et aI, "Space Proven GaAs Solar Cells - Main Power Generation for CS-3," Proc. 21st 

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Specialists Conference, 97-100 (1968). 

[40] C.H. Liebert, "Solar Cell Performance at Jupiter Temperature and Solar Intensity," Proc. 7th IEEE 

Photovoltaic Specialists Conference, pp. 92-96 (1968). 

[41] IEEE, session 2B, "Temperature-Illumination Studies," Proc. 8th IEEE Photo voltaic Specialists 

Conference, pp. 110-168 (1970) 

[42] Y. Ichikawa, "Large-Area Amorphous Silicon Solar Cells With High Stabilized Efficiency," Proc. 

23rd IEEE Photovoltaic Specialists Conf., pp. 27-33 (1993). 

[43] K. Mullaney et al., "Infra-Red Reflective Coverglasses: the Next Generation," Proc. 23rd IEEE 

Photo voltaic Specialists Conf., pp. 1363-1368 (1993). 

[44] w.T. Beauchamp, T.T. Hart and M.L. Sanders, "Blue/Red Reflecting Solar Cell Covers for GaAs 

Cells," Proc. 23rd IEEE Photovoltaic Specialists Cont. , pp. 1487-1490 (1993). 

[45] G.A. Landis, "Satellite Eclipse Power by Laser Illumination" (paper IAF-90-053), Acta Astronautica, 

Vol. 25 No. 4, 229-233 (1991). 

[46] G.A. Landis, "Space Power by Ground-based Laser Illumination," IEEE Aerospace and Electronics 

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[47] G.A. Landis,"Photovoltaic Receivers for Laser Beamed Power," Journal of Propulsion and Power, 

Vol. 9No. 1, 105-112 (1993); presented at 22nd IEEE Photovoltaic Specialists Conference, Las 

Vegas NV, Oct. 1991 , Vol. II, 1494-1502. 

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F. DeAngelo, "High Efficiency, Low Bandgap InGaAs PV Device Development for TPV Power 

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16 1994 (proceedings in press). 

399 

--
I~ · 

, 

I 

I 

I 

1/ 31 '9o;")t"i 

N9S-20S42 

EFFECTS OF PLASMA HYDROGENATION ON TRAPPING PROPERTIES OF 

DISLOCATIONS IN H!=TEROEPITAXIAL InP/GaAs 1 

S.A. Ringel and B. Chatterjee 

The Ohio State University 

Columbus, Ohio 

ABSTRACT 

In previous work, we have demonstrated the effectiveness of a post-growth hydrogen plasma 

treatment for passivating the electrical activity of dislocations in metalorganic chemical vapor deposition 

(MOCVD) grown InP on GaAs substrates by a more· than two order of magnitude reduction in deep level 

concentration and an improvement in reverse bias leakage current by a factor of - 20 (refs. 1-2). These 

results make plasma hydrogenation an extremely promising technique for achieving high efficiency large 

area and light weight heteroepitaxial InP solar cells for space applications. In this work we investigate the 

carrier trapping process by dislocations in heteroepitaxial InP/GaAs and the role of hydrogen passivation 

on this process. It is shown that the charge trapping kinetics of dislocations after hydrogen passivation are 

significantly altered, approaching point defect-like behavior consistent with a transformation from a high 

concentration of dislocation-related defect bands within the InP bandgap to a low concentration of individual 

deep levels after hydrogen passivation. It is further shown that the "apparent" activation energies of 

dislocation related deep levels, before and after passivation, reduce by - 70 meV as DL TS fill pulse times 

are increased from 1 usec. to 1 msec. A model is proposed which explains these effects based on a 

reduction of Coulombic interaction between individual core sites along the dislocation cores by hydrogen 

incorporation. Knowledge of the trapping properties in these specific structures is important to develop 

optimum, low loss heteroepitaxial InP cells. 

INTRODUCTION 

Dislocations within InP layers grown on lattice mismatched substrates such as GaAs, Si and Ge 

are currently a major limitation for the development of efficient heteroepitaxial InP solar cells for space 

applications. The - 8% mismatch in lattice constant for InP/Si, and 4% for both InP/Ge and InP/GaAs, 

typically result in threading dislocation densities in the range 1-1 Ox 108 cm2 in the InP layers which has 

severely reduced heteroepitaxiallnP cell efficiencies thus far (ref. 3). This has prompted investigations of 

a number of approaches to reduce threading dislocation density toward 1 x1 05 cm2

, the theoretically 

predicted value necessary to achieve efficiencies comparable to homoepitaxial InP/inP (ref.3) _ These 

approaches include the use of compositionally graded buffer layers, thermally cycled growth and hydrogen 

passivation (refs. 1-5)_ 

For the case of hydrogen passivation, where the focus is on reducing the electrical activity of 

dislocations rather than total dislocation concentration, it has been shown that exposure of p-lnP grown on 

GaAs substrates to a hydrogen plasma at 250°C reduces the dislocation related deep level concentration 

from - 6x1014 cm3 to - 3x1012 cm-3 in the InP layer (ref. 1)_ This passivation was found to be stable up 

to - 550°C, which combined with dopant reactivation occurring at 380 °C, opens a 170°C post-passivation 

processing window for cell completion_ Furthermore, it was shown that reverse leakage currents of 

heteroepitaxial InP diodes on GaAs were significantly improved as the result of dislocation passivation_ 

These results are summarized in Figures 1 and 2. 

1 Work supported by NASA under grant no_ NAG-1461 . 

401 

NCl PAGE BLANK OT Fl EO 

________ .J 
L 

In addition to reducing the concentration of deep levels resulting from dislocations, hydrogen 

incorporation is found to change the mechanisms of the actual carrier trapping process. Knowledge of the 

dislocation trapping properties is critical to understand current transport characteristics of heteroepitaxial 

cells and must be accounted for in heteroepitaxy-specific cell design optimization. In this paper we present 

a study of the charge trapping properties of dislocations in p-type InP grown by MOCVD on GaAs 

substrates, both prior to and after hydrogen passivation. 

EXPERIMENTAL 

InP was grown on (100) GaAs substrates by low pressure MOCVD. Growth details have been 

reported elsewhere (ref. 1). The InP structure consisted of a 21lm thick layer of Zn doped p-type (1x10'7 

cm':l) InP grown on a 0.5 Ilm thick n+ InP layer on n+ GaAs substrates. Ohmic contacts were formed by 

electron beam evaporation and patterning of Au/Zn/Au and Ni/Ge/Au on the p-type InP and n-type GaAs 

substrate, respectively, followed by a 3800 C anneal in N2. This device structure, shown in Figure 3, allows 

us to probe the depletion region as measured from the buried junction and up toward the InP surface. 

Hydrogenation was performed in a Technics Planar Etch II 30 kHz parallel plate plasma reactor. 

Hydrogen exposure was performed using a 30 sccm flow rate of semiconductor grade H2 , chamber 

pressure of 530 mTorr, power density of 0.08 W/cm2 and a substrate temperature of 2500 C. Hydrogen 

exposure time was either 1.5 or 2 hours, as noted. Immediately prior to hydrogen exposure, all samples 

were capped with a thin, hydrogen-permeable, - 20 nm layer of SiN. to prevent surface degradation by 

preferential loss of phosphorous during hydrogenation. The nitride layer was etched off in dilute HF before 

metallization. DL TS measurements were performed using a Biorad DL4602 DL TS spectrometer, 

respectively based on the double boxcar data acquisition method. Fill pulse times were controlled using 

the Biorad pulse timing circuit and by an external pulse generator. 

Note that all data of hydrogenated samples discussed in this paper received a post hydrogenation 

dopant reactivation anneal for 380 °C for 5 minutes. 

RESULTS AND DISCUSSION 

Figure 4 shows the DL TS spectra for non-hydrogenated heteroepitaxial InP measured at a rate 

window of 1000/sec. as a function of fill pulse time. The fill pulse time dependence of the three main peaks 

(T1, T2 and T3) is more quantitatively depicted in Figure Sa (before hydrogenation) and 5b (after 

hydrogenation) where it is seen that the peak height variation of T1 and T2 display a linear relationship with 

the natural logarithm of the fill pulse time. Such behavior is consistent with earlier reports for charge 

capture by extended defects and dislocations in other semiconductors which has been described by a rate 

equation given as (ref. 6) 

(1 ) 

where cn,p are capture rates for electrons and holes, en,p are emission rates for electrons and holes, PT is 

the concentration of traps filled with holes, NT is the total trap concentration, and <I>(t) is a time dependent 

potential barrier associated with the charging of a dislocation. Note that this differs from the usual point 

defect rate equation only by the presence of the exponential term which acts to slow the trap filling rate . 

The solution of this equation, after some simplification and in terms of the trap fill pulse time, tp' used in 

the DLTS experiment is given by (ref. 6) 

(2) 

The term 't in equation (2) is proportional to the height of this potential barrier and Pr is proportional to 6.C/C 

as given in Figure 5. Note that in each of these plots of Figure 5, T1 and T2 follow the logarithmic behavior 

402 
- - - - - - - - - ---

and extended fill pulse time dependence predicted for dislocations, whereas T3 follows the usual point 

defect behavior. In fact, we have earlier shown that T3 was indeed associated with a point defect since 

it was the only level evident in the DL TS spectra of homoepitaxiallnP/lnP structures (ref. 7). By comparing 

Figures Sa and Sb, it can be seen that hydrogen passivation results in a significant change in the slope of 

the plotted data. Calculations have shown that the slope for the T1 and T2 curves reduce by approximately 

two orders of magnitude after hydrogenation, indicating that the trap state density is decreased as 

expected, and that the barrier to carrier trapping, CP(t) , has been lowered. 

While the logarithmic capture kinetics is one characteristic of the unusual trapping properties of 

dislocations in heteroepitaxial InP, it is also clear from Figure 4 that T1 and T2 broaden on the low 

temperature side of the DL TS peaks as the fill pulse time is increased, resulting in a shift toward lower 

temperatures of the peak maximum with fill pulse time. This is shown more quantitatively for T2 in Figure 

6 where the FWHMITp is plotted against the fill pulse time (Tp is the temperature at which the DL TS peak 

is maximum). It can be shown that for a single deep level, FWHMITp = 0.104 for a rate window of 

1000/sec., which should not vary with fill pulse time. As seen in the figure, T2 is clearly broadened 

compared to this value and furthermore, varies with fill pulse time until it saturates at ~ - 0.1 msec. This 

behavior can be understood if one first considers that dislocations are generally thought to cause a band 

of deep levels, rather than non-interacting isolated levels, due to wavefunction overlap of closely spaced 

dislocation core states. In this case, only the lower energy states within the band are filled for very short 

fill pulse times. As fill pulse time increases, higher energy states can trap carriers, eventually saturating 

as the band becomes filled. It can be shown that the higher energy states for holes (this corresponds to 

states closer to the valence band), which result in lower activation energies (Er -Ev)' will always contribute 

DL TS peaks at lower temperatures for the same rate window and cross section, compared to lower energy 

states (higher activation energies). This would explain why the DL TS peaks broaden only on their low 

temperature sides. It should be noted here that a similar analysis was performed for T1 . A complete 

analysis will be presented elsewhere. 

Figure 6 also includes FWHMITp data for hydrogen passivated samples. Note that in comparison 

to non-hydrogenated data, the FWHMITp characteristic of peak T2 after hydrogen passivation is much 

closer to that of an ideal point defect, although it is still somewhat broadened. This is likely the result of 

a reduction in the active core site density which causes an increase in the average active core site spacing 

and a change in the dislocation density of states. 

The result of the peak broadening and shifting for DL TS interpretation is most striking in Figure 7, 

which shows the measured trap activation energy for T1, T2, and T3 as a function of fill pulse time following 

the usual Arrhenius method. First consider the non-hydrogenated data. As can be seen, the measured 

activation energy shifts by approximately 80 meV for T1, from - 880 eV for a 1 usec. fill pulse time, to -

800 eV for a 10 msec. fill pulse time. T2 shows a shift of similar magnitude. In contrast, the point defect 

level, T3, shows no such dependence, as would be expected. For the complicated case of dislocations, 

the Arrhenius method for DL TS activation energy actually yields the weighted average energy of the filled 

states within the dislocation energy band that contribute to carrier emission. For short fill pulse times, only 

those states farthest from the valence band are filled with holes, resulting in larger activation energies. One 

would therefore expect a more "point defect-like" behavior due to reduced dislocation charging. This is 

corroborated by Figure 6 which shows less peak broadening for short fill pulse times. For longer fill pulse 

times, the band fills downward with higher energy holes (toward the valence band edge). Therefore the 

average energy of the re-emitted holes is higher and the measured average activation energy decreases, 

as observed. Eventually, for longer fill pulse times, the band will saturate with trapped holes. At this point, 

the measured activation energy is indicative of the weighted average over the entire band. Note that the 

variation of activation energy in Figure 7 is completely consistent with the FWHMIT p variation shown in 

Figure 6. These results suggest that care should be taken when comparing DL TS measured activation 

energies for dislocations, with particular attention to the fill pulse time used in the experiment. 

The effect of hydrogen passivation on the activation energy behavior is also shown in Figure 7 for 

each DL TS peak. Again, negligible change in activation energy is observed for the point defect level, T3, 

403 
while a significant change of activation energy with fill pulse time is still evident for both T1 and T2. 

However, from the figure it is clear that the activation energies of T1 and T2 are consistently higher for all 

fill pulse times after hydrogenation. It should be noted that the shape of the activation energy dependence 

on fill pulse time also changes after hydrogen passivation. This is likely a result of reduced density of 

states at certain energies due to passivation, which allows a faster change, or response, to fill pulse time. 

This suggest that such a plot of activation energy versus fill pulse time may provide information about the 

actual shape of the dislocation density of states. This is currently being looked at in more detail. 

CONCLUSIONS 

Hydrogen passivation by plasma hydrogenation has been shown to be effective in reducing the 

electrical activity of dislocations in p-type InP grown on GaAs substrates. Deep level concentrations were 

reduced by more than 2 orders of magnitude and diode leakage currents were suppressed. These 

beneficial effects were found to be stable up to a post hydrogenation annealing temperature of 550 °C. 

Detailed analysis of the dislocation trapping kinetics prior to hydrogenation showed the expected logarithmic 

trapping kinetics as reported for dislocations in other materials. Furthermore, it was shown that this was 

related to the observed DL TS peak broadening toward lower temperatures with fill pulse time and a large 

dependence of measured trap activation energy on fill pulse time. It was found that hydrogen passivation 

Significantly altered the fundamental trapping properties of the dislocations, resulting in a decrease in the 

broadening of the dislocation related DL TS peaks in addition to a shift in the average energy of the 

dislocation density of states away from the valence band. These effects suggest that hydrogen passivation 

reduces the interaction between adjacent dislocation core sites and results in a trapping behavior that 

approaches that expected for point defects. 

REFERENCES 

1. B. Chatte~ee, S.A. Ringel, R. Sieg, R. Hoffman and I. Weinberg, "Hydrogen Passivation of Dislocations 

in InP on GaAs Heterostructures," Appl. Phys. Lett., vol. 65, pp. 1-3, 1994. 

2. B. Chatterjee, SA Ringel, R. Sieg, R. Hoffman and I. Weinberg, "Deep Level Characterization and 

Passivation in HeteroepitaxiallnP," Mat. Res. Soc. Symp. Proc., vol. 325, pp. 125-130, 1994. 

3. R.K. Jain and D.J. Flood, "Influence of the Dislocation Density on the Performance of Heteroepitaxial 

Indium Phosphide Solar Cells," IEEE Trans. Electron. Dev., vol. 40, pp. 128-1934, 1993. 

4. C.J. Keavney, S.M. Vernon, V.E. Haven, S.J. Wojtczuk and M.M. AI-Jassim, "Fabrication of n+/p InP 

Solar Cells on Si Substrates," Appl. Phys. Lett., vol. 54, pp. 1139-1141, 1989. 

5. I. Weinberg, C.K. Swartz and D.J. Brinker, "HeteroepitaxiallnP Solar Cells on Si and GaAs Substrates," 

Tech. Digest Int. PVSEC-5, pp. 81-84,1990. 

6. P. Omling, E.R. Weber, L. Montelius, H. Alexander, and J. Michel, "Electrical Properties of Dislocations 

and Point Defects in Plastically Deformed Silicon," Phys. Rev. B., vol. 32, pp. 6571 -6581, 1985. 

7. B. Chatterjee, SA Ringel, R. Sieg, R. Hoffman and I. Weinberg, "Identification and Passivation of 

Mismatch Related Deep Levels in Heteroepitaxial InP on GaAs Substrates," Proc. 184th Meeting 

Electrochem. Soc., New Orleans, 1993 (in press). 

404 
I 

I 

I 

I 

l 

I 

I 

-10 ....... -----------.,----.-

1.3.4.5.6 

:f 

1: As-grown 

2 

2 

3~ 2: Hyd. 2 hrs. Unannealed 

4 

c ....... l) < t 

::l .=!: 

u 

-24 . 

-1.5 

, 

VR a 

APPLIED BiAS [\1] 

3: Hyd. 2 hrs. Annealed 400C 

4: Hyd. 2 hrs. Annealed 500c 

5: Hyd. 2 hrs. Annealed 580C 

6: Hyd. 2 hrs. Annealed 600C 

1.5 

Figure 1. Current-voltage characteristics of p-lnP/n+-lnP diode grown on GaAs, measured before and after 

hydrogenation and after various post-hydrogenation anneals at the temperatures indicated. 

I 

[ 

= t :2 r 

~ u 

< 

100 

150 200 

T3 

150 

T2 

200 

(a) 

A AS GR OW!" o B 600 ( 

( 5800 ( 

D 500 0 ( 

250 30(J 

TEMPERATURE IK J 

3~ O 

Figure 2. Effect of hydrogenation and post-hydrogen 

anneals on DL TS spectra for InP/GaAs, measured at 

a rate window of 1000/sec. The inset measured at 

SO/sec rate window, clearly shows trap T3, which 

is prominent at higher rate windows. 

405 

Au{Zn/Au 

frlnP 

17 ~ 1x10 em 

n-lnP 

n + GaAs substrate 

Au/Ge/Ni 

1.9 urn 

0.5 urn 

Figure 3. Device structure 

used for this study. 
L _ 

NON.II\'IlIW(;ENATD) (IOUU/St:(·) 

600 

5 m) ... ·c 

~ TZ I m)l'C 

~UU I'''C 

0 100 I',ec Tt 

<l IIU ",C( 

400 6U ",rc 

~ ","( 

4U ",ec 

1U IlSC( 

10l'sec 

200 

o~~~~~~~LU~~~LW~~~wU~~~ui~~ 

100 150 200 250 300 350 

TEMPERATURE [KJ 

Figure 4. DLTS spectra measured at 1000/sec. rate window for the different fill pulse times as shown. 

6.5E+14 

NON-HYD • "? • • • ~ 5.3E+14 

~ • • 

• 

0 4.1E+14 

z 

z < 29E+14 . • T1 

u • T2 

-- 1.7E+14 U •• <l • • T3 

C\l • 5E+13 

1.00E-07 1.00E-06 1.00E-05 1.00E-04 1.00E-03 1.00E-02 

Tfill[SEC) 

Figure 5(a). Capture characteristics for traps T1, T2 and T3 before hydrogenation. T1 and T2 exhibrt 

ex1ended dependence on fill pulse time that is characteristic of dislocations and follows eqn. (2). 

406 
I 

I 

I 

I 

I 

I 

I 

I 

I 

HYD 1.5HRS 

4E+13 /: C',' • • • ::;! • • • , 

Q . 0 //. z 

, 2E+13 

« • T1 

z 

• • T2 

0 

...... • • • 

0 .- - T3 

<l 

C\I 5E+12 1-- ·--- 1 

100E-07 1.00E-06 1.00E-05 1.00E-04 1.00E-03 1.00E-02 1.00E-01 

Tfill[SEC) 

Figure 5(b). Capture characteristics for traps T1, T2 and T3 after hydrogenation. Note the reduction in 

slope for T1 and T2 compared to figure 5(a). 

o NON-HYD. • HYD. 1.5 HRS. I 

0.125 0 0 0 0 

0.12 0 0 

T2 ~ 0 i 0.115 0 

• • • • 

~ 0.11 • • • • 

0.105 POINT DEFECT 

0.1 

1.00E-05 1.00E-04 1.00E-03 1.00E-02 

nUl [SEC] 

Figure 6. FWHMfTp data for T2, as a function of fill pulse time measured before and after hydrogenation. 

The tine at FWHMfTp = 0.104 is calculated for an ideal point defect. 

407 

____ _ J 
910 

~ oS 875 • • • 

>- • o T1, NON-HYD t!) 

a: • z • T1 , 1,SHRS HYD w 840 0 0 

Z • 

0 0 

~ 0 

80S 0 I- • u 0 • • • 

< a 0 0 0 

770 

U)OE-<l6 1.00E-OS 00E~ 1.00E-03 1.00E-02 

Tfill (SEC] 

• <465 • r <4SO 

• o T2, NON-HYD 

>- 0 0 • • T2, HYD 1.SHR 

t!) <435 a: 0 

w 0 z 

w <420 • • z 0 

0 405 • 

~ 0 • • • 

390 

u b 0 

< 375 0 

0 0 

360 

1.00E-<l6 1.00E-OS 00E~ 1.00E-03 1.00E-02 

Tfill (SEC] 

260 o T3, NON-HYD. 

~ 256 ~ 

• T3, HYD. 1.S HRS 

t; 

a: 

w 2S2 • 

z 0 

w e II 

z 

0 2<48 

~ 244 u C < 

2.0 

1.00E-<l6 8.00E-<l6 1.10E-OS 1.60E-OS 2. 1 OE-OS 

Tfill (SEC] 

Figure 7, Variation of DL TS activation energy with fill pulse time , measured by the Arrhenius method for 

as-grown and hydrogenated samples for (a) T1, (b) T2, and (c) T3 . 

408 
,- - - - - - - ---

3Qt()PI 

N95- 20543 

NUCLEATION, GROWTH, AND STRAIN RELAXATION OF LATTICE-MISMATCHED III-V d'trSEMICONDUCTOR EPITAXIAL LAYERS 1 

R.E. Welser and L.J. Guido 

Yale University 

New Haven, Connecticut 

Summary 

We have investigated the early stages of evolution of highly strained 2-D InAs layers and 3-D InAs 

islands grown by metal-organic chemical vapor deposition (MOCVD) on (100) and (111)8 GaAs 

substrates. The InAs epilayer / GaAs substrate combination has been chosen because the latticemismatch is severe (- 7.2%), yet these materials are otherwise very similar. 8y examining InAs-on-GaAs 

composites instead of the more common InxGa1_xAs alloy, we remove an additional degree of freedom (x) 

and thereby simplify data interpretation. A matrix of experiments is described in which the MOCVD growth 

parameters -- susceptor temperature, TMln flux, and ASH3 flux -- have been varied over a wide range. 

Scanning electron microscopy, atomic force microscopy, transmission electron microscopy, and electron 

microprobe analYSis have been employed to observe the thin film surface morphology. In the case of 3-D 

growth, we have extracted activation energies and power-dependent exponents that characterize the 

nucleation process. As a consequence, optimized growth conditions have been identified for depositing 

- 250 A thick (100) and (111)8 oriented InAs layers with relatively smooth surfaces. Together with 

preliminary data on the strain relaxation of these layers, the above results on the evolution of thin InAs 

films indicate that the (111)8 orientation is particularly promising for yielding lattice-mismatched films that 

are fully relaxed with only misfit dislocations at the epilayer / substrate interface. 

Introduction 

In comparison with Si and Ge, III-V semiconductor alloys are well-suited for space photovoltaic 

applications because of improved radiation hardness and the potential for heterostructure devices with 

high power conversion efficiency. Unfortunately, the choice of starting substrate for the subsequent 

growth of I II-V epitaxial layers is limited by the practical requirements of a large strength-to-weight ratio and 

low cost. The single crystal materials that satisfy these constraints -- Si, Ge, and, to a lesser degree, GaAs 

-- also possess lattice constants that differ significantly from those of the III-V semiconductor alloys of 

interest (fig. 1). Consequently, the lattice-mismatched epitaxial layer / substrate composite undergoes an 

elastic deformation that (1) causes a transformation from the desired 2-D layer-by-/ayer growth mode to a 

3-D island mode and (2) provides the driving force for dislocation generation and propagation. The result 

is a rough surface and a high density of defects in the overlying epitaxial layers (fig. 2a). What is needed 

instead for device applications is a smooth 2-D surface with misfit dislocations at the epitaxial layer / 

substrate interface which relieve the strain caused by the lattice-mismatch, but do not thread the epitaxial 

layer and consequently degrade the device quality (fig. 2b). 

Several different mUlti-step fabrication schemes have been attempted to improve the structural 

properties of lattice-mismatched epitaxial layers. In fact, the practices of using either a low-temperature 

buffer layer or a vicinal substrate to initiate growth, and of incorporating a compositionally graded layer or a 

superlattice dislocation filter between the substrate and active region, have all met with limited success 

(ret. 1 -4). Post-growth annealing techniques have also been explored but have proven less effective 

(ref. 5). Despite some encouraging results. even these multi-step methods have failed to yield defect 

densities low enough for high-performance minority carrier devices « 1 04 cm-2). 

1 The authors gratefully acknowledge the financial support of the National Science Foundation via the 

Presidential Faculty Fellowship Program (Grant # ECS-9253760) and of the NASA Graduate Student 

Researchers Program (Grant # NGT-50832). 

409 
l 

__ 

Recent reports suggest that epitaxial growth conditions may strongly influence the kinetics of 

strain relaxation (ref. 6). Although the minimum energy configuration of a highly strained material is 

characterized by a rough, three-dimensional surface morphology, the exact nature of this equilibrium 

configuration, or even whether or not it develops at aH, also depends on kinetic limitations. By implication, 

the efficacy of a particular buffer layer technique depends upon the experimental conditions during each 

step of the deposition process. In this paper, we describe the influence of the kinetic pathways of 

nucleation, growth, and strain relaxation on the surface morphology and the defect structure of highlystrained epitaxial thin films. 

The Initial Stages of Deposition 

While a comprehensive theoretical description of III-V heteroepitaxy which incorporates the 

complexities of MOCVD is lacking at present, a generic model of atomic deposition provides a useful 

reference (ref. 7). Figure 3 illustrates the initial stages of film deposition from single atoms which arrive at 

the substrate at a rate R and diffuse across the surface with diffusion coefficient D. The surface density of 

these single atoms (n1) is then determined primarily by the competition between the arrival rate and the 

rate at which single atoms are lost to re-evaporation and incorporation into a growing island or pre-existing 

step edge. During the nucleation stage, single atoms can also couple to other atoms (binding energy Ej) 

to form sub-critical clusters of atoms whose density (nj) is inversely proportional to the number of atoms in 

the cluster. Above a critical size i*, the sub-critical cluster is more likely to grow into a stable island than to 

diSintegrate. The density (nx) and the geometry of these growing islands are the two most important 

parameters determining the surface morphology and defect structure of thicker epitaxial layers. 

The three traditional modes of growth (ref. 8) can be described in terms of the island width-toheight ratio (w / h). In the limit of large w / h, the islands are two-dimensional even at the point of 

coalescence, and growth proceeds in a layer-by-Iayer fashion (Frank-van der Merwe). In the limit of small 

w / h, on the other hand, islands grow in a three-dimensional or islanding mode (Volmer-Weber). If the 

w / h ratio changes from a large to a small value as the islands evolve, then layer-by-Iayer growth is followed 

by islanding (Stranski-Krastanov). Growth may also proceed in a more complicated fashion for 

intermediate values of w / h. Specifically, as the island density increases and island coalescence begins, 

there is a range of w / h values which produces films that are nearly indistinguishable from those grown via 

the layer-by-Iayer mode. 

As stated above, the density of growing islands is proportional to the density of critical clusters, 

which is inversely proportional to the number of atoms in the critical cluster and directly proportional to the 

density of single atoms and their binding energy. The processes depicted in Figure 3 can be 

quantitatively described by rate equations which govern the density of single atoms, sub-critical clusters, 

and stable clusters (growing islands) on the substrate surface and highlight the interrelation between 

nucleation and growth. Solving this set of equations in a self-consistent manner yields a nucleus 

saturation value (Ns) 

where No is the density of absorption sites, Ea is an activation energy, and the exponent p is proportional 

to the critical cluster size (ref. 7). 

In many cases, the stress in thick lattice-mismatched films is large enough to drive dislocation 

nucleation and propagation (ref. 2). However, the thickness at which dislocations nucleate and propagate 

in 2-D films is very sensitive to the kinetic pathways of epitaxial growth (ref. 9) and to heterogeneities in 

material properties (ref. 10). For example, recent work has shown that islands can form before the onset of 

dislocation generation (ref. 11). As these coherent islands grow, there is an increase in the driving force 

for dislocation generation. Thus, there is a direct connection between the evolving surface morphology 

and defect generation in highly-strained epitaxial layers. 

410 
Experimental Approach and Summary of Results 

The InAs samples have been deposited in a horizontal-geometry MOCVD reactor at a base 

pressure of 100 torr, with an H2 carrier gas flow of 12 slm and TM In and AsH3 sources. All growths have 

been performed simultaneously on (100) and (111)8 Si-doped GaAs substrates. After a standard oxide 

desorption process, the substrate temperature (T s) and ASH3 flow (fAsH3) are adjusted to the desired 

values, and the growth of InAs is initiated directly on the substrates by switching in a pre-stabilized TMln 

flow (fmln). The surface morphology has been observed by Nomarski optical microscopy (NOM), 

scanning electron microscopy (SEM), and atomic force microscopy (AFM). Transmission-electron 

microscopy (TEM) has also been performed to determine structural quality of the films. The constancy of 

the island density with time indicates that, except for the eventual decrease in density because of 

coalescence, the measured values are indeed the nucleus saturation densities. The pertinent results of 

these experiments are summarized in the following discussion. 

The impact of substrate temperature, TMln flow, and AsH3 flow on InAs nucleation has been 

examined in a 23 factorial experiment on nominally 250-A thick films. This approach is designed to 

efficiently evaluate the prinCipal and interactive effects of a large set of variables (ref. 12). There are 

several noteworthy trends evidenced in the data of Table I. First, the island densny on the (100) substrate 

is approximately 100x larger than that for the (111)8 case. In addition, the island density on both 

substrates increases with either a decrease in Ts or an increase in fmln, as shown in more detail in Figures 

4a and 4b. An increase in fAsH3 also has a modest effect, but with the opposite sense for the two 

orientations. Next, the island geometry is completely different in the two cases (fig. 5a and 5b). For the 

(100) substrate, AFM reveals pyramids with a rectangular base aligned to the <011> crystallographic 

directions and side-walls best fitted by {111} planes. In contrast, the islands on the (111)8 substrate are 

flat polygons with the base aligned to <211> directions, and extremely abrupt side-walls which are most 

likely {011} cleavage planes. The w / h values of the islands are also very different on the two substrate 

orientations, with a value of - 2 on the (100) substrate and - 50 for the (111)8 case. The exponential 

dependence of Ns with T 5 and the power dependence of Ns with fTMln agrees qualitatively with the 

conventional picture of nucleation. The lower densities on the (111)8 substrate can be explained by 

assuming that the critical cluster size is larger than on the (100) substrate. However, the effect of AsH3 

flow on the nucleus saturation density cannot be accounted for via this simple single-species nucleation 

model. 

Using these results, the surface morphology can be transformed from a 3-D into a 2-D surface. 

Figure 6 shows that lowering the temperature to 525 °C, increasing the TMln flow to 720 sccm, and 

increasing the ASH3 flow to 500 sccm leads to a nearly featureless 250 A thick InAs layer on the (111)8 

substrate. Conditions were also found for a similar occurrence on (100) substrates. Figure 7 shows that 

the morphology transition on the (100) substrate is achieved by lowering the temperature to 475 oC, 

increasing the TMln flow to 950 seem, and lowering the ASH3 flow to 82 seem. While the films shown in 6b 

and 7b are both "specular," a more quantitative analysis has been obtained via AFM. The surface 

roughness of the strained (100) InAs layer « 50 A) is higher than that of (100) GaAs « 1 0 A ) grown under 

similar conditions. However, in comparison to a (111)8 GaAs layer, the surface roughness of the strained 

InAs is 30x lower. The flat surface and large w / h ratiO of InAs islands on (111)8 GaAs substrates is 

consistent with the observation that 2-D films form more readily (at higher temperature and lower TMln 

flow) than on (100) substrates. The small w / h ratio of InAs islands on (100) substrates also accounts for 

the relatively large surface roughness. While TEM analysis indicates that both films are fully relaxed, the 

(100) film exhibits a large density of threading dislocations. However, for the (111)8 InAs film, preliminary 

analysis reveals misfit dislocations only; no threading dislocations are observed in cross-sectional TEM 

images. 

411 

- - -.----.---.~---- ------ - - - --- --~ ----- --- --' 
Conclusions 

Using InAs-on-GaAs as a model high-strain system and a factorial design of experiments 

approach, we have demonstrated that the thin-film surface morphology is quite sensitive to epitaxial 

growth parameters. The temperature and TMln effects can be described, at least qualitatively, by 

traditional one-species nucleation theory. However, a more complex model is needed to account for the 

effects of AsH3 flow. Two-dimensional morphologies have been demonstrated on both (100) and (111)8 

substrate orientations; however, the (111)B orientation is more promising because only misfit dislocations 

are present at the epilayer / substrate interface. 

References 

01. M. S. Abrahams, L. R. Weisberg, G. J. Busocchi, and J. Blanc, J. Mater. Sci. 4, 223 (1969). 

02 . J. W. Matthews and A. E. Blakeselee, J. Cryst. Growth 32, 265 (1976). 

03. M. Akiyama, Y. Kawarada, and K. Kaminichi, J. Cryst. Growth 68, 21 (1984) . 

04 . R. Fisher, D. Neumann, D. Zibel, H. Morkoc, C. Choi, and N. Ostuka, Appl. Phys. Lett. 48, 1223 

(1986). 

05. J. W. Lee, H. Shichijo, H. L. Tsai, and R. J. Matyi, Appl. Phys. Lett. 50, 31 (1987). 

06. E. A. Fitzgerald, V-H. Xie, D. Monroi, P. J. Silverman, J. M. Kuo, A. R. Kortan, F. A. Thiel, and B. E. 

Weir, J. Vac. Sci. Technol. B 10, 1807 (1992). 

07 . J. A. Venables, G. D. T. Spiller, and M. Hanbucken, Rep. Prog. Phys. 47, 399 (1984). 

08. E. Bauer, Z. Kristallogr. 110, 372 (1958). 

09 . B. W. Dodson and J. Y. Tsao, Appl. Phys. Lett. 51,1325 (1987). 

10. E. A. Fitzgerald, J. Vac. Sci. Technol. 87, 782 (1989). 

11 . D. J. Eaglesham and M. Cerullo, Phys. Rev. Lett. 16, 1943 (1990). 

12 . G. E. D. Box, W. G. Hunter, and J. S. Hunter, Statistics for Experimenters, John Wiley & Sons, 

1977. 

41 2 
Run 

Number 

1 

2 

3 

4 

5 

6 

7 

8 

~ 

:> 

4) 

'-" 

~ Co:) 

"C 

a 

..0 I 

>- bl) 

'""' 4) 

c:: 

~ 

Table I : Matrix of Growths of InAs on GaAs - A 23 Factorial Design 

Growth H2 Flow In 

Temp,(°k} TMln (seem) W!! 

525 140 100 

600 140 100 

525 720 100 

600 720 100 

525 140 400 

600 140 400 

525 720 400 

600 720 400 

3.0 

2.5 AlP 

2.0 

1.5 

0 1.0 Si 

0.5 

0.0 

5.4 5.5 5.6 5.7 5.8 

Lattice Constant (A) 

Island Density (cm-2) 

(100) 

Substrate 

5.6 x 108 

1.4 x 108 

1.0x109 

1.7 x 108 

4.3 x 108 

4.9 x 107 

6.7x 108 

1.0 x 108 

• Direct 

o Indirect 

InAs 

(111 ) 

Substrate 

coalesced 

1.9 x 106 

coalesced 

4.5x106 

coalesced 

2.7x106 

coalesced 

7.8 x 106 

5.9 6.0 6.1 

Figure - 1 Plot of energy-gap vs. lattice constant for elemental and III-V compound semiconductors and 

alloys. Note that InP has a lattice constant that is 3.8 % larger than that of GaAs and 8.1 % larger than Si. 

Also note that the (AlxGa1-x)O.65lno.3SAs alloy with a lattice constant of 5.8 A covers an energy-gap range 

from 1 eV to 2 eV, which is the optimum energy-gap spectrum for a multijunction solar cell configuration. 

413 
Epitaxial 

Layer 

Substrate 

Epitaxial 

Layer 

Substrate 

(a) 

(b) 

Figure· 2 (a) Relaxed epitaxial layer with a rough surface morphology and high threading dislocation 

density. (b) Relaxed epitaxial layer wijh only misfit dislocations. 

sub-critical 

cluster 

w 

Figure· 3 A stepped surface with a sub-monolayer deposit consisting of single atoms, sub-critical 

clusters of atoms, and growing islands. Single atoms arrive at the substrate at a rate R, move across the 

surface with diffusion coefficient D, and can either re-evaporate at a rate n1 I 'ta, form sub-critical clusters of 

atoms which are finally transformed into growing islands, or incorporate directly into a pre-existing step 

edge. 

414 

l __________ _ __ --l 
rI 

I 

E = 1.36 eV ..r..... a __ ~ _ - J:' 

~--~ ~ 

12.5 13 

---(100) 

--o--(l11)B 

13.5 14 

l/kT (eV- 1) 

(a) 

p = .93 -9 ---

_ ---9-

100 

• (100) 

- -0- - (111)B 

14.5 15 15.5 

p = .71 

1000 

H2 Flow in TMln (seem) 

(b) 

Figure - 4 Nucleus saturation density as a function of temperature in (a), and TMln flow in (b). 

415 

, 

____ J 
l_ 

[011] 

t 

--1 [011) 

(a) 

- , -......., .... --.-,;", 

,/ 

~: 

<211> 

(b) 

Figure - 5 InAs is )I1ds on a (100) orier,ted G'1 l1,s ubstrCll~ in (a) , and a (111 )6 substrate in (b). Growth 

conditions are T = 600 oC , fTMln = 140 ~cen I ,\,,11 3 = 100 seem, tavg - 100 A. 

416 

I 
I--

I 

10 ~m 

10 1 m 

Figure·6 Optical micrographs for the (11 1)8 GaJ\s substrate d~pieting til e tral1s ition from a 3-D 

morphology in (a) to a nominally 2-D InAs surface In (b) Growtn conditions are T = 600 oC, fTMln = 140 

seem, fAsH3 == 100 seem , tavg - 250 A for (a) . and T = 525 oC, fTMln = 720 seem, fAsH3 == 500 seem, tav9 -

250 A for (b). 

417 

I 

~~ ___ ~ ______ J 
 

_ 

I eli 10 pm 

(b) 

Figure - 7 Optical micrographs for the (1 OO) GaAs substr?le depicting Ire transition trom a 3-D 

morphology in (a) to a nominally 2-D suriace in (b) Growtrl conditions are r = 525 oC , fTMln = 720 scem, 

fAsH3 = 125 seem. tavg - 250 A for (a). and T = 475 °C. fTMln = 950 seem. fAsH3 = 82 .5 secm, tavg - 250 

A for (b) 

418 

I 

I 
SUMMARY OF WORKSHOP ON InP: STATUS AND PROSPECTS 

R.J. Walters 

U.S. Naval Research Laboratory 

Washington, DC 

and 

I. Weinberg 

NASA Lewis Research Center 

Cleveland, Ohio 

This paper presents a summary of the workshop on InP solar cells. The overall purpose of this 

workshop was to: 

1) determine the primary objective of the present InP research programs 

2) establish the immediate prospects for use of InP solar cells 

3) state the current status of the various InP research projects 

4) identify the current major problem areas in the development of InP cell technology 

This report address each of these topics in turn. 

1) OBJECTIVES 

The primary objective of most of the programs in InP solar cells is the development of the most 

radiation hard solar cell technology. In the workshop, it was generally agreed that the goal is a cell 

which displays high radiation tolerance in a radiation environment equivalent to a 1 MeV electron fluence 

of about 1016 cm·

2 • Furthermore, it is desired that the radiation response of the cell be essentially flat out 

to this fluence - i.e. that the power output of the cell not decrease from its beginning of life (BOL) value 

in this radiation environment. 

It was also agreed in the workshop that the manufacturability of InP solar cells needs to be 

improved. In particular, since InP wafers are relatively dense and brittle, alternative substrates need to 

be developed. Research on hetero-epitaxial InP cells grown on Si, Ge, and GaAs substrates is currently 

underway. The ultimate goal is to develop hetero-epitaxial InP solar cells using a cheap, strong, and 

lightweight substrate. 

2) PROSPECTS 

The prospects for use of InP solar cells are primarily in high radiation earth orbiting satellites. 

There has been an expressed need in the Navy and in commercial interests (particularly in the 

communication industry) to fly satellites in orbits where the equivalent 1 MeV electron fluence for a five 

year mission is greater than 3 x 1015 cm·

2

• Calculations made at the Naval Research Laboratory (NRL) 

and NASA Lewis have shown hetero-epitaxial InP cells grown on Si wafers to be the most cost-effective 

technology for these missions. Also, even low-earth-orbits (LEO), especially polar orbits, can be a 

severe radiation environment due to solar flares. A satellite power system based on radiation hard InP 

419 

j 
solar cells would be relatively insensitive to solar flares. Also, extending geostationary mission lives to 

beyond 10 years has been considered. Such a mission would experience a significant amount radiation. 

Radiation hard InP solar cell technology has the potential to significantly improve the performance of 

these missions. 

Another area where InP solar cell technology has been considered is in the alpha and beta voltaic 

power sources. By there very nature, such devices are very sensitive to radiation effects. The radiation 

hardness of InP seems well suited for this application. 

Although the focus of the present conference is space photovoltaics, the possibility of terrestrial 

applications for loP solar cells was identified. One particular application was that of a concentrator array. 

While no space system has ever used a concentrator array, terrestrial systems make ample use of this 

technology. Modeling results reported by the National Renewable Energy Laboratory (NREL) have 

shown the InP/Gllo.47IIlo.s3As tandem cell to be the best band gap combination for concentrator 

appl ications. 

3) STATUS 

The status of the present research in InP solar cells is best described by summarizing the current 

programs: 

at NRL: 

1) Hetero-epitaxial InP on Si cells 

This is an SBIR program with Spire Corporation. Phase I was recently completed and 

phase II is currently underway. The goal of this project is to produce a large number (> 100) 

2x2 cm cells with BOL efficiencies of 16% (1 sun, AMO, 25 C) which virtually do not degrade 

after an equivalent 1 MeV electron fluence of 1016 cm2

• 

2) InP/GaO•

47Ino.$lAs Tandem Solar Cells 

This program includes NREL for the cell growth and NRL for the cell characterization. 

The current best cell efficiency is 22 % . The program was not funded in FY 94 but has good 

chances for funding in FY 95. The next step in the program is to grow the tandem on a Si 

substrate. 

3) Basic Research - Annealing of Radiation Damage 

NRL has a basic research program studying displacement damage effects in InP solar 

cells. At present, the main research topic is the annealing characteristics of irradiated InP cells. 

at NASA Lewis: 

1) Hetero-epitaxial InP on Si 

NASA is funding Matrix Sciences to grow InP cells on Si substrates. This is a phase two 

contract. 

2) Hetero-epitaxial InP on Ge 

NASA Lewis funded Spire Corporation to grow InP cells on Ge. This was a phase I 

420 

L __________ . ______ . __ . ____ _ 

-I 

I 
rI 

- - -- --

SBIR and has been completed. 

3) Hetero-epitaxial InP on Ge 

NASA Lewis also has an "in-house" program to develop InP cells on Ge substrates. 

4) Hydrogen Passivation 

NASA Lewis is funding Ohio State to study the effects of hydrogen on the dislocations 

which occur in a hetero-epitaxial cell. 

at Space Vacuum Epitaxy Center. University of Houston: 

The research at Space Vacuum Epitaxy Center is developing chemical beam epitaxy as 

a growth technique for photovoltaic devices. As part of this development, these researchers are 

growing InP/Gao.47Iilo.s3As t,mdeITi c~l1s. 

4) PROBLEM AREAS 

The main problem area encountered in the present research is that the heteroepitaxial cells must 

be p+n cells while most cells grown to date have been n+p cells. Therefore, the major research focus 

is in optimizing homojunction p+n cells to the level of the n+p cells. 

Another research focus is the reduction of the deleterious effects of the dislocations which form 

in InP cells grown on Si, Ge, or GaAs substrates. The use of graded and possibly strained layers as an 

intermediating layer between the substrate and the cell active layers to prevent the propagation of the 

dislocations into the active region is being investigated . Also, the research at Ohio state is investigating 

the possibility of passivating the defect levels created by the dislocations with hydrogen. 

SUMMARY 

In general, the workshop concluded that the InP solar cells are being developed as an enabling 

technology which, by virtue of its superior radiation resistance, will allow space flights in high radiation 

orbits which are not possible with Si or GaAs solar cell technology. 

421 

_______ - _. ___ - _________ J 
SYSTEMS, ARRAYS AND APPLICATIONS WORKSHOP 1 

G.J. Pack 

Lockheed Missiles and Space Company 

Sunnyvale, California 

and 

Lothar Gerlach 

ESAIESTEC 

Chelmsford, Essex, United Kingdom 

The charter of this workshop was to evaluate photovoltaic technologies from the broad 

perspective of future mission needs and operational requirements. We were given a 

set of six questions listed in Figure 1 to start the discussion, however; these were 

viewed as sample questions which didn't constrain our deliberations. 

Our primary objective in discussing the provided questions and other items of interest 

to the group was to answer the question: How should NASA spend its scarce space 

power resources? It was clear from the papers presented in the main session and the 

depth of technical talent present in the audience, that there are significantly more good 

projects available than there is funding for. Thus NASA is faced with the problem of 

deciding whether to allocate their resources across a range of projects or to focus on a 

small subset of tasks that are mission enabling. 

Our discussion arrived at two primary findings: 

• NASA management has failed to articulate a vision of where the agency is going. 

• Consequently, the questions listed in figure 1 are irrelevant. 

The failure of management to provide leadership and a consistent direction is 

apparent in the space power arena. Photovoltaics has been the mainstay power 

source for U. S. space missions for the past thirty years. However, for the last decade, 

NASA (and the Air Force) has continually wasted more money on frivolous pursuits of 

nuclear, solar dynamic, and other poorly justified energy sources than they have 

invested in the photovoltaic arena. We must therefor face up to the reality that space 

photovoltaics research has slipped out of the mainstream and into the eddies. 

We have seen the mission emphasis change yearly from missions to Mars, to missions 

to Earth. It has thus become traditional for the research centers to try to fund a little 

something for everybody and keep all options open. The subset of questions posed 

for this group follows that path ranging from interplanetary to near term LEO 

commercial missions and everything in between. Insufficient funding to make 

significant progress in a timely manner renders these efforts irrelevant to the ultimately 

selected mission. 

The R&D technical community cannot expect to redefine the agency. The following 

suggestions can reposition space power to take maximum advantage of the resources 

that are available. 

423 

PftlelDtNCI PAGE BLANK NOT FILMED 

- - --------------------------------------~ 
• Set technology improvement goals at 100% minimum. 

The key word for the 90l

s should be focus. Since the resources will be very limited 

go for the gold. Select only high risk, high payoff, mission enabling technologies 

for consideration. Anything less than 100% improvement at the system level is 

evolutionary and will not achieve rapid market acceptance. 

• Time to market (Faster) is just as important to technology efforts as 

commercial efforts. 

The advantage scientists and engineers have over artists is that we 

shouldnlt have to be dead to see our work applied and appreciated. Focus! 

• Efficiency 

Efficiency drives everything: weight, volume, cost, etc. 

• Look for revolutionary, mission enabling systems. 

In order to tackle this suggestion, researchers and NASA center personnel will have to 

get out of their labs and offices and go talk to the users. Kind of a novel suggestion? 

Just remember, the customer is always right. Go find out what would generate new 

programs and public support. Tackle those problems. Focus! 

• What should the maximum operating voltage of a solar array be? 

• Are arc-proof arrays required for future space PV? 

• Do we need new array technology for the next generation of commercial satellites? 

If so, what is required? 

• Expendable arrays for complex missions -- Should arrays be expendable? 

• What is the operational range of PV in the solar system -- PV for Pluto? 

• Do we need new array technology for intermediate orbit applications? 

1The following paper presents the general results of the workshop and does not 

necessarily represent the views of any individual participant or company. 

424 

L ________ _ 
l--

I 

FUTURE DIRECTIONS IN PV CELL DEVELOPMENT: SUMMARY OF THE 

WORKSHOP AT SPRAT-XIII 

Geoffrey A. Landis 

NYMA, Inc. 

Brook Park, Ohio 

Workshop co-chairs: Geoffrey A. Landis (NYMAINASA Lewis) and Ed Gaddy (NASA Goddard) 

MISSIONS 

The "mission" of this workshop was to identify what areas of PV cell development would be most 

fruitful to direct NASA's scarce research money toward in order to have the greatest impact on future 

space power systems. 

Before analyzing what advances customers need, it is necessary to decide who are the customers 

for improved solar cells will be, and what orbits the cells will be required to operate in. The following list 

of customers was generated: 

NASA: Earth orbit missions: Low Earth Orbit (LEO) 

Intermediate Earth Orbit 

Geosynchronous Earth Orbit (GEO) 

Solar-system exploration: Inward (Sun, Mercury, Venus) 

Outward (Mars, Asteroids, Outer Planets) 

Commercial communications LEO 

Molniya orbit 

GEO 

Military communications, navigation, observation 

Other Government Agencies weather, navigation, observation 

It was decided that NASA research should be directed toward applications by NASA and 

commercial users . Representatives of the military and other government agencies at the workshop 

made no comment. Workshop co-leader Ed Gaddy of NASA Goddard Spaceflight Center suggested 

that there will be only a small number of NASA missions devoted to solar system exploration, and a 

large number of future missions to LEO and GEO orbits , where LEO missions will include many polarand near-polar (sun synchronous) orbits. [The representatives from NASA JPL were all in the other 

session on array technology, so no one challenged this statement.] The consensus of the workshop was 

that commercial missions will also be primarily to LEO and GEO orbits. Thus, the overall consensus of 

the group was that the most important applications for future solar cells will be for satellites in the orbits 

which are important today: LEO and GEO. 

Two strong trends were identified, with important implications for the future: a trend toward small 

satellites (less than a few kilowatts, as opposed to the 10 kW+ projects envisioned just a few years ago), 

and an industry-wide trend toward fast cycle times for development of new technology. 

425 

--__ I 
GOALS 

The next question is: in which solar cell parameters are advances most needed by the user 

community? 

The answer is : cost . Array volume and reliability of interconnects were mentioned, but the 

overwhelming answer of the participants was that cost is critical. However, in the ensuing discussion, it 

became clear that the important cost is not cell cost, but life-cycle system cost. 

The typical purchase cost to users of solar arrays ranges from $1000 to $2000 per watt today. 

Some of this is non-recurring cost, since power requirements are different for each satellite. One vendor 

said that they are providing arrays for $800 per watt to a military customer, where the non-recurring cost 

is amortized over many satellites. Frank Ho provided the following "typical" numbers. For a GaAs array 

of a few kilowatts, roughly 25% of the cost of the power system is the cost of the solar panels. Of the 

panel cost, roughly half of the cost is cell cost (for GaAs/Ge cells). Thus, about 12% of the cost of a 

power system is attributable to the cost of the cells for (relatively expensive) GaAs cells, and a few 

percent for (relatively cheap) silicon cells. From this we conclude that cutting the cell cost can have at 

most a 12% impact on the power system cost, and, considering launch costs and other system costs , is 

likely to have a much lower impact. In itself, cell cost is not a major issue. 

To achieve low system cost, the workshop participants suggested that the single most important 

factor is conversion efficiency, since an increased efficiency reduces the entire array cost. In addition, in 

order to get a new technology into the marketplace, investors require a low development cost, and a fast 

cycle time. 

Lew Fraas emphasized that low development cost is critical. He said that Boeing estimated that the 

development cost to bring their 30% efficient tandem GaAs/GaSb concentrator system to market would 

be $100 M, and that this high cost made it impossible to attract investors. Representatives of the spacecell industry said that development of the GaAs on Ge cell [1 ,2] required "lots of millions of dollars" and 

took over three years, but that it had a strong selling point in that the cells already had a customer, since 

they were direct replacements for existing GaAs on GaAs cells developed for an unnamed (presumably 

military) customer. 

It was also mentioned that low development cost means that the capital cost for production has to 

be reasonably low as well. There was a discussion of what low capital cost means , applied to space 

solar arrays. George Vendura of TRW pointed out that amorphous silicon may have a high capital cost if 

a new production facility must be built for the space product, but that TRW was pursuing a low capital 

cost approach for a-Si arrays by leveraging the huge (many megawatt) a-Si production capability in 

place for terrestrial markets . On the other hand, it was pointed out that a typical amorphous silicon 

production facility has a capital cost of about $10M. If this were expensed over a year's production of 50 

kW, the cost would be $200 per watt. This is only a small fraction of the current space-cell production 

cost, and if other costs (such as the cost of assembling the array) were reduced, it might be acceptable. 

The current industry trend is toward extremely fast cycle time: getting a product to market as swiftly 

as possible. Several of the participants suggested that for a new produce to fly, development time ought 

to be three years or less. Frank Ho said that getting MO-CVD GaAs cells to market took four years from 

the 1982 manufacturing technology (mantech) program. GaAs on Ge cells took three years after the 

mantech. Lew Fraas said that his experience at Boeing was that they had their research breakthrough 

in GaSb in 1989, found a flight opportunity in 1992 for a flight in 1994-- and the program was terminated 

by Boeing in 1992. The time scale of 5 years from technology to flight test was too long. 

It was debated whether a 3 year cycle time was possible. It was concluded that it may be possible 

for developments with low technical risk and the ability to use existing system heritage, as the GaAs/Ge 

cell did, where system components other than the cell can be transferred unchanged. 

It was concluded that space experience was the big stumbling block to short cycle times . It is 

important for NASA to use advanced cells on actual missions, in order to get the space heritage 

demanded by mission designers. A scientific satellite, for example, could be designed so that one of the 

panels of an array is made with advanced cells. 

426 

L __________________________ ~ _________ ! 
SOLAR CELL TECHNOLOGIES 

Seven different advanced cell technologies were discussed in some detail. 

Amorphous silicon , copper indium diselenide, and cadmium telluride thin films were discussed as 

systems that could have lower cost at the cell and array levels, and have the potential for very low mass 

and good radiation tolerance [3,4]. However, it was expected that to take maximum advantage of these 

systems, new array technologies would be needed. The workshop was divided on this issue. 

Ultra-thin (5 micron) gallium arsenide was discussed. The costs were considered higher, but the 

reduced cell mass would improve the specific power of arrays. 

High-efficiency monolithic tandem cells, such as GainP 2 on GaAs/Ge [5], and GaAs on active 

germanium, were discussed as ways of improving efficiency. Since these cells could be used directly as 

replacement for existing GaAs/Ge cells in existing arrays, this was considered a very promising 

approach. 

Indium phosphide was discussed, as well as the heteroepitaxial InP approaches such as InP/Ge 

and 'nP/Si. The cost is high today, but it was agreed that ultimately the cost of 'nP/Si or InP/Ge could be 

made competitive. It was agreed that these cells may have an application in orbits which see high 

radiation environments. 

The concentrator approach was discussed. This is not a direct replacement into existing arrays, 

and may find some resistance from program managers due to pointing requirements. However, they 

have the potential for high efficiency and good radiation tolerance [6]. 

Finally, it was mentioned that new generation silicon cells with efficiency of 20% have performance 

as good as GaAs cells. However, the radiation tolerance of these cells is yet to be determined, and they 

are not yet space qualified. 

Thermophotovoltaic (TPV) cells were mentioned as a promising new use for photovoltaics, but 

since there was a separate workshop on these concepts, they were not discussed in detail. 

DISCUSSION SESSION 

In the summary session, Geoffrey Landis took issue with the consensus that future systems will be 

primarily LEO and GEO. He suggested that the most significant commercial space system in the next 

decade will be the emplacement of a worldwide communications satellite network for portable 

telephone systems, with an investment of tens of billions of dollars, and that these satellite systems may 

be significantly different from currently operated GEO satellites. He presented results from an 

unpublished study [7] that shows that the number of satellites required to provide global phone 

coverage can be reduced by a factor of four if an intermediate orbit of 3200 kilometers is chosen instead 

of the low Earth orbit proposed. A page from these results is shown in figure 1. 

Andrew Meulenberg agreed with this conclusion. He said that a study done by Comsat on behalf 

of Inmarsat concluded that Inmarsat could save nearly a billion dollars on their worldwide telephone 

satellite constellation "Inmarsat-P" due to the reduced number of satellites required if they went to 

intermediate orbits instead of low orbit. According to articles in Space News [8], the price of the 

Inmarsat-P system reduces from 3 billion dollars to 2 billion, and the number of satellites is reduced from 

54 to 12-15, if intermediate orbit of 10,300 km is chosen instead of LEO. (This data point is shown to the 

right of the curve shown in figure 1). A recent study published in Space News indicated that there will 

be a market for as many as four of these worldwide communications satellites systems, and that these 

will produce a revenue of $9 billion per year [9] . 

These intermediate orbits see an intense radiation environment. At 3200 km, the radiation dose 

received from trapped protons in one day is approximately the same as that seen in geosynchronous 

orbit in a year! This implies that radiation tolerant solar cells may be critical components of future 

communications satellite networks , and thus could have considerable commercial value. Data 

presented by Landis showed that InP cells (and possibly other radiation-resistant cell types as well) may 

be able to stand up to this environment. 

427 

___ ~ __________________ . ____________ -l 
100 

Iridium 

eo (66 satellites) 

Ares 

N 60 

G10balstar 

(48 satellites) 

40 

20 

1000 2000 

Altitude Oem) 

Russian 'Signal' 

(24 satellites) 

30 00 4 000 

Inmarsat-P 

( 12- 15 sateWtes, 

)0,000 km) 

I--.-

Figure 1. Number of satellites required for a phone network to provide worldwide 

continuous coverage, as a function of altitude (data from Bailey and Landis [In . Solid line is 

theoretical curve, for polar-orbit satellite constellations; data points shown are for proposed 

American and Russian global coverage systems. 

Some further comments of note from the discussion of the workshop summary: 

Geoffrey Sommers said that telephone satellites will not have a lot of fancy electronics; they will be 

simple relays , and could be built so that the electronics will be (relatively) radiation tolerant. Thus, if the 

solar arrays could be made radiation tolerant, it would indeed be possible to utilize high-radiation orbits. 

Irving Weinberg said that radiation-tolerant cells such as InP are important for commercial 

applications in GEO, not just intermediate orbits. He said that a satellite in GEO accumulates a radiation 

dose of 1015 electrons over ten years, and that this results in degradation in power of 30 to 40 percent. 

Further, he notes that the next generation of commercial satellites are going to extended lifetimes of 

fifteen years and longer, making radiation-limited lifetime important. 

Finally, Lew Fraas concluded by reminding us that research aimed at near-term markets is a job for 

industry. The government should think in the long term, and fund technology development, so that we 

maintain a technology base for industry to draw on in the future. 

CONCLUSIONS 

Cost is the main issue for space photovoltaics , but eel/cost is only a minor (10%) component of the 

cost. The parameter that is most desired out of next-generation photovoltaic technology is high 

conversion efficiency. To get a product to market required fast cycle time and a low development and 

qualification cost. One thing that aids low development cost the ability to directly replace existing cells in 

existing array designs, so that a new array design doesn't have to be developed. 

There is a good argument that development of radiation tolerant cell technology could open up a 

new range of intermediate orbits, with potentially high commercial value. This may be a strong 

argument for continued development of InP and other radiation tolerant cell designs. 

REFERENCES 

(1) B. Smith, M. Gillanders, P. Vijayakumar, D. Lillington, H. Yang and R. Rolph , "Production Status of 

GaAs/Ge Solar Cells and Panels," Space Photovoltaic Research and Technology 1991, NASA CP3121, 9-1 (1991). 

(2) C.L. Chu , PA lies , and W. Patterson , "Recent Advances in GaAs/Ge Solar Cells ," Space 

Photovoltaic Research and Technology 1991 , NASA CP-3121 , 8-1 (1991). 

428 

1 ________ _ 
[3) G.A. Landis and A.F. Hepp, "Thin-Film Photovoltaics: Status and Applications to Space Power," Proc. 

European Space Power Conf., 2-6 Sept. 1991 , Florence Italy; pub. ESA SP-320, 517-522 (1991). 

[4] R.L. Mueller and B.E. Anspaugh, "Results of Some Initial Space Qualification Testing on Triple 

Junction a-Si and CulnSe 2 Thin Film Solar Cells," Proceedings of the XII Space Photovoltaic 

Research and Technology Conference, 108-117 {1992}. 

[5] M.F. Piszczor and M.J . O'Neill, "Recent Developments in Refractive Concentrators for Space 

Photovoltaic Power Systems," Proceedings of the XII Space Photovoltaic Research and Technology 

Conference, 206-216 (1992). 

[6] J.M. Olson, S.R. Kurtz, A.E. Kibbler, K.A. Bertness, and D.J. Friedman, "GalnP2/GaAs Tandem Cells 

for Space Applications," Space Pholovollaic Research and Technology 1991, NASA CP-3121, 12-1 . 

[7) S.G. Bailey and G.A. Landis, submitted to 1 st World Conference on Photovoltaic Energy Conversion, 

Hawaii, December 1994. 

[8] P. deSelding, "Inmarsat Drops LEO Concept," Space News, Aug . 2-8 1993; page 3 and "Inmarsat 

Takes First Step to Privatization," Space News, May 16-22 1994, page 1. 

(9] "Study Claims LEO Market Can Support Four Systems," Space News, June 13-19, 1994. 

429 
- - - - --- - - - - - - - ---

ARRAY TECHNOLOGIES WORKSHOP I 

Douglas Allen 

W.J. Schafer Associates 

North Olmsted. Ohio 

and 

Andrew Meulenberg 

Comsat Laboratories 

Clarksburg. Maryland 

The following is a summary of responses to questions posed to the workshop and related 

discussions. Approximately 40 people attended the workshop and included representatives of 

satellite design and fabrication companies, cell development and manufacturing companies, 

panel development and manufacturing companies, Universities, and several government 

organizations. Commercial, DoD, and civil applications were all well represented by workshop 

attendees, as were rigid, flexible, and concentrator array technologies. Most of the time in the 

workshop was spent discussing questions 1 and 2, the remaining questions received only 

minimal deliberation. 

Question 1. What will determine the selection of a specific array type (rigid, flexible, 

concentrator)? 

Workshop attendees all agreed that the first criteria for selecting an array on any satellite 

program is past use. This is because of the cost of designing and qualifying any new technology, 

which can add about $15M to the cost of a satellite program and add risk. Any time a new 

technology is integrated into a spacecraft design, there are many "what ifs" that must be studied 

and answered which is the cause of the high cost. 

There are only two ways that a new array technology will be used on a spacecraft in 

nearly all circumstances: I) the technology "enables" the mission or 2) The technology 

provides system level benefits that overcome the extra cost and risk. Enabling technology means 

that the new technology allow the mission to be accomplished when no existing technology will. 

For example, concentrator array technology may be enabling for some non-nuclear deep space 

missions or for orbital missions in the radiation belts. The second criteria allows spacecraft 

builders to take advantage of new technology if the risklbenefit analysis shows significant system 

level payoff for integrating a new array technology. In general, the new array technology must 

provide about a 2X improvement in an important parameter (volume, cost, mass, etc.) to meet 

this criteria. Otherwise, spacecraft builders do not feel it is worthwhile to accept the extra cost 

and risk of integrating a new technology. 

One positive sign that was discussed is that there appears to be a growing number of 

opportunities for space experiments and demonstrations that will help new array technologies 

overcome the past use criteria. 

After past use, the next issue for a new array technology is schedule. Many satellites are 

being designed, built, and launched on a 2-3 year schedule now and if a technology option can 

not meet the schedule for a program, the program will not wait. Another similar go/no go 

criterion on many satellite programs is stowed volume and configuration. The array must be able 

to be packaged into the planned launch vehicle or it will not be considered. 

After passing through the above gates, the criteria for selecting an array technology are 

cost, then area, then mass. It was noted that array mass tends to be much less important than the 

other criteria for most satellites. Area is important due to several considerations like the 

431 

pRMMWlMi PAGE LANK NOT FILMED 
spacecraft moment of inertia's impact on the attitude control subsystem and the need for drag 

makeup propellant in low orbits. 

The workshop also spent time discussing the growing importance of small satellites in the 

marketplace and the need for new technologies to be compatible with this trend. There was no 

consensus on the definition of what is a "small satellite." Some of the definitions offered 

included power below 1 kW, program cost below $50M, and a satellite that can be launched on a 

Pegasus. It was also noted that there may be an misperception about the importance of small 

satellites in the marketplace. For example, many of the new LEO communications constellations 

will include large numbers of satellites, but the satellites do not come close to any of the above 

definitions of small satellites. Teledesic at ] I kW (BOL) and Ellipsat at about 5 kW (EOL) were 

cited as examples. 

Other topics discussed in the course of answering the first question included the variety 

of possible concentrator arrays (concepts with concentration ratios of anywhere from 1.5X to 

lOOOX have appeared recently in the literature), the relationship between array stiffness and 

pointing tolerance, the ability of concentrator arrays to perform under spacecraft error conditions, 

and the fact that the number of satellites in a constellation and other constellation parameters 

changes the criteria for selection and the importance of each criteria. 

A concentrator array's ability to locate and track the sun when a spacecraft tumbles was 

the highlight of the other topics, as it folIowed up on discussion started in the morning session. 

Significant comments included suggestions about using hybrid arrays or body mounting a 

backup panel to overcome this problem. It was also noted that the degree of the potential 

problem was spacecraft dependent contingent on factors such as what percentage of the 

spacecraft's power is necessary to keep the spacecraft alive until the attitude problem is solved 

and the relative cost of losing one spacecraft (critical in a one of a kind science mission, but 

somewhat less important in a large constellation). It was also noted that a program may be 

willing to take the risk if the concentrator array provides enough system level cost and mass 

benefits. 

Finally, the workshop attendees agreed that the team that developed GaAs/Ge should be 

congratulated as it is an excellent model of how to achieve wide use of a new spacecraft 

technology in a short time span. The downside of this success is that any new technologies 

might require the same size budget and commitment to be successful. 

Question 2. Are 300 W/kg, 300 W/m 2, and $300/w achievable goals for rigid, flexible, and 

concentrator arrays? If not, what is practically achievable for each array type? 

Many of the workshop participants were very hesitant to specify numbers for these 

parameters for any of the array types. Many reasons were given: 

- Specific numbers vary greatly depending on orbit, design life, what's included (whole 

array or panel), amount of funding to be invested, timeframe, design of the spacecraft, etc. 

- Many times, it is not smart to optimize the mass or other parameters at the panel or 

array level - trades must be done at the spacecraft level and include other parameters such as 

reliability, cost, etc. 

- Performance predictions depend on what assumptions are made including how good 

predicted performance is, assumed investment to be made in the technology, etc. (can result in 

"fantasy" predictions) 

432 

l _____ _ 

"I 

I 

I 
1-

- Some of the parameters may be possible, but can not be practically achieved on a real 

spacecraft due to real world considerations (e.g. APSA technology was demonstrated at 130 

W /kg, but implemented on EOS at 32 W Ikg) 

With the above qualifiers, the group did fill out the table below to specifically answer the 

question. The column of "300 days" was added to stress the growing importance of delivering 

array hardware in less than one year (from receipt of the specification and order) to meet short 

spacecraft program schedules. In each box in the table, there are two sections. The upper section 

is the capability of each technology today . The lower section is the workshop's assessment of 

what is practically achievable in the foreseeable future for each array type, for each parameter. If 

the word "yes" appears, it means that 300 is achievable, if not, the number or range represents 

what is achievable. 

Array Type (NOW/POSSible> 300 W/rrt2 300 Wlkg 300 $IW 

Rigid 

Flexible 

Concentrators 

50 -75 

(20 

16 -

6 

150 -

250 

200 -

250 

300 Days 

Several observations were made about the table once it was completed. First, there is 

surprisingly little difference between the array types for some of the parameters such as in 

W/m 2. After reflection, this is probably due to the ability of each array type to use similar 

photovoltaic cell technology, although implementing advanced cell technology is different for 

each array type. Another observation that got a lot of discussion is that flexible array mass is 

strongly dependent on satellite integration. Some examples quoted were Hubble at 16 W /kg, 

Hubble technology available today at 24 W/kg, EOS at 32 W/kg, SAFE at 66 W/kg, and APSA 

at 130 W/kg. 

Flexible array mass is also very dependent on the power level. For any satellite design, 

there is a crossover power level above which flexible arrays are lighter than rigid arrays and 

below that point they are heavier. However, where this point is depends strongly on the 

assumptions you make and on the spacecraft design. Two studies were quoted: an ESA study 

433 
L 

that set the crossover at 7 kW (later updated to 15 kW) and a JPL study that placed the crossover 

at 700 W. This shows the importance of studying options for a specific satellite with available 

technology to select the optimum array type. 

A final comment was made about the reality of the numbers. Even though 50+ W /kg is 

possible with current rigid array technology, th ~ actual specific power is usually 20 - 25 W /kg on 

real satellites. Therefore, all the numbers on the table should be taken with the disclaimers listed 

at the beginning of this section in mind. 

Question 3. Why is there no general consensus amount the industry with regard to the 

future trends in array technology and is there (or should there be) a trend toward 

standardizing array technology to be applicable to a variety of missions? 

As was evident in the answer to question 2, different missions and spacecraft designs 

drive the design trades. This results in different types of arrays (rigid or flexible) being the best 

solution for different programs. Another important point from question 2 is that some 

parameters for the different types of arrays tend to be similar, which results in no one type of 

array standing out over the others for a wide variety of missions. 

With regard to standards, it was generally agreed that there will never be a "standard solar 

array." This is because mission requirements vary significantly and because individual 

aerospace companies want to advertise an edge or a benefit with their design. However, there 

essentially are already several mini-standards, as each major aerospace company uses their own 

standard array design when they bid most programs. 

It was noted that in lieu of standard arrays, the best way to reduce program costs would 

be to standardize an all-encompassing set of requirements to minimize the non-recurring testing 

required for each program. This could have a major impact on program costs since the recurring 

cost of the solar arrays is typically a small part of the total cost to a program. 

Question 4. What are the operational pointing requirements for concentrator arrays? Is 

two axis tracking worth the effort or are linear concentrators better? 

In general, at least ±20 of sun acceptance angle at the individual concentrator level should 

allow the use of standard, off-the-shelf array components such as sun sensors and array tracking 

gimbals. The exact requirement for this angle is dependent on several factors including 

manufacturing precision of the concentrator elements, positioning accuracy of the concentrators 

into the panel, panel to panel alignment, gimbal pointing accuracy, sun sensor accuracy, thennal 

distortions, and tolerance of defects in the concentrator concept. A well managed error budget 

taking all these factors into account will result in determination of the allowable acceptance angle 

for a given concentrator design. 

For a linear concentrator, a sun acceptance angle of at least ±20 to 300 in the linear axis 

allows easy integration into a wide range of satellites, as this is a typical requirement on satellites 

that only track the sun in one axis (mission design and attitude control result in maintaining 

nominal pointing in the other axis). 

As was discussed in question 1, it was again noted that minimizing the time for sun 

acquisition is an important consideration in designing a satellite concentrator array . This is 

important to minimize the impact of possible attitude control subsystem problems and to 

434 
minimize the required battery size for initial operation of the spacecraft when deployed from the 

launch vehicle. 

Regarding 2 axis vs. 1 axis tracking concentrator arrays, the trade on which will better 

meet a spacecraft's requirements is very mission dependent. However, in general there is more 

applicability for single axis tracking concentrators. One reason for this is the cost and reliability 

of a 2 axis tracker. Another is that many spacecraft now use only a single axis tracker and 

changing to a two axis tracking system can have significant impact on other spacecraft 

subsystems and the overall design concepts. DoD and NASA have decided to move forward 

with the single axis systems first due to several factors including lower development and 

recurring costs, applicability to a wider range of satellites (any satellite that can use a two axis 

concentrator can also use a one axis system, but the reverse is not true), and lower risk for flight 

testing a full size array. Once the technology is demonstrated, an individual program may decide 

that the benefits of two axis technology is worth the added investment to optimize that 

spacecraft. 

Question 5. What are the advantages of integrating the solar array into the overall satellite 

to further optimize the satellite and is there interest in doing this? 

(Due to time limitations, there was not a lot of discussion about this question.) The 

advantages of complete integrated design of the array into a satellite vary by the individual 

satellite and mission. Some examples include direct drive of high voltage (300V) electric 

propulsion thrusters to eliminate the power conditioners and optimizing the voltage to match 

high load users to minimize cable and electronics mass. It was generally agreed that every 

company is already looking for ways to do this for every satellite, and it is really required for 

companies to stay competitive. 

435 

- ---.---
WORKSHOP SUMMARY 

THERMOPHOTOVOLTAICS AND NON-SOLAR ENERGY CONVERSION 

Donald L. Chubb 

NASA Lewis Research Center 

Cleveland, Ohio 

and 

Larry C. Olsen 

Washington State University 

Richland, Washington 

The workshop was well attended (about 20) with the discussion limited to alpha/betavoltaics and 

thermophotovoltaics (TPV). TPV was the major part of the discussion. Both space and terrestrial 

applications were presented for TPV at various power levels. However, alpha/ betavoltaics appear to be 

limited to very low power (mw) level applications. Reference 1 lists several low power applications for 

betavoltaics. 

ALPHA/BETAVOLTAICS 

In an alphavoltaic or betavoltaic energy converter charge carriers are produced in a p-n junction 

semiconductor by alpha particles or beta particles rather than photons as in a conventional photovoltaic 

energy converter. One of the key issues discussed for these devices was radiation damage. For a 

betavoltaic device the threshold for damage begins for beta particles with energies greater than 200-300 

KEV. Promethium (Pm 147) and tritium (H3

) were mentioned as possible beta sources with energies less 

than the damage threshold. For beta sources with energies greater than the damage threshold, 

strontium (Sr<') and thallium (T1 204

) were mentioned. Since the half life for each of these sources is long, 

the potential lifetime for an alpha/betavoltaic device is long. Botti InP and SiC were discussed as 

possible semiconductor materials suitable for alpha/betavoltaics. To make these devices feasible, the 

radiation damage problem must be solved. 

437 

G NK r 0 FILMED 
THERMOPHOTOVOLTAICS (IPV) 

The TPV discussion centered around possible applications and the key research areas. 

TPV Applications 

TPV has both low power (:s 100 w) and high power space applications. The low power 

applications are for deep space missions such as the Pluto flyby where the thermal energy is supplied 

by radioisotope decay. For higher power applications such as earth orbit or a moon base, solar energy 

can be used as the thermal source. In this case TPV has an advantage over the conventional PV-battery 

system since thermal energy storage can be used. The TPV system thus will have a lower mass, 

especially for a moon base if the lunar soil can be used as a storage material. 

It is the commercial applications of TPV that has caused the great interest in TPV in the last 

several years. Although the discussion was limited on this subject the following applications were 

mentioned: portable power supplies for recreational vehicles and Army field units, cogeneration of 

electrical power for natural gas appliances such as electrical power for furnace blowers, and hybrid 

electrical vehicles. 

TPV Research 

Two important developments have made efficient TPV energy conversion at moderate 

temperatures « 2000 K) possible. One is low bandgap energy « 1 eV) PV cells such as GaSb and 

InGaAs and the other is efficient selective emitters. Research in these areas was thus the main topic of 

discussion. For high temperature (~ 2000K), high efficiency can be attained using Si PV cells. 

There are two approaches to attaining an efficient TPV emitter. First of all a selective emitter 

that emits mainly in the wavelength region where the PV celis have maximum efficiency and secondly a 

grey body emitter with a band pass filter to make the emitter behave like a selective emitter. Discussion 

was centered on rare earth selective emitters. The mantle type emitter made of small (5-10 f.Jm) 

diameter rare earth oxide fibers, such as Yb203 , has demonstrated good efficiency. Research is 

continuing on the fiber emitter with different geometry than the mantle. Also, a new rugged, rare earthgarnet emitter shows promising spectral emittance. 3 

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Currently there is considerable research on low bandgap energy PV cells. Probably the most 

developed low bandgap energy, Eg, cell is GaSb with Eg = .72 eV. Also being actively researched is 

InxGa,_xAs on !nP substrates. This system yields .36 :s Eg :s 1.42 eV depending on the value of x. Also, 

just beginning is research on InxGa,_xSb which has .17 :s Eg:s .72eV depending on the value of x and 

InxGa,_xAsySb,_yl which has .17 :s Eg :s 1.42eV depending on the value of x and y. This later quantenary 

system will allow lattice matched growth on GaSb substrates in the energy range .3 :s Eg :s .72eV. Two 

other PV materials that are not currently being considered, but should be considered, are HgxCd,_xTe 

and Ge. The HgCdTe system allows lattice matched growth for 0 :s Eg :s 1.56eV. The main advantage of 

Ge (Eg = .66 eV) is that it potentially should be the lowest cost. 

Most people felt that low bandgap energy PV cell development will occur before an efficient 

emitter is developed. The main reason for this is the large amount of cell research compared to emitter 

research. For an efficient TPV system both an efficient emitter and PV cell are required. 

CONCLUSION 

Both low bandgap energy PV cell research and emitter research are required to make efficient 

TPV energy conversion possible. The many potential applications of TPV more than justify the research 

effort. Alpha/ betavoltaic energy conversion will be viable if the radiation damage problem can be 

solved. 

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References 

1. P. W. Sims, L. C. DiNetta and A. M. Barrett, "High Efficiency GAP Power Conversion for 

Betavoltaic Applications," Proceedings of the XIII Space Photovoltaic Research & Technology 

Conference (SPRAT XIII), Cleveland, Ohio, June 1994. 

2. R. E. Nelson, in "Proceedings of 32nd !nternational Power Sources Symposium," 

(Electrochemical Society, Pennington, NJ, 1986), pp. 95-101 . 

3. R. A. Lowe, D. L. Chubb, S. C. Farmer and B. S. Good, "Rare Earth-Garnet Selective Emitter," 

Appl. Phys. Letts., 64, 3551 (1994). 

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13. ABSTRACT (Maximum 200 words) 

The Thirteenth Space Photovoltaic Research and Technology (SPRAD Conference gathered representatives from 26 

commercial corporations, 8 universities and 7 governmental agencies, including Europe, for two and a half days of 

presentations and discussions regarding the status and future of space photovoltaics, The conference was well attended, 

with over 100 attendees, and included 38 technical papers, 6 program reviews and 5 workshop discussions. The effects of 

shrinking research and development budgets were evident in the focus and tone of this SPRAT. Most attendees appeared 

to be oriented toward near term, system oriented projects and fewer were involved in long term , high risk research. It was 

generally agreed that space power requirements would continue to move toward small er «2kW) power levels. Most 

future applications are believed to be in traditional orbits (LEO, GEO) although interesting opportunities may be found in 

high radiation, mid-altitude orbits useful for global communication networks. New solar cell devices and materials will be 

difficult to introduce unless they are mission enabling, or offer significant cost and/or performance benefits. The attendees 

were unable to come to a consensus regarding the type of array (eg. rigid, concentrator, flexible) suitable for specific 

missions. Many factors outside the realm of photovoltaics influence the selection process. These topics and many more 

are covered in the following pages of this record. 

14. SUBJECT TERMS IS. NUMBER OF PAGES 

452 

Photovoltaics; Space power; Solar cells 16. PRICE CODE 

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