18th Space Photovoltaic Research

and Technology Conference

NASA/CP—2005-213431

April 2005
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18th Space Photovoltaic Research

and Technology Conference

NASA/CP—2005-213431

April 2005

National Aeronautics and

Space Administration

Glenn Research Center

Proceedings of a conference

held at Ohio Aerospace Institute

and sponsored by NASA Glenn Research Center

Brook Park, Ohio, September 16–18, 2003
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Available electronically at http://gltrs.grc.nasa.gov
NASA/CP—2005-213431 iii

Foreword 

The 18th Space Photovoltaic Research and Technology Conference, SPRAT XVIII, was held from 

September 16 to 18, 2003. The purpose of the SPRAT conference is to bring members of the space solar 

cell community together to discuss the recent developments in solar cell technology and future directions 

of the field. The SPRAT conference is convened roughly every 2 years. The venue for SPRAT XVIII was 

the Ohio Aerospace Institute (OAI), located just outside the gate of the NASA Glenn Research Center at 

Lewis Field. SPRAT is sponsored by the Photovoltaic and Space Environmental Effects Branch at the 

NASA Glenn Research Center. The conference organizers were all NASA and OAI staff. They did an 

outstanding job and it was my pleasure to serve as chairman. 

Representatives of industry, government, and universities presented topical papers on space solar cell 

and array technology. The SPRAT conference also holds a series of half-day workshops on current issues 

in space photovoltaics. The papers and workshop summaries are included in the SPRAT proceedings. 

As the SPRAT XVIII chairman, I had the pleasure of awarding the Irving Weinberg Award to a 

researcher who has made significant contributions to space solar cell research and technology. The 

SPRAT XVIII Irving Weinberg Award went to Dr. Bruce Anspaugh of the Jet Propulsion Laboratory. He 

made significant contributions in the field of solar cell calibration and radiation damage. Dr. Anspaugh 

retired from JPL this year. I wish Bruce well and hope he has a long and happy life. 

Phillip Jenkins 

Chairman of the 18th Space Photovoltaic Research and Technology Conference 

NASA/CP—2005-213431 v

Contents 

Foreword................................................................................................................................................... iii 

Invited Papers 

TOWARDS REALIZING SUPER-HIGH-EFFICIENCY AND RADIATION-RESISTANT 

MULTIJUNCTION SOLAR CELLS 

 Masafumi Yamaguchi, Toyota Technological Institute; Tatsuya Takamoto, Sharp 

Corporation; Aurangzeb Khan, University of South Alabama; Mitsuru Imaizumi, NASDA; 

Sumio Matsuda, NASDA; and Nicholas J. Ekins-Daukes, Toyota Technological Institute.............. 1 

MULTIJUNCTION GaAs PHOTOVOLTAIC ASSEMBLY TECHNOLOGY FOR VERY 

SEVERE LEO ENVIRONMENT MISSIONS 

 A. Caon and C. Signorini, European Space Agency/European Space Research and 

 Technology Centre; and A. Bals and K. Dettlaff, European Aeronautic Defence and 

 Space Astrium .................................................................................................................................... 9 

CURRENT STATUS OF PHOTOVOLTAIC DEVELOPMENT AT AFRL 

 John Merrill, Paul Hausgen, Donna Senft, and Pawel Tlomak, Air Force Research 

 Laboratory, Space Vehicles Directorate, Kirtland Air Force Base .................................................... 18 

Session I 

Systems and Flight Experiments 

RESULTS OF FLIGHT DEMONSTRATION OF TERRESTRIAL SOLAR CELLS BY MDS−1 

IN GTO 

Mitsuru Imaizumi, Taishi Sumita, Shirou Kawakita, Osamu Anzawa, and Sumio Matsuda, 

National Space Development Agency of Japan (Current name Japan Aerospace Exploration 

 Agency) .............................................................................................................................................. 24 

SOLAR ARRAY VERIFICATION AND ANALYSIS TOOL (SAVANT) ANALYSIS OF THE 

MICROELECTRONICS AND PHOTONICS TESTBED (MPTB) SPACE SOLAR CELL 

DATA 

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

 SFA, Inc.; and T.L. Morton, Ohio Aerospace Institute ...................................................................... 31 

RECENT PROGRESS ON THE STRETCHED LENS ARRAY (SLA) 

 Mark O’Neill and A.J. McDanal, ENTECH, Inc.; Michael Piszczor and Patrick George, 

 NASA Glenn Research Center; Michael Eskenazi and Matthew Botke, ABLE Engineering 

 Company; David Edwards and David Hoppe, NASA Marshall Space Flight Center; and 

 Henry Brandhorst, Auburn University ............................................................................................... 40 

MATERIALS ON THE INTERNATIONAL SPACE STATION⎯FORWARD TECHNOLOGY 

SOLAR CELL EXPERIMENT 

 R.J. Walters, J.C. Garner, S.N. Lam, J.A. Vazquez, W.R. Braun, and R.E. Ruth, U.S. Naval 

 Research Laboratory; J.R. Lorentzen, SFA, Inc.; CDR R. Bruninga (Ret.), U.S. Naval 

 Academy; P.P. Jenkins, Ohio Aerospace Institute; J.M. Flatico, QSS; and D.M. Wilt, M.F. 

 Piszczor, L.C. Greer, and M.J. Krasowski, NASA Glenn Research Center....................................... 51 
NASA/CP—2005-213431 vi

PROMISING RESULTS FROM THREE NASA SBIR SOLAR ARRAY TECHNOLOGY 

DEVELOPMENT PROGRAMS 

 Mike Eskenazi, Steve White, Brian Spence, Mark Douglas, Mike Glick, Ariel Pavlick, and 

 David Murphy, ABLE Engineering; Mark O’Neill and A.J. McDanal, ENTECH, Inc.; and 

 Michael Piszczor, NASA Glenn Research Center.............................................................................. 59 

Session II 

Multijunction Solar Cells 

DEVELOPMENT OF A HIGH-EFFICIENCY MECHANICALLY STACKED 

MULTIJUNCTION SOLAR CELL 

 Daniel Aiken, Paul Sharps, and Mark Stan, EMCORE PhotoVoltaics; Harry Atwater, Anna 

 Fontcuberta i Morral, and James Zahler, California Institute of Technology; and Mark 

 Wanlass, National Renewable Energy Laboratory............................................................................. 95 

RECENT ADVANCES IN HIGH-EFFICIENCY III−V MULTIJUNCTION SOLAR CELLS 

FOR SPACE APPLICATIONS: ULTRA TRIPLE JUNCTION QUALIFICATION 

 Hojun Yoon, Jennifer E. Granata, Peter Hebert, Richard R. King, Christopher M. Fetzer, 

 Peter C. Colter, Kenneth M. Edmondson, Daniel Law, Geoffrey S. Kinsey, Dmitri D. Krut, 

 James H. Ermer, Mark S. Gillanders, and Nasser H. Karam, Spectrolab, Inc. .................................. 102 

HIGH-EFFICIENCY, MULTIJUNCTION CELLS WITH MONOLITHIC BYPASS DIODES 

 P.R. Sharps, M.A. Stan, D.J. Aiken, B. Clevenger, J.S. Hills, and N.S. Fatemi, EMCORE 

 PhotoVoltaics ..................................................................................................................................... 108 

DENDRITE GROWTH AND DEGRADATION IN MULTIJUNCTION SOLAR CELLS 

 Donna Speckman, Dean Marvin, and John Nocerino, The Aerospace Corporation .......................... 116 

INDIUM PHOSPHIDE WINDOW LAYERS FOR INDIUM GALLIUM ARSENIDE SOLAR 

CELLS 

 Raj K. Jain, NASA Glenn Research Center ....................................................................................... 126 

Session III 

Calibration and Measurement of Solar Cells 

DESIGN AND PERFORMANCE OF A TRIPLE SOURCE AIR MASS ZERO SOLAR 

SIMULATOR 

 Phillip Jenkins, Ohio Aerospace Institute; David Scheiman, Ohio Aerospace Institute; and 

 David B. Snyder, NASA Glenn Research Center .............................................................................. 134 

MEASUREMENT AND CHARACTERIZATION OF CONCENTRATOR SOLAR CELLS II 

 Dave Scheiman, Ohio Aerospace Institute; Bernard L. Sater, PhotoVolt, Inc.; Donald Chubb, 

 NASA Glenn Research Center; Phillip Jenkins, Ohio Aerospace Institute; and David B. 

 Snyder, NASA Glenn Research Center.............................................................................................. 139 

HIGH-ALTITUDE AIR MASS ZERO CALIBRATION OF SOLAR CELLS 

 James R. Woodyard, Wayne State University; and David B. Snyder, NASA Glenn 

 Research Center.................................................................................................................................. 148 
NASA/CP—2005-213431 vii

HISTORICAL PRECISION OF AN OZONE CORRECTION PROCEDURE FOR AM0 SOLAR 

CELL CALIBRATION 

 David B. Snyder, NASA Glenn Research Center; and Phillip Jenkins and David Scheiman, 

 Ohio Aerospace Institute .................................................................................................................... 166 

EVALUATION OF THIN-FILM SOLAR CELLS FOR SPACE APPLICATIONS 

 S.H. Liu, E.J. Simburger, J. Matsumoto, A. Garcia III, J. Ross, and J. Nocerino, The 

 Aerospace Corporation....................................................................................................................... 170 

Session IV 

Thin Films⎯I 

DEVELOPMENT OF CIGS2 THIN FILMS ON ULTRALIGHTWEIGHT FLEXIBLE LARGE 

AREA FOIL SUBSTRATES 

 Neelkanth G. Dhere, Vivek S. Gade, Ankur A. Kadam, Anant H. Jahagirdar, Sachin S. 

 Kulkarni, and Sachin M. Bet, Florida Solar Energy Center ............................................................... 179 

Cu(InGa)Se2 SOLAR CELLS ON A FLEXIBLE POLYMER WEB 

 Robert Birkmire, Erten Eser, Shannon Fields, and William Shafarman, Institute of Energy 

 Conversion, University of Delaware .................................................................................................. 190 

WIDE-BANDGAP CIAS THIN-FILM PHOTOVOLTAICS WITH TRANSPARENT BACK 

CONTACTS FOR NEXT-GENERATION SINGLE AND MULTIJUNCTION DEVICES 

 Lawrence M. Woods, Ajay Kalla, Damian Gonzalez, and Rosine Ribelin, ITN Energy 

 Systems, Inc........................................................................................................................................ 195 

DEVELOPMENT OF A THIN-FILM SOLAR CELL INTERCONNECT FOR THE 

POWERSPHERE CONCEPT 

 Edward J. Simburger, James H. Matsumoto, Thomas W. Giants, Alexander Garcia III, and 

 Simon Liu, The Aerospace Corporation; Suraj P. Rawal, Alan R. Perry, and Craig H. 

 Marshall, Lockheed Martin Corporation; John K. Lin and Stephen E. Scarborough, ILC 

 Dover Incorporated; and Henry B. Curtis, Thomas W. Kerslake and Todd T. Peterson, NASA 

 Glenn Research Center....................................................................................................................... 202 

Session V 

Radiation Effects 

RESEARCH TRENDS TO ENHANCE THE RADIATION TOLERANCE OF Si SPACE 

SOLAR CELLS 

 A. Khan and Al-Khatib, University of South Alabama; M. Yamaguchi, Toyota Technological 

 Institute; T. Abe, Shin-Etsu Handotai; and M. Imaizumi and S. Matsuda, National Space 

 Development Agency of Japan........................................................................................................... 208 

USING SRIM TO CALCULATE THE RELATIVE DAMAGE COEFFICIENTS FOR SOLAR 

CELLS 

 Scott R. Messenger, SFA, Inc.; Edward A. Burke, Consultant, SFA, Inc.; Robert J. Walters 

 and Jeffrey H. Warner, Naval Research Laboratory; and Geoffrey P. Summers, Naval 

 Research Laboratory and The University of Maryland, Baltimore County ....................................... 217 
NASA/CP—2005-213431 viii

Session VI 

Advanced Cell Structures 

PHYSICS OF QUANTUM STRUCTURES IN PHOTOVOLTAIC DEVICES 

 Ryne P. Raffaelle and John D. Andersen, Physics Department, Rochester Institute of 

 Technology......................................................................................................................................... 226 

SINGLE WALL CARBON NANOTUBE-POLYMER SOLAR CELLS 

 Sheila G. Bailey, NASA Glenn Research Center; Stephanie L. Castro, Ohio Aerospace 

 Institute; and Brian J. Landi, Thomas Gennett, and Ryne P. Raffaelle, NanoPower Research 

 Laboratories, Rochester Institute of Technology ............................................................................... 233 

HIGH-TEMPERATURE SOLAR CELL DEVELOPMENT 

 Geoffrey A. Landis, NASA Glenn Research Center; Danielle Merritt and Ryne P. Raffaelle, 

 Rochester Institute of Technology; and David Scheiman, Ohio Aerospace Institute ........................ 241 

INTEGRATED PHASE ARRAY ANTENNA/SOLAR CELL SYSTEM FOR FLEXIBLE 

ACCESS COMMUNICATIONS (IA/SAC) 

 E.B. Clark, R.Q. Lee, A.T. Pal, D.M. Wilt, and B.D. McElroy, NASA Glenn Research 

 Center; and C.H. Mueller, Analex Corporation.................................................................................. 248 

Session VII 

Spacecraft Charging 

ELECTROSTATIC CHARGING OF MIRRORS IN SPACE: A PLAUSIBLE CAUSE OF 

SOLAR PANEL ANOMALIES ON SATELLITES 

 Shu T. Lai, Space Vehicles Directorate, Air Force Research Laboratory, Hanscom Air 

 Force Base .......................................................................................................................................... 252 

EXPERIMENTAL STUDY OF ARCING ON HIGH-VOLTAGE SOLAR ARRAYS 

 Boris Vayner, Ohio Aerospace Institute; and Joel Galofaro and Dale Ferguson, NASA Glenn 

 Research Center.................................................................................................................................. 256 

Session VIII 

Thin Films⎯II 

LIQUID-PHASE DEPOSITION OF SINGLE-PHASE ALPHA-COPPER-INDIUMDISELENIDE 

 J. Cowen, L. Lucas, F. Ernst, and P. Pirouz, Case Western Reserve University; and A. Hepp 

 and S. Bailey, NASA Glenn Research Center.................................................................................... 273 

RECENT PROGRESS IN CuInS2 THIN-FILM SOLAR CELL RESEARCH AT NASA GLENN 

 M.H.-C. Jin, K.K. Banger, and C.V. Kelly, Ohio Aerospace Institute; J.H. Scofield, Oberlin 

 College; and J.S. McNatt, J.E. Dickman, and A.F. Hepp, NASA Glenn Research Center................ 285 

A SILICONE/POSS COATING FOR THIN-FILM SOLAR CELLS AND CONCENTRATORS 

 Henry W. Brandhorst, Auburn University; Mark J. O’Neill, ENTECH, Inc.; and Joseph D. 

 Lichtenhan, Hybrid Plastics, Inc. ....................................................................................................... 290 
NASA/CP—2005-213431 ix

Workshops 

FUTURE DIRECTIONS FOR THIN FILMS IN SPACE WORKSHOP AT SPRAT XVIII 

 Aloysius F. Hepp and John E. Dickman, NASA Glenn Research Center.......................................... 294 

NANOSTRUCTURES FOR SOLAR CELLS 

 Ryne P. Raffaelle, Rochester Institute of Technology ....................................................................... 297 

ADVANCED SOLAR CELL TESTING AND CHARACTERIZATION 

 Sheila Bailey, Henry Curtis, and Michael Piszczor, NASA Glenn Research Center......................... 299 

Weinberg Award....................................................................................................................................... 303 

2003 Group Photograph............................................................................................................................ 304 
TOWARDS REALIZING SUPER-HIGH-EFFICIENCY

AND RADIATION RESISTANT MULTI-JUNCTION SOLAR CELLS

Masafumi Yamaguchi,

Toyota Technological Institute, Nagoya 468-8511, Japan

Tatsuya Takamoto,

Sharp Corporation, Nara 639-2198, Japan

Aurangzeb Khan

University of South Alabama, Mobile 36688, USA

Mitsuru Imaizumi

NASDA, Tsukuba 305-8505, Japan

Sumio Matsuda

NASDA, Tsukuba 305-8505, Japan

Nicholas J. Ekins-Daukes

Toyota Technological Institute, Nagoya 468-8511, Japan

1 INTRODUCTION

The idea that a multi-junction solar cell can achieve high efficiency has a long history, dating back to 1957 [1]. However the

actual realization of efficient multi-junction solar cells that demonstrate the advantages predicted by the early papers, has taken

approximately 30 years. Currently there is cause for optimism, as the InGaP/GaAs/Ge solar cell has begun to penetrate the

space market and considerable efforts are underway to bring highly efficient multi-junction solar cells to terrestrial concentrator

systems. This paper will provide an overview of important developments over the last decade or so, with particular emphasis

on work performed in Japan.

2 KEY ISSUES FOR REALIZING HIGH-EFFICIENCY MULTI-JUNCTION SOLAR CELLS

A key point in the development of Japanese multi-junction solar cells was the Super-High Efficiency Solar Cell project, that

started in the fiscal year 1990. The programme marked a long term commitment to develop a 40% efficient solar cell and faced

considerable technical challenges. The multi-junction approach was, and still is, the only realistic means of attaining such high

efficiency in the medium term.

Table 1 summarises the past, present and anticipated future trends of high efficiency multi-junction solar cells. In the 1980’s

much effort was invested in the monolithic AlGaAs/GaAs tandem solar cell, but progress was slowed by oxygen contamination

of the AlGaAs. Extremely low surface recombination velocities were observed in the more oxygen tolerant InGaP material [2]

and it soon became the material of choice for multi-junction solar cells, leading to a number of notable achievements using the

InGaP/GaAs tandem structure and mechanical stack designs. A study to optimise the InGaP top cell showed that the lowest

surface recombination was obtained using an AlInP window layer and the longest minority lifetime obtained with a buffer layer

and optimised growth temperature [3]. This resulted in a record efficiency 18.5% InGaP solar cell.

The tunnel junction forms another important aspect of the multi-junction structure. It should be optically transparent and

connect the component cells in the multi-junction structure with the minimum of electrical resistance. The optical absorption

in the tunnel junction can be minimised by using thin layers of wide band-gap material, although care must be exercised as

the tunnelling current decreases exponentially with increasing band-gap energy. A double heterostructure (DH) was initially

proposed for AlGaAs/GaAs tandem cells with the DH layers blocking dopant diffusion [4]. Similar advantages were found when

NASA/CP—2005-213431 1
Table 1: Key issues for realizing high-efficiency multi-junction solar cells

Key Issue Past Present Future

Top Cell AlGaAs InGaP AlInGaP

3rd Layer None Ge GaInNAs etc...

Substrate GaAs Ge Si

Tunnel Junction DH-GaAs DH InGaP DH wide/narrow Eg

Lattice Matching GaAs In0.01GaAs Lattice Mismatch

Carrier Confinement InGaP BSF AlInP BSF Widegap BSF

Photon Confinement None None Bragg reflector etc..

Back Contact

n+ (In)GaAs

n+ AlInP [Si]

n+ InGaP [Si]

p InGaP [Zn]

p AlInP [Zn]

p++ AlGaAs [C]

n++ InGaP [Si]

n+ AlInP [Si]

n+ (In)GaAs [Si]

p (In)GaAs [Zn]

p+ InGaP [Zn]

p Ge Substrate

p++ AlGaAs [C]

n++ InGaP [Si]

n+ GaAs : 0.1μm

n+ (In)GaAs [Si]

n G e

Bottom Cell

InGaP

Top Cell

InGaAs

Middle Cell

Buffer Layer

Front C ontact AR Coating

Tunnel Junction

Tunnel Junction

0

20

40

60

80

100

Wavelength (nm)

400 800 1200 1600

4

6

8

10

12

14

16

AM1.5G, 1cm2

Jsc: 14.46 mA/cm2

Voc: 2.552 V

FF: 85.8 %

Eff.: 31.7 %

In0.49Ga0.51P

In0.01Ga0.99As

Ge

Quantum Efficiency (%)

Current Density (A/cm2)

0 0.5 1 1.5 2 2.5 3

Voltage (V)

0

2

Fig. 1. Schematic structure, light-IV and spectral response for the high efficiency InGaP/InGaAs/Ge 3-junction cell fabricated

on a Ge substrate. (From Ref. [8])

the DH tunnel junction was applied to the InGaP/GaAs tandem solar cell. Compared to a plain InGaP tunnel junction, the DH

tunnel junction with AlInP barriers achieved a peak tunneling current of 2A/cm2 while the plain InGaP tunnel junction achieved

a peak tunnelling current of only 5mA/cm2. This improvement is due to the lower diffusion coefficient of Zn in the AlInP and

InGaP materials [5]. An additional benefit of the DH tunnel junction is the provision of a back-surface-field and window layer

for the adjacent p/n junctions.

By growing the InGaP/GaAs device on a Ge substrate it is possible to create a third junction, thereby marginally increasing

the voltage of the multi-junction device. A further improvement can be made by paying careful attention to the lattice parameter

of the Ge substrate. It is commonly assumed that the 0.08% lattice mismatch between GaAs and Ge is negligibly small, but it

is still sufficient to introduce misfit dislocations when a thick GaAs epilayer is grown. By replacing the GaAs with In0.01GaAs, a

fully lattice-matched cell can be made with a negligible misfit dislocation density. A further advantage of the In0.01GaAs layer

is a marginal increase in short-circuit current over the GaAs, on account of the lower band-gap of the In0.01GaAs [6,7].

3 HIGH EFFICIENCY & RADIATION RESISTANT InGaP/GaAs/Ge 3-JUNCTION SOLAR CELLS

The most recent InGaP/GaAs/Ge solar cell design uses a C-doped AlGaAs/ Si-doped InGaP heterostructure tunnel junction [9] with DH AlInP barriers, described previously. A schematic of the device is shown in figure 1 together with light-IV and

quantum efficiency results from a 1cm2 device. Such solar cells have achieved efficiencies of 31.7% (1cm x 1cm) and 31.2%

(5cm x 5cm) under 1 sun AM1.5G and were fabricated by Japan Energy Co. [8]. Under AM0, a 29.2% efficient 2x2cm2 device

has been demonstrated; the light-IV is shown in figure 2.

Other notable cells include:

– InGaP/GaAs/InGaAs mechanical stack that achieved a world record 33.3% efficiency under AM1.5G, following joint work by

Japan Energy Co., Sumitomo Electric Co. and Toyota Tech. Inst. [10], shown in figure 3.

– AlGaAs/GaAs monolithic tandem cell with a 1 sun AM1.5 efficiency of 27.6% by Hitachi Cable Co [11].

NASA/CP—2005-213431 2
0

20

40

60

80

0 500 1000 1500 2000 2500 3000

VOLTAGE (mV)

Size: 2x2 cm2

Voc: 2567 mV

Jsc: 17.9 mA/cm2

FF: 0.858

Eff: 29.2 %

AM0, 135.5mW/cm2

Current (mA)

Fig. 2. Light-IV curve for the InGaP/GaAs/Ge 3-junction cell under AM0 illumination.

0

5

10

15

20

25

30

0 0.5 1 1.5 2 2.5

Voltage (V)

InGaAs (6.2%)

InGaP/GaAs (27.1%)

AM1.5G Total: 33.3%

Current (mA/cm2)

JQA

Fig. 3. Schematic structure for the InGaP/GaAs/InGaAs mechanically stacked 3-junction solar cell and light-IV curve showing

the world-record efficiency under 1 sun AM1.5G. (From Ref. [10]

– GaAs/GaInAsP mechanical stack tandem cell with a 1 sun AM1.5 efficiency of 31.1% by Sumitomo Electric Co. [12].

– InGaP/GaAs/Ge concentrator cell with an efficiency in excess of 36% for concentrations from 100-500 suns [13]; the increase

in efficiency under concentration is shown in figure 4.

Due to the series connection in the multi-junction solar cell, the radiation resistance of the cell as a whole will be dominated

by the worst performing layer. Figure 5 shows the degradation in both Voc and Isc for the sub-cells of a InGaP/GaAs/Ge multijunction cell under 1MeV electron irradiation. The effect of radiation damage is first a reduction in the minority carrier diffusion

length followed by majority carrier removal under heavy irradiation [14,15]. It is clear that the InGaP is reasonably radiation

hard, maintaining a high Isc, while the Ge and GaAs junctions fall quite quickly. However, as the Ge junction is current rich,

its effect on the overall multi-junction cell radiation performance is small. It is then the degradation of the GaAs junction that

dominates the degradation of the multi-junction cell under irradiation. Various schemes such as thinning the emitter and base

layers and field assisted collection can be employed to improve the radiation response of each sub-cell in a multi-junction device.

When considering the multi-junction cell as a whole, the standard means for maintaining high efficiency is to design the solar

cell such that the GaAs junction is current rich at the beginning of life (BOL) and becomes current matched to the InGaP at

the end of life (EOL), as shown in figure 6. Nevertheless, this compromises the BOL efficiency, so it is desirable to replace the

GaAs junction with a more radiation resistant material, such as InGaAsP [16].

NASA/CP—2005-213431 3
28

30

32

34

36

38

1 10 100 1000

Concentration Ratio

Efficiency (%)

Fig. 4. Efficiency of the 3J concentrator cell shown as a function of solar concentration. (From Ref. [13])

0.85

0.9

0.95

1

1E+13 1E+14 1E+15 1E+16

1MeV Electron Fluence (cm -2 )

Remaining Factor of Voc

InGaAs

middle

Ge

bottom

InGaP

top

0.85

0.9

0.95

1

1E+13 1E+14 1E+15 1E+16

1MeV Electron Fluence (cm -2 )

Remaining Factor of Isc

Ge

bottom

InGaAs

middle

InGaP

top

Fig. 5. Radiation resistance of sub cells in a triple-junction cell. (From Ref. [13])

4 SHORT & LONG TERM TARGETS

It is worth noting that the triple junction concentrator solar cell is an important driving force in Japanese R&D on multijunction solar cells. A consortium supported by NEDO and comprised of Sharp Co., Daido Steel Co. and Daido Metal Co. are

working to demonstrate a 40% efficient solar cell under 500 suns illumination by the end of March 2006. As shown in figure 4,

efficiencies in excess of 36% were recently demonstrated from an InGaP/InGaAs/Ge concentrator solar cell [13]. It is very likely

that advances achieved under this programme will bring improvements to the AM0 multi-junction space solar cell, as there are

many shared problems which we now will discuss with reference to table 1.

The InGaP/GaAs band-gap combination is not ideally suited to the AM0 solar spectrum, the InGaP being current rich and

the GaAs current poor. In the past, the situation has been helped by thinning the top cell, allowing unabsorbed sunlight to

penetrate to the lower junction [17]. This raises the current in the lower junction and thus the overall current in the multijunction cell, but the top cell open-circuit voltage is compromised. A preferable, but more challenging approach is to find a

higher band-gap material to replace the InGaP material, thereby increasing the photocurrent in the lower junction and increasing

NASA/CP—2005-213431 4
15

20

25

1E+13 1E+14 1E+15 1E+16

1MeV Electron Fluence (cm-2)

Photocurrent (mA/cm

2)

Ge

InGaAs

InGaP

BOL EOL

3J Cell

Fig. 6. Photogenerated current in each sub cell as a function of 1MeV electron fluence; the overall triple junction cell current

is shown by the broken line. (From Ref. [13])

0

2

4

6

8

10

12

0 0.4 0.8 1.2 1.6

VOLTAGE (V)

CURRENT (mA)

AlInGaP 2.5μm

InGaP 0.6μm

AM1.5G

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

300 400 500 600 700 800

WAVELENGTH (nm)

QUANTUM EFFICIENCY (%)

InGaP

(0.6μm)

AlInGaP

(2.5μm)

Eg=1.96eV

Eg=1.87eV

Fig. 7. Light-IV and QE curves for a thick AlGaInP cell and a thin GaInP cell under AM1.5G with no AR coating. (From

Ref. [13])

the top-cell open-circuit voltage. A recent demonstration of this is shown in figure 7, comparing a thick 1.96eV AlInGaP device

to a thin 1.87eV InGaP device. The short-circuit currents are identical but the AlInGaP device yields a higher open-circuit

voltage [13]. It is anticipated that this improvement alone will raise the multi-junction AM0 efficiency from 29% to almost 31%.

The Ge junction provides a fairly robust substrate for the multi-junction cell, but its properties as a third junction are far

from ideal. The band-gap is too low for efficient conversion of the sub-GaAs photon flux and it would be preferable to use a

material with a band-gap of around 1eV. Considerable effort has been invested in dilute nitride materials such as GaInNAs for

this purpose and although results to date have been poor on account of low minority carrier diffusion lengths [18], a greater

understanding of the material and its defects is emerging [19]. Recent progress in growing high-quality lattice mismatched

materials may offer a shorter term solution, to the problem of a 1eV junction. While the efficiencies of lattice-matched devices

NASA/CP—2005-213431 5
0

200

400

600

800

1000

1200

1400

0 10 20 30 40 50 60

AM0 Efficiency (%)

MJ/Ge(5.5mil)/rigid Thin/rigid

MJ/Ge(5.5mil)/light MJ/Ge(2mil)/light

MJ/Si(2mil)/light Thin-Film/SS(1mil)/light

Thin-Film/Poly(1.5mil)/light

Specific Power (W/Kg)

Fig. 8. Specific power vs BOL AM0 efficiency for complete space panels. The filled circles represent conventional rigid panels,

while the open symbols represent a light-weight panel technology where the cell weight is the dominant factor.

still exceed those of lattice-mismatch devices, the flexibility offered by the lattice mismatch approach is very attractive. For

example, encouraging results on an ambitious 5-junction lattice-mismatched device were recently reported [20].

The substrate material also plays an important role. Traditionally, the choice of GaAs and Ge substrates have constrained

the multi-junction solar cell to a limited range of materials, due to the usual requirement for lattice matching. However, the

recent progress in growing high quality lattice mismatched materials noted above, frees the choice of substrate to some extent.

While Ge provides a more robust substrate than GaAs, further gains could be made by replacing the Ge substrate with a Si. For

space, the mechanical strength of Si combined with the high-efficiency and radiation tolerance of the III-V cells would give a

highly attractive power to weight ratio when combined with a light-weight panel structure. This is illustrated in figure 8 where

the specific power for a complete panel is plotted as a function of BOL AM0 efficiency. The current rigid panel technology is

indicated by the filled circles, for both multi-junction cells on thick (5.5mil) Ge substrates and thin devices (e.g. 3mil Si). In this

case, the area of the rigid panel dominates the specific power, so multi-junction cells outperform the thin devices [21]. However,

if a lightweight panel technology is developed, then the overall specific power becomes dominated by the cell’s specific power.

Here the choice of substrate for the multi-junction solar cell becomes critical and a move to lighter substrates, such as 2mil Si,

is required for the multi-junction cells to maintain superior specific power over the less efficient thin-film devices.

In the longer term, it is worth considering if the multi-junction cell is the only means for achieving high efficiency solar

cells. In this respect, the next generation of high-efficiency solar cells is likely to draw from breakthroughs in semiconductor

materials. If a material were found where a third (intermediate) band exists in addition to the usual conduction and valance

band, then it would be possible to surpass the efficiency of a 2 junction solar cell [22,23]. Various hot carrier approaches have

been suggested, some maintaining a high voltage by sustaining the hot carrier population for longer than the carrier collection

time [24] with others yielding a high current through impact ionization [25]. The physics of these hot carrier proposals are

closely related [26] and yield identical limiting efficiencies. It is interesting to note that some of these concepts were proposed

almost simultaneously to the multi-junction solar cell [27], so while developments in compound semiconductor technology have

lead to the InGaP/GaAs/Ge multi-junction cell, so it is hoped that advanced semiconductor materials may lead to a practical

implementation of these photovoltaic concepts.

NASA/CP—2005-213431 6
5 SUMMARY

Recent progress in high-efficiency multi-junction solar cells has been reviewed, together with the technological improvements

that have led to InGaP/InGaAs/Ge multi-junction solar cells with efficiencies of 29.2% under AM0 and 31.7% under AM1.5G. In

a mechanical stack configuration, a GaInP/GaAs/InGaAs device holds the world record efficiency of 33.3% under 1 sun AM1.5G.

Under concentrated light, efficiencies in excess of 36% have been demonstrated between 100-500X. Near term improvements

include the use of an AlInGaP top cell which is projected to raise the AM0 efficiency to 31%.

REFERENCES

[1] Masafumi Yamaguchi. Super-high efficiency III-V tandem and multijunction cells. Clean electricity from photovoltaics,

1:347–375, 2001.

[2] J.M. Olson, R.K. Ahrenkiel, D.J. Dunlavy, B. Keyes, and A.E. Kibbler. Ultralow recombination velocity at Ga0.5In0.5P/GaAs

heterointerfaces. Appl.Phys.Lett, 55(12):1208–1210, 1989.

[3] M-J. Yang, M. Yamaguchi, T. Takamoto, E. Ikeda, H. Kurita, and M. Ohmori. Photoluminescence analysis of InGaP top

cells for high-efficiency multi-junction solar cells. Solar Energy Materials and Solar Cells, 45:331–339, 1997.

[4] H. Sugiura, C. Amano, A. Yamamoto, and M. Yamaguchi. Double heterostructure GaAs tunnel junction for AlGaAs/GaAs

tandem solar cell. Jpn.J.Appl.Phys, 27:269, 1998.

[5] T. Takamoto, M. Yamaguchi, E. Ikeda, T. Agui, H. Kurita, and M. Al-Jassim. Mechanism of Zn and Si diffusion from a

highly doped tunnel junction for InGaP/GaAs tandem solar cells. J.Appl.Phys., 85(3):1481, 1998.

[6] T. Takamoto, T. Agui, E. Ikeda, and H. Kurita. High efficiency InGaP/In0.01Ga0.99As tandem solar cells lattice matched

to Ge substrates. Solar Energy Materials & Solar Cells, 66:511–516, 2001.

[7] T Takamoto. Lattice matched solar cell and method for manufacturing the same. U.S. Patent, 6,300,558, 2001.

[8] T. Takamoto, T. Agui, E. Ikeda, and H. Kurita. High efficiency InGaP/InGaAs tandem solar cells on Ge substrates. In

Proc. 28th IEEE PV Specialists Conference, pages 976–981, USA, 2000. IEEE.

[9] D. Jung, C. A. Parker, J. Ramdani, and S. M. Bedair. AlGaAs/GaInP heterojunction tunnel diode for cascade solar cell

application. J. Appl. Phys., 74(3):2090–2093, 1993.

[10] T. Takamoto, E. Ikeda, H. Kurita, T. Tanabe, S. Tanaka, H. Matsubara, Y. Mine, S Takagishi, and M. Yamaguchi.

InGaP/GaAs and InGaAs mechanically stacked triple-junction solar cells. In Proc. 26th PV Specialists Conference. IEEE,

1997.

[11] K Takahashi, S. Yamada, Y. Minagawa, and T. Unno. Characteristics of Al0.36GaAs/GaAs tandem solar cells with pp−n−n

structural AlGaAs solar cells. Solar Energy Materials and Solar Cells, 66:517, 2001.

[12] T. Yamada, A. Moto and S.Tanaka, T.Tanabe and S.Takagishi. to be published.

[13] T. Takamoto, T. Agui, K. Kamimura, M. Kaneiwa, M. Imaizumi, S. Matsuda, and M. Yamaguchi. Multi-junction solar

cell technologies - high efficiency, radiation resistance and concentrator applications. In Proc. 3rd World Conference and

Exhibition on Photovoltaic Energy Conversion, 2003.

[14] Aurangzeb Khan, Masafumi Yamaguchi, Jaques C. Bourgoin, and Tatsuya Takamoto. Thermal annealing study of 1MeV

electron-irradiation induced defects in n+p InGaP diodes and solar cells. J.Appl.Phys, 91(4):2391–2397, 2002.

[15] Nethaij Dharmarasu, A. Khan, Masafumi Yamaguchi, Tatsuya Takamoto, Takeshi Ohshima, Hisayoshi Itoh, Mitsuru

Imaizumi, and Sumio Matsuda. Effects of proton irradiation on n+p InGaP solar cells. J.Appl.Phys., 91(5):3306–3311,

2002.

[16] Nethaij Dharmarasu, Masafumi Yamaguchi, Aurangzeb Khan, Takashi Yamada, Tatsuya Tanabe, Shigenori Takagishi,

Tatsuya Takamoto, Takeshi Ohshima, Hisayoshi Itoh, Mitsuru Imaizumi, and Sumio Matsuda. High radiation-resistant

InGaP, InGaAsP and InGaAs solar cells for multijunction solar cells. Appl.Phys.Lett., 79(15):2399–2401, 2001.

[17] S.R. Kurtz, P. Faine, and J.M. Olson. Modeling of 2-junction, series-connected tandem solar-cells using top-cell thickness

as an adjustable-parameter. J. Appl. Phys., 68(4):1890–1895, 1990.

[18] J.F. Geisz and D.J. Friedman. III-N-V semiconductors for solar photovoltaic applications. Semicond. Sci. Technol., 17:769–

777, 2002.

[19] S.B. Zhang and S. Wei. Nitrogen solubility and induced defect complexes in epitaxial GaAs:N. Phys.Rev.Lett., 86(9):1789–

1793, 2001.

NASA/CP—2005-213431 7
[20] F. Dimroth, C. Baur, M Meusel, S. van Riesen, and A.W. Bett. Multi-junction III-V solar cells for space applications. In

Proc. 3rd World Conference and Exhibition on Photovoltaic Energy Conversion, 2003.

[21] Navid S. Fatemi, Howard E. Pollard, Hong Q. Hou, and Paul R. Sharps. Solar array trades between very high-efficency

multi-junction and Si space solar cells. In Proc. 28th IEEE PV Specialists Conference, pages 1083–804, USA, 2000. IEEE.

[22] A. Luque and A. Mart ́ı. Increasing the efficiency of ideal solar cells by photon induced transitions at intermediate levels.

Phys.Rev.Lett., 78(26):5014–5017, 1997.

[23] P. W ̈urfel. Improvement of solar cell efficiencies by impurity transitions. Sol. Energy Mater. Sol. Cells, 79(1):153–161,

2003.

[24] Robert T. Ross and Arthur J. Nozik. Efficiency of hot-carrier solar energy converters. J.Appl.Phys, 53:3813–3818, 1982.

[25] P. W ̈urfel. Solar energy conversion with hot electrons from impact ionisation. Sol. Energy Mater. Sol. Cells, 46(1):43–52,

1997.

[26] Jenny Nelson. The physics of solar cells. Imperial College Press, London, U.K., 2003.

[27] M. Wolf. Limitations and possibilities for improvement of photovoltaic solar energy convertors. Proc. IRE, 48:1246–1263,

1960.

NASA/CP—2005-213431 8
MULTI-JUNCTION GaAs PHOTOVOLTAIC ASSEMBLY TECHNOLOGY 

FOR VERY SEVERE LEO ENVIRONMENT MISSIONS 

A. Caon, C. Signorini 

ESA/ESTEC, Noordwijk, The Netherlands, 2200AG 

A. Bals, K. Dettlaff 

EADS Astrium, Munich, Germany, 81663 

Abstract

The paper reports the results of a development and testing programme dedicated to photovoltaic assembly 

technology, based on multi-junction GaAs/Ge solar cells, for the very severe low earth orbit environments typical 

of some earth observation missions. 

The selection of material, components and technologies are discussed, with respect to the environmental 

requirements: 

• a combination of high number of thermal cycles (about 25,000) and extreme thermal conditions (between 

+160°C and -170°C); 

• all parts and materials had to be ATOX (Atomic Oxygen) resistant or had to get an ATOX protection. 

In particular, the paper reports about the optimisation of the welding process on two types of ATOX resistant solar 

cell interconnectors, gold (Au) and silver plated molybdenum (MoAg) and the evaluation of an ATOX protection 

process for the panel structure front side (Kapton face sheet). 

Finally, the cell characterization and the extensive environmental tests performed at cell, SCA (Solar Cell 

Assembly) and coupon level are described. 

The results of the evaluation show how the MoAg interconnector passed all tests successfully, while the Auinterconnector requires further optimisation for very severe LEO environments

1. INTRODUCTION 

For orbits below 650 km the density of Atomic Oxygen is such, that ATOX sensitive parts and materials of the 

solar array must be either replaced by ATOX resistant materials or they have to be protected by an ATOX 

resistant coating. As far as EADS Astrium solar arrays are concerned, special care has to be taken for the 

insulating 50 μm Kapton top-layer of the panel substrate and for the solar cell interconnectors and bus bars which 

are made of silver (Ag) for all Geostationary Earth Orbit (GEO) applications and all Low Earth Orbit (LEO) 

applications in orbits higher than 650 km. 

In this R&D activity, as a reference, it was considered an orbit characterized by an altitude of 250km, an 

inclination of 96.5° and an Atomic Oxygen fluence of 1023 cm-2 per year; a mission lifetime of about 3 years was 

taken into account. 

In addition to the ATOX environment, the solar array components have to withstand a high number of thermal 

cycles (up to 25.000) in a temperature range which goes up to +160°C / -170°C. 

In the following, some design solutions, compatible with the very sensitive GaAs Triple Junction solar cells, are 

presented. The design was tested on several coupons with different panel substrates, which were submitted to 

thermal cycling. 

NASA/CP—2005-213431 9
2. PROCESSES AND MATERIALS 

2.1 ATOX Protection of Kapton 

There are ATOX resistant Kapton foils on the market, however it was decided to use a standard Kapton foil and to 

protect it with an ATOX resistant silicone adhesive; this process is very flexible because it allows the use of a 

standard panel skin. A homogeneous layer of siliconic adhesive was obtained using a glass fibre sheet of the 

same thickness (100μm) and bonding it with the siliconic resin on the Kapton layer by means of a complex curing 

process; this adhesive (RTV S 691 from WACKER) is used for the solar cell lay-down and is very well known. This 

additional layer of glass fiber and adhesive has an additional positive effect: it helps to decouple the sensitive 

solar cells from possible stresses caused by the panel structure. 

2.2 ATOX resistant solar cell interconnectors and bus bars 

The standard interconnector used at EADS Astrium is a 12.5 μm thick Ag interconnector which is produced 

(punching and stress relief bending) out of a silver ribbon by a fully automated solar cell - interconnector welding 

machine. The most cost effective solution, which is considered, consists of replacing the ATOX sensitive silver by 

ATOX resistant gold and to keep the interconnector manufacturing process unchanged. These Au interconnectors 

were successfully tested and used in the past on 200 μm thick Silicon solar cells and 200 μm thick single junction 

GaAs/Ge cells for temperature ranges of +100°C / -100°C and up to 25000 thermal cycles. However, there were 

indications that with the much more severe temperature range in combination with the 160 μm thick GaAs Triple 

Junction cells the Au interconnectors with the actual design might go over there limit. 

A more expensive solution is a MoAg interconnector, which has already been used on the HUBBLE SPACE 

TELESCOPE solar array and on several other Silicon (Si) solar arrays for LEO missions. This interconnector 

consists of 15 μm Mo and a 5μm Ag plating on each side; the silver plating is necessary to make the 

interconnector weldable. After welding and lay-down of the solar cells the weld area is protected by the RTV 

laydown adhesive while the exposed Ag will be eroded by the ATOX and will flake away; for missions with very 

sensitive optical instruments which do not allow this flaking, it is possible to get the interconnectors with the Ag 

etched away in areas which are not used for welding. 

These robust MoAg interconnectors had never been used before on the much more sensitive 160 μm GaAs/Ge 

Triple Junction cells. To reach a good yield for the interconnector welding operation, it was necessary to optimise 

the weld electrodes as well as the weld parameters (electrodes pressure, weld duration and energy). 

The test hardware has been equipped with both Au and MoAg interconnectors. 

At the solar cell string ends in the standard design for high LEO and GEO, a 50 μm thick, ATOX sensitive, Ag bus 

bar with stress relief loops is used as an interface to the string cabling. 

As an alternative a 50μm Au bus bar with the same design as the standard Ag bus bar was selected and tested. 

2.3 Type of tested solar cells 

The welding optimisation activity was focused on different solar cells: 

• The GaInP2/GaAs/Ge Triple Junction cells named GAGET1 from RWE Space, which were based on an 

EMCORE epitaxy and which had an integral protection diode which is protecting the adjacent cell. 

• The GaInP/GaAs/Ge Triple Junction cells named GAGET 2 ID 2 cells from RWE Space, which were 

based on a Spectrolab epitaxy. 

The several coupons which were submitted to long term thermal cycles were equipped with the GaInP/GaAs/Ge 

Triple Junction cells named GAGET 2 ID cells from RWE Space, which were based on a Spectrolab epitaxy. 

One additional coupon equipped with GaInP/GaAs/Ge Triple Junction cells named ATJ solar cells, from EMCORE 

and characterized by a monolithic Schottky protection diodes was added in the thermal cycling test. 

NASA/CP—2005-213431 10
3. TESTING AND PRE-QUALIFICATION OF CELLS, SCA’S AND COUPONS 

3.1 Testing of cells and SCA’s

The purpose of this task was to verify the optimisation of the parallel gap resistance welding process applied to 

the above mentioned Triple Junction GaAs based solar cells and the Gold and Silver plated Molybdenum 

interconnectors. Several iterations, based on the adjustments of the welding process including the electrode 

design and the verification testing, were done to achieve a fine tuning of the existing welding process to the multijunction solar cells. 

The subsequent validation of the complete solar cell assembly process was performed on the basis of the 

following test plan: 

• Visual Inspection of welded solar cell and SCA 

• Electrical Performance measurements of cell and diode 

• SAM (Scanning Acoustic Microscope) investigation of weld spots 

• Pull test at Beginning of Test (BOT) on cells and SCAs 

• Pull test at End of Test (EOT) on cells and SCAs 

• Up to 7500 thermal cycles –100°C / 160°C 

The Test Articles were 10 GAGET 1 solar cells integrated with Au interconnectors (see Fig.1) and 20 GAGET 2 ID 

solar cells integrated both with Au and with MoAg interconnectors (see Fig.2); the used coverglasses were the 

CMX100 μm for the GAGET 1 solar cells and the CMO 100 μm for the GAGET 2 solar cells. The size of the cells 

was 40 mm x 80 mm with cropped corners. 

The results of the testing activity showed the optimization of the welding processes both for the Au and MoAg 

interconnectors. In particular, the visual inspection and the electrical performance measurements of the cells 

highlighted the welding process and the thermal loads derived from the thermal cycling do not cause any 

degradation of the cells, neither mechanical (e.g. cracks) nor electrical. Furthermore, the SAM pictures 

demonstrated the excellent repeatability of the welding and the non degradation of the welding joints in the 

thermal environment; finally, the pull force values (4N for Au interconnectors at 90° and 10N for MoAg at 0° with 

respect to the solar cell plane) showed the good quality of the joints and the non degradation of them in the 

thermal environment. 

FIG.1: GAGET 1 SCA with Au interconnector 

NASA/CP—2005-213431 11
FIG. 2: GAGET 2 ID with MoAg interconnector 

3.2 Test of Coupons 

3.2.1 Pre-Testing 

In a first pre-test step, 4 coupons characterized by different sizes, substrate and interconnector types have been 

build and tested. The aim of this test was to check the capability of all components and processes (including the 

demonstration of SCA and bus bar repair) to withstand a thermal environment characterized by 6,030 thermal 

cycles (30 cycles +160°C/-170°C, 6000 cycles +160°C/-100°C) and to verify their electrical and mechanical 

behaviour. Coupons 1 & 4 have been submitted to additional 9000 cycles reaching the 15000 cycles. 

The test coupons were equipped as follows: 

FRONT SIDE 

• GaAs Solar Cells RWE (GAGET I) (80 mm x 40 mm, 160 μm, with integrated diode, cropped corner) 

• GaAs Solar Cells RWE (GAGET II) (80 mm x 40 mm, 160 μm, with integrated diode, cropped corner) 

• One diode strip per coupon 

• Coverglass from Thales Space Technology (TST) CMO 100 AR 

• Gluing spot for solar cells to structure, 

• Cell Interconnectors 12,5 μm gold and 25 μm MoAg 

• Bus bars 50 μm gold 

• Kapton insulated wire type 

• APL (Atox Protection Layer) consisting of RTV-S 691 and glass filament 

• The front side layout was characterized by one complete string of 14 cells (Coupon 1,4), 1 cell (Coupon 2) 

NASA/CP—2005-213431 12
and 8 cells (Coupon 3) 

• The strings of the Coupons 1-3 and the Coupon 4 (String B+C) were manufactured with 12.5μm gold cell 

interconnector and 50μm gold bus bars while the string A of the Coupon 4 was realised with 25μm MoAg cell 

interconnector 

REAR SIDE 

• Kapton insulated wire type AWG 20 and 22 

• Blocking Diode 1N5417 

• Bleed Resistor RWR 89 2,8 kΩ

• LRB (Line Reduction Board) 

• Cable clamps with conductive gluing spots 

After thermal cycling the ATOX Test (Atomic Oxygen Exposure) was performed on Coupon 2 (Au interconnector). 

The aim of the test was to determine if there would be any degradation in the materials on the surface of the array 

after exposure to a reduced atomic oxygen fluence of approximately 2x1020 atoms cm -2. The average beam 

velocity during the exposure was 7.4 ± 0.4 kms-1, the background pressure was approximately 5 x 10–5mbar and 

the temperature of the sample plate was 23±3°C. 

Using high resolution optical imaging and scanning electron microscopy, no discernible degradation of the 

exposed materials due to atomic oxygen erosion could be detected at the magnifications used (see Fig.3). 

The ATOX resistivity of the MoAg interconnector was already demonstrated in previous programs. 

Fig. 3 Microanalysis spectrum of the surface of an Au-contact 

After completion of all testing, the electrical performance did not change more than 3% at operational voltage, 

(except for the strings with gold interconnectors on coupon 4 showing some cracks on some interconnector 

fingers). All parts, except the Au interconnector on the coupons have passed the pre-test program successfully. 

NASA/CP—2005-213431 13
3.2.2 Pre-Qualification Testing with Coupons 5 & 6

The aim of this test was to demonstrate the capability of all components, processes (including the demonstration 

of SCA repair) to withstand a thermal environment extended up to 22,545 thermal cycles of which 45 cycles with 

extreme high/low temperature of +160°C/-170°C and 22,500 cycles with a temperature range tailored for a 

specific mission (1800 cycles +160/-100°C, 7200 cycles +150°C/-100°C, 900 cycles +160°C/-80°C, 3600 cycles 

+150/-80°C, 1800 cycles +160/-0°C, 7200 cycles +150/-0°C). 

The 2 coupons represented a flight panel especially designed for high temperatures with flight parts, processes 

and electrical connection that could be used for future LEO solar array. The front side of the coupon 5 (see Fig.4) 

was covered with RWE GaAs Triple Junction (GAGET II) solar cell assemblies and for the coupon 6 EMCORE 

Triple Junction cells have been used (see Fig.5). On both coupons, the SCAs had both Au interconnectors and 

MoAg interconnectors (see pictures below). 

The cell lay down was divided into two strings. The string 1 was manufactured with Au interconnector while the 

string 2 with MoAg interconnector; the termination tabs to the string ends were made of Au. 

Fig.4: Front side of Coupon 5 

NASA/CP—2005-213431 14
Fig.5: Front side of Coupon 6 

The rear side of coupon 5 (see Fig.6) was covered with harness, diode, resistor, LRB and cable clamps. 

Fig.6: Rear side of Coupon 5 

NASA/CP—2005-213431 15
The detailed list of the components of the two coupons is reported here below: 

FRONT SIDE 

• 13 GaAs Solar Cells GAGET II RWE/SPL, 40 mmx80 mm, 160 μm with cropped corner, BOL Efficiency (Eff.) 

of 26.8% at 28°C with integrated diode on Coupon 5 

• 2 diode strip on Coupon 5 

• 14 GaAs Solar Cells ATJ Emcore, 40 mmx80 mm, 160 μm, with cropped corner, BOL Eff. 27.0% at 28°C with 

monolithic Schottky diode 

• Coverglass TST CMO 100 AR (40.15 mm x 80.15 mm, 100 μm, with cropped corner) 

• Gluing spot for solar cells to structure, thickness 80 μm 

• Interconnector Au 12.5 μm 

• Interconnector MoAg 25 μm 

• Bus bars Au 

• Kapton insulated wire type AWG 24 (SPL 10-24), 22 (SPL 10-22) 

• ATOX protection layer (RTV S691 + glasfilament) 

REAR SIDE 

• The rear side cable harness was manufactured with different size of American Wire Gauge (AWG 20/22 

connected to the diode (welded) resistor (welded), LRB (welded) and cable clamps in welding and crimping 

version (for testing leads only). 

• Kapton insulated wire type AWG 20 (SPC 10-20) and 22 (SPL 10-22) 

• Blocking Diode 1N5417 

• Cable clamps with conductive gluing spots 

• Resistor RWR 89 2.8 kΩ

• LRB 

The coupons have been submitted to the following tests: 

• Inspections and electrical performance measurements 

• Insulation measurements & continuity measurements 

• 5 Thermal vacuum cycles –170°C / 160°C (before and the thermal cycling) 

• 45 Thermal cycles –170°C / 160°C

• 22500 Thermal cycles up to–100°C / 160°C 

The main results are summarized here below: 

After completion of all testing the electrical performance measurement did meet the requirement only for the 

MoAg string; the value changed for the string with Au-interconnector more than 3% at operational voltage due to 

some cracks on some interconnector fingers. 

A new replacement method for SCAs with MoAg interconnector was included and passed the pre-qualification 

successfully. 

One cell hairline crack has been detected, which had no impact on the electrical performance. 

All parts, except the Au-interconnectors, on the coupons (front and rear side) have passed the pre-qualification 

program successfully. 

NASA/CP—2005-213431 16
4. Conclusions 

The test performed at the end of the development programme demonstrated that the assembly technology 

applied to the multi-junction GaAs/Ge solar cells has reached an excellent level of confidence, for very demanding 

low earth orbit environment. 

The welding process has been optimised for the Triple Junction GaAs solar cells with both Au and MoAg 

interconnectors. 

The design of the MoAg interconnector has withstood without any failure or degradation the thermal fatigue 

derived from the considered LEO environment testing. 

The actual design of the Au interconnector has shown a limit of applicability with respect to the chosen cell 

thickness and panel substrate configurations, in combination with the considered LEO environment; on the basis 

of these results, further optimisation of the Au interconnector design should be considered in the future R&D 

activities, in order to extend its potential use for these types of solar array and missions. 

The ATOX protection process applied to the kapton insulation layer of the panel front side has been successfully 

tested. This method is very flexible because it allows the use of a standard panel skin and provides an additional 

positive effect because it helps to decouple the sensitive solar cells from the possible stresses caused by the 

panel structure. 

Finally, it is important to underline that the selected Multi-Junction solar cells and photovoltaic assembly 

processes have been demonstrated to be compatible with the chosen panel structures in a very demanding LEO 

environment. 

NASA/CP—2005-213431 17
CURRENT STATUS OF PHOTOVOLTAIC DEVELOPMENT AT AFRL 

John Merrill, Paul Hausgen, Donna Senft, and Pawel Tlomak 

Air Force Research Laboratory, Space Vehicles Directorate 

Kirtland AFB, NM USA 87117-5776 

 ABSTRACT: Advances in design and manufacturing technology of crystalline multijunction solar cells have 

continued to push efficiencies higher. Triple junction cells with 28% efficiency are now available to the user and 30 % 

prototype cells have been demonstrated. Development of thin-film solar cell arrays is being aggressively funded for nextgeneration high power space platforms. Initial modules of thin-film photovoltaics comprised of amorphous silicon and 

Cu(In,Ga)Se2 devices have been demonstrated. Space-compatible technologies for thin-film cell integration, contacts, 

and protective coatings are being developed. Cell-level development efforts aimed at increasing performance are 

continuing and progress is being made toward a 15% efficiency goal for cells on lightweight, flexible substrates. High 

temperature polymer substrates for Cu(In,Ga)Se2 cells have been demonstrated. Thin-film solar arrays are promising for 

3-5 times reductions in specific power (W/kg), array cost ($/W), and stowed volume (W/m3) compared to state-of-the-art 

rigid panel solar arrays. A proposed thin-film solar cell flight experiment is also discussed. 

1. INTRODUCTION 

Rigid panel arrays populated with crystalline solar cells are currently the only practicable power generation option for earth 

orbit missions, with multijunction crystalline solar cells the primary option for DoD and most commercial spacecraft. Air 

Force Research Lab (AFRL) multijunction solar cell programs continue to increase the efficiency of these cells at a steady rate.

Multijunction solar cells can be directly retrofitted into existing arrays and offer increased power, mass, and fairing volume 

budgets for spacecraft designers. Triple junction cells with 28 % efficiency [1] are now commercially available. Impact of 

efficiency increase is specific to a given mission/spacecraft design, but one recent study showed a 30 % increase in efficiency

(27 to 35 %) would result in ~ 70 % increase in power available to the payload. 

FTFPV (Flexible Thin-Film Photovoltaics) have been touted as the next-generation technology for space power [2,3] for 

their advantages over crystalline multijunction cells. With launch costs exceeding $10,000/kg to LEO, the specific power of 

the solar array is often a critical metric. Current solar arrays comprised of multijunction solar cells grown on single crystal

wafers, mounted on Al honeycomb panels, and protected with a relatively heavy cover glass have specific power levels of less 

than 100 W/kg. In addition, multijunction solar cell arrays can create problems in terms of stowage volume for high power 

spacecraft (8-13 kW/m3

). Thin-film photovoltaic arrays, where the photovoltaic materials are deposited on flexible, lightweight 

metal foils or polymers, are capable of an EOL specific power of 235 W/kg or greater. Thin-film arrays are ultimately 

projected to have efficiencies of 15% and a stowed volume on the order of 45-60 kW/m3

.

FTFPV technology has advantages over state-of-the-art technology in terms of cost and radiation resistance in addition to 

mass and stowage volume. Thin-film deposition processes are low-cost and have been developed for the terrestrial market, 

where the long-term cost target is $1/W. The two major types of thin-film solar cells are amorphous silicon (a-Si) and 

polycrystalline Cu(Ga,In)Se2 (CIGS). Space-qualified adaptations of these terrestrial products are more expensive than their 

terrestrial counterparts, but costs are not expected to exceed $100/W for FTFPV arrays. Multijunction panel costs typically 

exceed $400-500/W. CIGS solar cells in particular have been found to be radiation resistant, with post-irradiation results that

compare to InP. 

Lightweight FTFPV arrays will be enabling for certain classes of high power satellites and microsatellites. A 

microsatellite having a total mass of less than 100 kg and a power requirement of 650 W, for example, will need to allocate 

14% of its total mass to the solar array using a crystalline multijunction array. A thin-film array with a conservative specific

power of 255 W/kg could reduce the requirement for the solar array to 2.5% of the total mass. For a very large 100 kW system, 

a thin-film solar array would reduce the mass by over 1700 kg and the stowage volume by over 8 m3; this mass and volume 

could be reallocated to the payload. 

To realize the full advantages from FTFPV technology, first the technology to construct space qualified modules from 

thin-film cells must be developed. Second, specialized lightweight array support structures must be designed and developed. 

Finally, attention must be paid to the issues of space environmental effects on thin-film photovoltaics and the protective layers 

necessary to ensure long-term reliable operation. 

NASA/CP—2005-213431 18
2. MULTIJUNCTION SOLAR CELL DEVELOPMENT 

State-of-the-art multijunction solar cells are now commercially available from EMCORE Corp. and Spectrolab, Inc. with 

nominal efficiencies of 28%. Prototype triple junction cells (4 cm2

), also with the GaInP/GaAs/Ge device structure, have been 

demonstrated at efficiencies of up to 30%. AFRL programs are pursuing multiple technical approaches to increase cell 

efficiency beyond 30%. Lattice-mismatch GaInp/GaInAs/Ge solar cells have been demonstrated at 28.8% efficiency with an 

ultimate practical efficiency of 31.5% predicted. The second technical approach involves the addition of a 1.0 eV-bandgap 

nitride-material junction. Previously, AFRL programs have focused on developing InGaAsN material to be used in a 4 junction 

device. The material has proven to be current limiting in the 4-junction cell design due to defects in the material, which has 

prevented the 4-junction design from reaching the predicted performance levels. A 5-junction approach 

(GaInP/GaInP/GaAs/GaInNAs/Ge) splits the top junction into two separate junctions, yielding a higher open circuit voltage, 

while reducing the current to levels capable of being produced by the InGaAsN junction. The overall effect is higher 

efficiency, with projections of >34 % and an open circuit voltage of 4.7 V. 

3. THIN-FILM ARRAY DEVELOPMENT AND SPACE QUALIFICATION 

 The goal of AFRL thin-film solar cell programs is the development of the technology necessary to achieve reliable 

operation of thin-film arrays in space. There is still significantly more work to do to demonstrate a space qualifiable array 

based on thin-film photovoltaics. Multijunction solar cell development yields a drop-in replacement using existing solar array 

designs; to take advantage of the benefits of thin-film solar cells, new solar arrays must be developed which are lightweight but

have the stiffness necessary for array pointing and maneuvering. The challenges to accomplishing this goal are three-fold: 1) 

improvements in cell technology on lightweight substrates compatible with space use, 2) integration of cells into modules and 

the design of arrays, and 3) space qualification of the array blankets. AFRL has ongoing thin-film solar cell and module 

development programs focused on these objectives. The initial progress in the three areas of FTFPV blanket development and 

plans for future work are detailed below. 

 3.1 Cell Development 

 In contrast to crystalline wafer solar cells that were developed initially for space and eventually adapted to terrestrial 

applications, the thrust behind the development of thin-film solar cells has come from the terrestrial market, primarily because

of the potential for lower production costs. As a result, some commonly used thin-film substrates and materials are unsuitable 

for space applications. For example, CIGS modules with efficiencies of 12% are commercially available with 3.2 mm-thick 

glass substrates. Less fragile substrates such as metal foils and Kapton films are more suitable for space. EVA and other 

polymeric encapsulants have been developed to protect thin-film cells from the effects of weather, but are insufficient for 

shielding arrays from the space environment. 

 AFRL-funded cell development programs are intended to make changes in cell technology geared to space application by 

increasing cell specific power via tilting performance/cost tradeoffs towards higher efficiencies and supporting the 

development of cell production processes with lightweight substrates such as metal foils and polyimide. A summary of the cell 

development programs for the two most promising thin-film technologies, a-Si and CIGS alloys, is listed in Table I. 

 For a-Si, the more mature technology of the two, the challenges have been to increase efficiencies above the 10% 

threshold, to integrate cells into modules, and to use lightweight substrates. United Solar Systems (USS) has increased the 

efficiency of their product to 9.8% for 929 cm2

 (1 ft2) cells and to 12% for small cells by the addition of an internal reflector 

layer to their triple junction cells and by reducing the a-Si deposition rates.[4] At the same time, USS has decreased the weight

of their terrestrial product by thinning the 127 μm stainless steel substrate to 25 μm. These improvements have increased the 

specific power at the cell level to 680 W/kg. Similar efficiencies have been demonstrated on polymer substrates at the research

and development level. These polymer substrate cells have a cell level specific power >1000 W/kg. 

 CIGS technology, while being less developed than a-Si, has the advantage of higher efficiencies than a-Si with the promise 

of further increases. The main challenge has been to develop a high efficiency CIGS cell on metal foil substrates. Since high 

temperature processing (>500 oC) is currently necessary to yield the optimal efficiencies in CIGS, our efforts have 

concentrated on metal foil substrates. Efforts are also underway to develop high temperature polymer films that can withstand 

the (>500 oC) processing temperature. 

 International Solar Electric Technologies (ISET) has successfully achieved efficiencies of 10.2% for 10 cm2

 cells on Mo 

foil substrates using a non-vacuum process. DayStar Technologies has achieved efficiencies of 15.2% for 1.1 cm2

 cells and 

>11% for the larger 24 cm2

 cells using batch evaporation processing. ITN Energy Systems, using a roll-to-roll evaporation 

production line has achieved efficiencies of 11% for sample sizes of 0.68 cm2

 and 7% for 58 cm2 samples. 

 3.2 Cell Integration and Array Design 

 Integration and array design involves producing a prototype thin-film array from individual cells, including cell layout, 

electrical contacts and interconnects, wire harnesses, and method of array deployment. A lightweight thin-film array support 

structure is being designed under the AFRL PowerSail program. Under this program, special attention is being paid to array 

dynamics and deployment mechanisms that will enable >20 kW arrays. Under the PowerSail FTFPV blanket qualification 

program (Table I), a prototype thin-film blanket as shown in Fig. 1 was fabricated. CIGS cells from ISET and a-Si cells from 

USS on 25 μm-thick metal foil substrates were spot-welded together to form the blanket without the use of a separate blanket 

support material. Hinges were developed so that the blanket can be stowed in a Z-fold configuration. During the initial round 

of blanket development, the cells were tested for adhesion using tape tests and under thermal cycling conditions for GEO 

orbits. Additional designs for a roll-out array are being constructed. 

NASA/CP—2005-213431 19
More extensive array blanket development programs are continuing under Dual Use Science and Technology (DUS&T)

programs. These two 3-year programs with Lockheed Martin and Boeing Satellite Systems involve 50-50% cost sharing 

between industry and government and will focus on the design and development of FTFPV modules.

The Lockheed Martin DUS&T program, with sub-contractors ISET and USS, has identified 5 kW, 10 kW, and larger thinfilm array designs using stabilized 10-15% efficient cells. The module electrical architectures have been optimized and include

an appropriate number of bypass and blocking diodes. The program is working to optimize cells for space, and has generated

specifications for module strength and structural support requirements. The effort has resulted in the technology to construct a

32 ft2

 module comprised of 30 USS a-Si solar cells. Smaller 3-cell modules have also been fabricated. A 3-cell module and

individual cells have been cycled from -175 °C to +120 °C for 100 cycles and 1100 cycles, respectively. Cell efficiencies after

thermal cycling were within 97% of pre-test values.

The Boeing DUS&T program has identified 1kW, 20 kW, and 50 kW arrays suitable for thin-film blankets. The Boeing

program is also developing modules, their electrical architectures, and structural support requirements. Sub-contractors on the

Boeing program are USS, and ITN Energy Systems. The program has produced a comprehensive requirements document 

covering all aspects of the thin-film blanket, array and interconnect design.

Figure 1: Prototype for the FTFPV module design under the PowerSail blanket qualification program.

3.3 Space Qualification of FTFPV

Despite some limited flight tests of thin-film solar cells, including space qualification of USS cells by Fokker and KVANT

[5], more rigorous and comprehensive testing of FTFPV cells and modules is needed to provide design data for degradation

rates and to uncover any potential failure mechanisms. Crystalline multijunction solar cells are typically covered with glass

that protects the cells against low energy protons, atomic oxygen, and micrometeorite damage. Thin-film cells, to maintain

their flexibility and avoid the weight penalty of the glass will be protected with a multi-purpose protective coating, the

development of which is discussed below. Space qualification will involve testing of this encapsulant along with the modules it

covers.

One concern is that the integrated cells withstand thermal cycling. Steep thermal gradients can cause buckling in PV 

blankets, and interconnect strength must withstand the strain. Thermal cycling has also been responsible for the delamination

of cells from substrates and degradation in electrical contact materials, particularly conductor-filled epoxies.

Although thin-film solar cells have been shown to be more radiation-resistant than crystalline cells and radiation damage

has been found to anneal from a-Si cells at 70-80 oC, comprehensive tests to measure and model the radiation response of both

a-Si and CIGS [6,7] materials are needed. Previous radiation testing has shown that the radiation resistance of CIGS exceeds

that of a-Si.[8] In addition, both thermal and light-induced annealing rates have been quantified for CIGS cells on glass 

substrates.[8]

As part of space qualification, the AFRL has an on-going program for the evaluation of radiation hardness of FTFPV solar 

cells. The goal of this program is to develop a predictive model and to demonstrate survivability in the different space

environments defined for given missions. The model is expected to be able to predict the end-of-life performance of FTFPV 

solar cells for a given orbit. Planned experiments will expose both a-Si and CIGS solar cells to both electrons and protons at 

different energy levels and fluences. Experimental conditions are based on 1.) potential orbit environments, and 2.) data needed

to support model development. Experiments will measure rates of annealing, light assisted annealing, and combined affects.

Uncoated CIGS and a-Si cells from a variety of manufacturers have been irradiated by electrons and protons of several 

different energies. Irradiation experiments will be repeated for coated cells, using cells coated by two or more 

materials/processes. Results of the radiation and annealing tests are expected to be published in 2004. In addition to solar cell

behavior under irradiation, coatings being developed for FTFPV for environmental protection and thermal management (as 

discussed in the next section), will be subjected to similar radiation conditions. 

Additional space qualification tests for thin-film modules will measure their mechanical durability, resistance to ultraviolet

light and atomic oxygen, high temperature performance, and response to micrometeorite bombardment. The module 

performance will depend upon the quality of the protective coating. The protective layer for the modules must shield against

NASA/CP—2005-213431 20
low energy protons, have high emissivity in order to reduce operating temperatures, have high optical transparency, and be 

resistant to micrometeorite damage and atomic oxygen. In addition, thin-film arrays will operate at >100 V in order to reduce 

wiring mass. The coating will be required to be pinhole-free and possess a high dielectric constant to eliminate high voltage 

arcing. 

3.4 Coatings Development 

State-of-practice crystalline solar cell technology employs a cover glass to protect cells from on-orbit radiation 

environments, arcing, and atomic oxygen. This cover glass adds mass to the solar array and increases costs. One of the 

advantages of thin-film solar cells is that they are relatively resistant to radiation and thus do not require a thick cover glass. 

However, these technologies do require development of flexible, lightweight protective coatings that would provide thermal 

management and environmental protection in the space environment. Emphasis on thermal management is necessary due to the 

high absorption and low emissivity characteristic of materials used for FTFPV. The key to controlling the temperature of 

FTFPV is the spectral separation between the solar power input and the thermal emission in the infrared range. Specifically, 

more than 95% of the Air Mass Zero (AM0) solar spectrum lies between 0.3 – 2.5 microns, whereas more than 95% of the 

energy radiated by a blackbody lies between 3.5 – 35 microns. Consequently, a coating with maximum transparency in the 

visible/NIR range while having maximum absorption in the IR range will significantly and effectively lower the operating 

temperature of the FTFPV acting as a passive thermal control system. 

 To reach these goals, a comprehensive coatings development program, establishing a scientific base for modeling and 

fabrication of thin-film multifunctional protective coatings, industrial scale-up, and concurrent qualification testing, has been 

initiated. The coating development program emphasizes the multifunctional role of FTFPV protective coatings. The key 

elements of the program are: (1) Design of coatings that are multifunctional in nature, i.e. simultaneously provide 

environmental protection, maintain high optical transmission, provide thermal management, and prevent charging, (2) 

Selection of suitable, low-temperature coating deposition techniques with high deposition rates, (3) qualification testing, and

(4) scalability and ability to tailor to specific mission needs. Currently, three different coating processes using SiOx and Al2O3

materials are being developed. The processes are based on proprietary sputtering and/or hybrid methods. Both single layer and 

multilayer, multi-component coatings are being investigated. 

The initial results of the coating programs will be used to downselect a coating process which will then be scaled up for 

large area deposition and production. Initial coatings have achieved target emissivity and optical performance goals. 

Environmental tests of coated solar cells, including thermal cycling, mechanical durability, and high voltage plasma are 

scheduled to begin in 2003.

3.5 Proposed flight experiment for FTFPV

One of the obstacles to acceptance of FTFPV is developing flight heritage. While some thin-film cells have been flown in 

space [5] more comprehensive data is needed to instill confidence in the technology and understand combined effects. These 

tests along with ground testing have given us a promising view of the performance of these cells in a space environment, but 

combined effects cannot easily be fully explored in ground testing. In addition, while conceptual thin-film array deployment 

structures have been built and tested using offloading to simulate zero-g environments, no actual tests of these large structures 

has occurred in space. In order to mitigate these concerns, a relatively large space experiment involving FTFPV has long been 

desired.

The AFRL Advanced Power Group is collaborating with the PowerSail program on a proposed flight experiment that will 

incorporate a relatively large area (~130 m2

, or the equivalent of a 15 kW array using 8.5% efficient solar cells) thin-film array 

in a large deployable structures experiment (DSX). The array will include thin-film solar cells from as many as five different 

vendors as shown in Figure 2. The array is not intended to be fully populated, but will demonstrate power levels greater than 3

kW. The array will have two areas of high voltage (100V and 300V) to test breakdown and arcing, and the proposed orbit will 

include high radiation levels. The array cells will be coated using the technology being developed as discussed in section 3.4.

This flight will be the first large scale demonstration of FTFPV in space, and should pave the way for routine use of FTFPV 

arrays for space applications. Successful development and demonstration of FTFPV technologies will enable a new generation 

of lightweight high-power spacecraft. 

NASA/CP—2005-213431 21
Figure 2: Proposed DSX (Deployed Structures Experiment) lightweight thin-film array showing layout of thin-film solar cell

experiment.

4. CHALLENGES FOR THE FUTURE

With AFRL programs in-place for increased efficiency of multijunction crystalline solar cells, future direction will be

geared to “quantum leap” type technologies with significantly higher specific power, Watts/m3

, and/or efficiency. One 

approach to achieve the quantum-leap goals is to devise a method for the low cost deposition of crystalline high efficiency

cells onto flexible, lightweight blankets. Other approaches include polymer solar cells and spectrum alteration.

For thin-film solar cells, further improvements in specific power from efficiency increases, lighter substrates, and

monolithic integration of cells are all challenges. Efficiency increases are expected to arise primarily from CIGS technology

where efforts to increase the bandgap and the research into multijunction cells are underway in several laboratories. Further

thinning of metal foils and a switch to lower density metals may also produce further increases in specific power density due to

mass reduction. Progress is also being made in polymeric substrates for both a-Si and CIGS. Monolithic integration is expected 

to reduce solar blanket production costs while decreasing areal losses, wiring mass, and complexity of cell interconnections.

Monolithic integration requires either a dielectric substrate such as polyimide or an insulating layer on a metal foil. Large-scale

pinhole-free coatings on metal foils at low cost have been problematic in the past, but AFRL programs are making headway.

Monolithically integrated a-Si is currently being produced on 1 mil polyimide at Iowa Thin Film Technologies (ITFT). ITFT 

sub-modules have AM0 efficiencies of 4-5% for 0.93 m2 (10 ft2

) areas. Under an AFRL-funded program, ITFT is developing a

higher efficiency product with a target goal of 10%. The development of high temperature polymers promises to enable

monolithic integration for CIGS cells on flexible substrates.

References and notes: 

[1] All quoted solar cell efficiency values are as measured with the Air Mass Zero or AM0 solar spectrum. 

[2] J.W. Tringe, "Trends in Thin Film PV Technology Development" 

[3] J. Tringe, J. Merrill, and K. Reinhardt, IEEE Photovoltaic Specialists Conference, 2000, 28, pp. 1242-1245.

[4] K.C. Palinginis, J.D. Cohen, S. Guha, J.C. Yang, Phys. Rev. 63, 2001, 201203.

[5] M. Kagan, V. Nadorov, S. Guha, J. Yang, and A. Banerjee, IEEE Photovoltaic Specialists Conference, 2000, 28, pp. 

1261-1264.

[6] U. Rau, A. Jasenek, H.W. Schock, et al, IEEE Photovoltaic Specialists Conference, 2000, 28, pp. 1032-1037.

[7] A. Boden, D. Bräunig, J. Klaer et al, IEEE Photovoltaic Specialists Conference, 2000, 28, pp. 1038-1041.

[8] A. Jasenek, U. Rau, H. Weinert, H.W. Schock, and J.H. Werner, 3rd WCPEC Proceedings, 2003. 

NASA/CP—2005-213431 22
Table I: Thin-film cell development and module/array development efforts funded by the AFRL as part of the space FTFPV 

array development program. 

Company or Program Name Program Technology Substrate

Blanket Qualification for PowerSail FTFPV Array Development a-Si and CIGS Metal foils 

Boeing Satellite Systems FTFPV Module Development a-Si and CIGS Metal foils 

Lockheed Martin FTFPV Module Development a-Si and CIGS Metal foils 

United Solar Systems Performance enhancements & monolithic 

integration on polyimide a-Si Stainless steel and 

polyimide 

ITN Energy Systems Sputtered fabrication process CIGS Stainless steel 

Iowa Thin Film Technologies (ITFT) Monolithic integration and efficiency 

improvements a-Si Polyimide 

International Solar Electric 

Technologies (ISET) 

Non-vacuum painted-on process & 

lightweight substrates CIGS Polymer 

Florida Solar Energy Company 

(FSEC) 

Increased bandgap and decreased 

temperature coefficient CIGSS Metal foil 

Foster-Miller

ITN Energy Systems High temperature polymer substrates CIGS Polymers 

Triton

Sommer Materials 

Dielectric coatings to enable monolithic 

integration CIGS Metal foil

ITN Energy Systems Tandem CIGS cells CIGS Metal foil 

Encapsulant Development Multi-functional protective coating 

development a-Si and CIGS Metal foils and polymers 

Grain Boundary Orientation Studies Study effect of grain boundary size and 

orientation CIGS Metal foils

Radiation Testing Measure and model effects of radiation 

on thin-film cells a-Si and CIGS Metal foils 

NASA/CP—2005-213431 23
RESULTS OF FLIGHT DEMONSTRATION OF TERRESTRIAL SOLAR CELLS BY MDS-1 IN GTO 

Mitsuru Imaizumi, Taishi Sumita, Shirou Kawakita, Osamu Anzawa, and Sumio Matsuda 

National Space Development Agency of Japan (NASDA)*, Tsukuba, Ibaraki, Japan 

*Current name: Japan Aerospace Exploration Agency (JAXA) 

INTRODUCTION 

It is very costly to develop special electronic devices for space use, and the market for such devices is not 

sufficiently large for business. As a result, semiconductor industries are less aggressive in developing space 

devices. Therefore, application of commercial electronic devices that should be robust against the space 

environment has been considered recently for space applications. 

Mission Demonstration-test Satellite No. 1 (MDS-1) was developed to test and demonstrate state-of-the-art 

technologies in space. One of the MDS-1 missions is to evaluate advanced terrestrial semiconductor devices in 

space environment in order to clarify their possibilities 

or capabilities as space applications.

MDS-1 was launched on February 4, 2002, by 

NASDA’s new launch vehicle H2A. The satellite was 

placed in a geostationary transfer orbit (GTO) with 

apogee of 500km, perigee of 36,000km, inclination 

angle of 28.5o, and period of 10 h 45 min. The orbit 

cuts across the inner and the outer Van Allen Belts so 

that the satellite is exposed to a very severe radiation 

environment compared to satellites in low Earth orbit 

(LEO) or geostationary orbit (GEO). Figure 1 

schematically illustrates the MDS-1 orbit. 

The mission duration was supposed to be one 

year, and the observation of cell degradation in space 

has been completed successfully in February 2003. 

This study presents the degradation flight data of the 

terrestrial solar cells and describes results of 

comparative analysis of the flight data and ground 

irradiation test results using the JPL [1, 2] and NRL [3] 

methodologies. 

TERRESTRIAL SOLAR CELL TEST BOARD 

 MDS-1 is equipped with a test board of solar cells developed for terrestrial use and called the “terrestrial 

solar cell test board” (TSC). Six kinds of terrestrial cells with different structures, materials, or manufacturing 

techniques (see Table I) were attached on the TSC. Two of each type of cell were prepared, and 500μm and 

100μm thick coverglasses were attached on the cells. In addition, two kinds of space solar cells, Si and GaAs, 

were mounted on the TSC for comparison. The combination of the GaAs single junction cell and a 300 μm–thick 

coverglass is the same to the cells on the solar paddles of the satellite. Figure 2 shows the TSC board under light 

irradiation from an AM0 solar simulator. 

The TSC was fitted on top of the satellite (see Fig. 3) and kept facing the sun at all times in orbit with an 

average angle of 10 degrees (MET < 15 days after launch) or 30 degrees (MET  15 days after launch). Five 

Fig. 1 Schematic orbit of MDS-1. The orbit cuts across the 

inner and the outer Van Allen Belts. 

Inner Van Allen Belt

Outer Van Allen Belt

MDS-1

NASA/CP—2005-213431 24
sensors monitored the TSC temperature; the average 

measured temperature, except during eclipse, is 

about 70 oC.

In order to obtain the cell performance, 16 points 

on the current-voltage (I-V) curve, including opencircuit voltage (Voc) and short-circuit current (Isc), of 

each cell were collected from the TSC using 16 loads 

with different resistances. The measured I-V data 

were packed with the temperature values and sent to 

a ground site via telemetry signals. The voltage and 

current values were corrected for temperature, angle 

to the sun, and solar irradiance (distance between the 

sun and the satellite) automatically at the ground site. 

Maximum power (Pmax) values were not 

obtained directly, but can be estimated from the I-V 

data. First, the point of greatest power and adjacent 

two points on an I-V curve are selected from among 

the 16 power values (i.e., products of current and 

voltage values). A fitting equation is then obtained. 

Finally, the maximum power is determined from the 

fitting curve. However, it is impossible to calculate 

conversion efficiency because the solar irradiance is 

not monitored on orbit. In this report, therefore, we will 

focus on the degradation of Isc and Voc. 

The initial I-V curves measured on orbit (MET=4 

days after launch) showed good agreement with the IV curves measured in our laboratory using an AM0 

solar simulator after the correction. This means that 

the cells, as well as the measurement system, were 

not damaged during the launch. 

The radiation environment on orbit is also 

measured by dose monitors (standard dose monitor: 

SDOM) as another mission. The SDOM produces the 

energy flux of the radiation electrons and protons 

detected by the loaded multi channel monitors (12 

channels for protons, 5 channels for electrons). 

Fig. 2 TSC test board under light illumination from an AM0 

solar simulator. 

Table 1 List of Solar Cells on the TSC board 

Cell Type Cover-glass

Thickness (μm) Note

U-type bulk poly-crystalline silicon 100/500 Cast

Bulk poly-crystalline silicon 100/500 Cast

N-type base single crystal silicon 100/500 Thin Emitter

InGaP/GaAs dual-junction tandem 100/500 MOCVD

Large-area Cu(In,Ga)Se2 poly-crystalline thin film 100/500 Wet-chemical deposition

High-efficiency Cu(In,Ga)Se2 poly-crystalline thin film 100/500 Dry-physical deposition

Bulk single crystal silicon (Base resistivity = 10 ȍcm) 100/500 Space cell (reference)

Bulk single crystal silicon (Base resistivity = 2 ȍcm) 100/500 Space cell (reference)

GaAs single crystal thin film on Ge sub. 300/300 Space cell on satellite paddle

Fig. 3 Image of MDS-1 and TSC test board on orbit. 

TSC

NASA/CP—2005-213431 25
FLIGHT DATA 

Prior to the launch, ground irradiation tests for the mounted cells were carried out at the Japan Atomic 

Energy Research Institute (JAERI). In the ground tests the cells were irradiated with unidirectional radiation 

beams without coverglass. The results demonstrated that the CIGS cells have excellent radiation tolerance, while 

the n-type base single-crystal silicon cells have the lowest one. Other test results have been described elsewhere 

[4].

Figure 4 depicts the Isc and Voc of eight types of solar cells on the TSC up to MET of about 400 days after 

launch, expressed as remaining factor. The data on MET=4 days after launch were used as the initial values. 

Each data point is the average of a date. 

The reference space silicon cell data (Figs. 4(g) and 4(h)) indicate that the remaining factor of Voc is higher 

for the lower resistivity (2 ȍcm) cells, while that of Isc is higher for the higher resistivity (10 ȍcm) cells. This 

tendency was expected from the ground test results as well as from experience with Si space cells, and the flight 

data confirmed it. Therefore, the flight data are not considered to express irregular degradation. Note that the 

degradation rate changes around MET=50 days after launch. Figure 6 presents the day fluence rate for a typical 

channel of dose monitors for protons (ch. 2: 1.08-1.46 MeV) and electrons (ch.4: 6.36-10.0 MeV) in SDOM system. 

According to the environmental data, the fluence rate of both electrons and protons increased by more than one 

order of magnitude at this time. The radiation fluence increase around MET=50 days is attributed to a solar flare 

and is thought to have increased the degradation rate of the TSC cells. 

The poly-crystalline silicon cells (Figs. 4(a) and 4(b)) exhibit radiation tolerance comparable to space singlecrystal silicon cells, supposedly due to their relatively low initial performance. However, this type of cell could be 

acceptable for low-power, low-cost space missions.

The n-type base single-crystal silicon terrestrial cells exhibit the greatest degradation (Fig.4(c)), which also 

coincides with the ground test results. This is believed to be not due to its structure but the material. The radiation 

damage in an n-type single crystal silicon cell is known to greatly decrease the minority-carrier lifetime due to 

generated recombination centers. 

The InGaP/GaAs dual-junction tandem cells exhibit good radiation tolerance (Fig. 4(d)), in spite of being 

designed without consideration for radiation damage. This radiation tolerance is thought to be mainly due to its 

dual-junction (DJ) tandem structure. The superior radiation tolerance of the InGaP top cell [5] is thought to 

contribute the tolerance although the bottom GaAs cell is current limiting in the DJ cell. The results imply that the 

cell can be applied without any modification to space missions for which the radiation environment is not severe. 

Almost no degradation on orbit was observed in the CIGS cell (Figs. 4(e) and 4(f)), especially with regard to 

Isc. This agrees with the ground test results, verifying the super radiation tolerance in space. Another study by the 

authors previously reported that radiation-induced damage in CIGS cells can be annealed out quickly at relatively 

low temperature [6]. According to the study described in ref. [6], an average temperature of 70oC of the TSC is 

high enough to recover most of radiation damage. 

Table 2 Equivalent 1 MeV electron fluence for TSC cells at MET=400 days after launch 

Cell Type Remaining Factor of Isc/Voc

(100μm Coverglass Cell)

Equivalent 1MeV e-

 Fluence

(cm-2)

U-type bulk poly-crystalline silicon 0.79 / 0.83 6E14 / 5E15

Bulk poly-crystalline silicon 0.80 / 0.86 6E14 / 4E15

N-type base single crystal silicon 0.67 / 0.70 1E14 / 4E15

InGaP/GaAs dual-junction tandem 0.87 / 0.91 2E14 / 2E15

Large-area Cu(In,Ga)Se2 poly-crystalline thin film 1.00 / 0.97 --- / 1E17

High-efficiency Cu(In,Ga)Se2 poly-crystalline thin film 0.96 / 0.92 --- / 1E17

Bulk single crystal silicon (Base resistivity = 10 ȍcm) 0.89 / 0.74 1E15 / 1E16

Bulk single crystal silicon (Base resistivity = 2 ȍcm) 0.83 / 0.80 6E14 / 7E15

NASA/CP—2005-213431 26
(a) U-type polycrystalline Si (b) Polycrystalline Si 

(c) Single crystal n-type base Si (d) InGaP/GaAs dual junction 

(e) Large area CIGS (f) High efficiency CIGS 

(g) Space Si (10 ȍcm) (h) Space Si (2 ȍcm) 

Fig. 4 Flight degradation data of Voc and Isc of solar cells on TSC. 

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

0.6

0.7

0.8

0.9

1.0

1 10 100 1000

MET (days after launch)

Remaining Factor

CG=100μm Isc

CG=500μm Isc

CG=100μm Voc

CG=500μm Voc

NASA/CP—2005-213431 27
DATA ANALYSIS 

Ground-based experiments of proton and electron irradiation on the TSC cells were performed prior to the 

launch. In particular, radiation response of the single crystal silicon space cells was investigated in detail and 

derived relative damage coefficients (RDC) for the cells since the cells are reference. As a result, the RDC agreed 

with the values exhibited in Solar Cell Radiation Handbook [1]. 

The equivalent 1 MeV electron fluence based on JPL methodology for degradation of each cell at MET=400 

days after launch is presented in Table 2. These values were obtained by comparing flight data of 100μm-thick 

cover-glass cells with the ground irradiation test data of bare cells [4]. According to the Monte Carlo simulation, 

1MeV electrons pass through the 100μm-thick cover-glass with almost no energy loss. Therefore, a shielding 

effect for incident electrons does not have to be considered in this case and the fluence values obtained are 

considered to be significant. Note that the equivalent fluence of Voc is generally about an order of magnitude 

higher than that of Isc. This might be due to differences in irradiated particles: One MeV electrons were irradiated 

for the ground test, while protons with various energies are major particles in the flight environment. Thus, 

damage in the shallow region of the cells including the 

space charge region due to especially low energy 

protons is likely to a reason of the difference. 

Cell degradation data from TSC and the ground 

tests were converted to the function of displacement 

damage dose (Dd), using the non-ionizing energy loss 

(NIEL) values for each cell material (Si, III-Vs, and 

CIGS) [7]. The NIEL value of GaAs has been used for 

InGaP/GaAs dual-junction cells because GaAs 

damage usually determines multi-junction cell 

degradation [8] and also the GaAs NIEL value is close 

to that of InGaP. 

Figure 5 shows the differential proton fluence 

spectra corresponding to the total incident protons at 

400 days after launch obtained from the SDOM. The 

diamond symbols correspond to the proton fluence 

detected by the 12 channels of the proton dose 

monitor. To account for the coverglass shielding effect, 

the transmitted proton fluence spectra were calculated 

by the model described by Haffner [9]. The calculated 

spectra are indicated in Fig. 5 with the coverglass 

thickness. The amount of Dd can be determined by 

summing the product of the differential spectrum with 

the NIEL values of the cells. 

Figure 6 presents the flight data and the ground test results of Isc and Voc of representative terrestrial cells 

on TSC as a function of Dd. The ground irradiation tests were carried out with 1MeV electrons and 380keV, 3MeV, 

and 10MeV protons [4]. However, in the case of the large-area CIGS cell, the ground test result of 1MeV electron 

irradiation was not plotted since no degradation was observed in ground tests even in the high fluence region 

(a1017cm-2). In addition, Dd value of electrons for the flight data was not included since the degradation was 

supposed to be mainly due to protons according to the observed environment. In this figure, degradation curves 

from flight data for 100μm and 500μm coverglass are identical, confirming that the calculation of transmitted 

proton energy spectrum was successful. 

Generally, degradation tendency of the flight data and that of ground test results for both Isc and Voc of all 

the cells show good agreement even for the large-area CIGS cell which exhibited almost no degradation in space. 

Especially for Isc, degradation curves are almost identical. This confirms the validity of the ground test 

method/conditions adopted in this study. 

From the results of the Dd analysis, it is possible to predict EOL performance on orbit. For example, in the 

case of the CIGS cell, the total Dd required to cause 10% degradation in Isc is estimated to be appropriately 

1.4u1012 MeV/g from Fig. 6(d). This is 650 times of the Dd value on orbit at MET=400 days after launch for a cell 

with a 100μm coverglass and 6700 times for a cell with a 500μm coverglass! These values imply that CIGS cells 

have practically infinite lifetime in space as far as remaining factor is concerned.

Fig. 5 Differential incident proton dose spectra on orbit and 

the transmitted spectra after passing through 100μm and 

500μm-thick coverglasses. 

0.1 1 10 100 1000

Proton Energy (MeV)

Differential Fluence (cm-2MeV-1)

SDOM Data

108

1011

1013

109

107

1014

1012

1010

100μm CG

500μm CG

NASA/CP—2005-213431 28
(a) Polycrystalline Si 

(b) Single crystal n-type base Si 

(c) InGaP/GaAs dual junction 

(d) Large area CIGS 

Fig. 6 Comparison of flight and ground test results for Isc and Voc as a function of Dd 

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Isc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Voc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Isc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Voc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Isc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Voc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Isc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

0

0.2

0.4

0.6

0.8

1

Dd (MeV/g)

Remaining Factor of Voc

TSC 100μm CG

TSC 500μm CG

Ground Test

106 1014 1010 108 1012

NASA/CP—2005-213431 29
However, for silicon cells (Figs. 6(a) and 6(b)), degradation of Voc on orbit exhibits slightly lower tolerance 

compared with the ground test degradation curve. It is likely because Dd for electrons on orbit was not included. If 

the Dd of incident electrons, which is calculated to be about 1u109

 MeV/g for MET=400 days after launch, is taken 

into account, the Voc degradation roughly corresponds with the ground test results. This fact suggests that 

degradation in Voc of silicon cells might involve an additional effect which has to be considered for NIEL analysis. 

Otherwise, some modification should be required in the NIEL calculation.

SUMMARY 

MDS-1 successfully monitored degradation tendencies of six types of terrestrial solar cells by flying in a 

severe radiation environment for more than a year. An increase in degradation rate of Voc and Isc was observed 

at MET~50 days after launch, which is due to an increase in the fluence rate of high energy electrons and protons 

caused by solar flare. Super-radiation tolerance of CIGS cells observed in ground irradiation tests was verified in 

actual space. Also, InGaP/GaAs tandem cells exhibited sufficient radiation tolerance, and poly-crystalline silicon 

cells, reasonable radiation tolerance. The results indicate the possibility and capability of the terrestrial cells for 

space application. 

The equivalent 1MeV electron fluence of Voc estimated at MET=400 days after launch was generally higher 

than that of Isc, indicating that degradation in Voc involves an additional effect that does not affect Isc degradation. 

According to the results of NIEL/Dd analysis, degradation tendency of the flight data and that of ground test 

results agree very well. This confirms the validity of the ground test method/conditions adopted in this study. 

ACKNOWLEDGEMENTS 

The authors would like to express their sincere appreciation to the New Energy and Industrial Technology 

Development Organization (NEDO) for their kind cooperation to this mission. They would also like to thank Dr. T. 

Hisamatsu of SHARP Corporation for fruitful discussion. 

REFERENCES

[1] H. Y. Tada, “Solar Cell Radiation Handbook, Third Edition” (JPL Publication, 1982) 

[2] B. E. Anspaugh, “GaAs solar Cell Radiation Handbook” (JPL Publication, 1996) 

[3] G.P Summers et al. Proceedings of the 1st World conference of Photovoltaic Energy Conversion, (1994) 2068 

[4] T. Aburaya, T. Hisamatsu, and S. Matsuda, Proceedings of 2nd World Conference and Exhibition on 

Photovoltaic Solar Energy Conversion, (1998) 3568 

[5] N. Dharmarasu, M. Yamaguchi, A. Khan, T. Yamada, T. Tanabe, S. Takagishi, T. Takamoto, T. Ohshima, H. 

Ito, M. Imaizumi, and S. Matsuda, Appl. Phys. Letters, 79 (2001) 2399 

[6] S. Kawakita, M. Imaizumi, M. Yamaguchi, K. Kushiya, T. Ohshima, H. Itoh, and S. Matsuda, Jpn. J. Appl. 

Phys., 41 (2002) L797 

[7] S. R. Messenger, G. P. Summers, E. A. Burke, R. J. Walters and M.A.Xapsos, Progress in Photovoltaics, 9

(2001) 103 

[8] T. Sumita, M. Imaizumi, S. Matsuda, T. Ohshima, A. Ohi and H. Itoh, Nucl. Instr. Meth. Phys. Res. B 206

(2003) 448 

[9] J.W.Haffner, “Radiation and Shielding in Space” (Academic, New York, 1967) 

NASA/CP—2005-213431 30
SOLAR ARRAY VERIFICATION AND ANALYSIS TOOL (SAVANT) ANALYSIS OF THE 

MICROELECTRONICS AND PHOTONICS TESTBED (MPTB) SPACE SOLAR CELL DATA 

R. J. Walters and G. P. Summers 

US Naval Research Laboratory, Code 6818 

S. R. Messenger 

SFA, Inc. 

T. L. Morton 

Ohio Aerospace Institute 

1. Introduction 

The United States Naval Research Laboratory (NRL), in collaboration with the Ohio Aerospace Institute (OAI) 

and NASA Glenn Research Center (GRC), have developed an improved space solar cell radiation response 

analysis capability and produced a computer modeling tool which implements the analysis. This was 

accomplished, in part, through analysis of data taken from the solar panel that powers the Micro-electronics and 

Photonics Test Bed (MPTB) experiment. This project was funded by the NASA Living with a Star Space 

Environments Testbed, NRA8-3. This project specifically addresses issues related to rapid technological change 

in the area of solar cells for space applications in order to enhance system performance, decrease risk and reduce 

cost for future missions. In this paper, an analysis of MPTB solar panel flight data is presented. The analysis 

focuses on the data for the photo-generated current from the MPTB solar array that was telemetered to the ground 

from the spacecraft. The concept of displacement damage dose (Dd) has been used to characterize the radiationinduced degradation of the solar array output, employing the methodology developed at NRL. The computer 

code Solar Array Verification and Analysis Tool (SAVANT) being developed by OAI under contract to GRC, 

was used to predict the solar array output as a function of time in orbit. The ultimate goal of this work was to 

validate the SAVANT code against the measured space data and to develop SAVANT into a user-friendly 

executable program that can be widely distributed. 

2. MPTB Solar Array Data 

The solar array under study in this project is mounted on the 

host spacecraft and is dedicated to providing power to the 

MPTB experiment. The array operates at a fixed voltage of 

31.2 V, but the current is not controlled. Therefore, any 

changes in the photocurrent of the array can be measured and 

analyzed. The array consists of 27 strings of cells with 54 

cells in each string. The solar cell technology is Spectrolab 

(SPL) 8 mil single-junction (SJ) GaAs/Ge. The solar array 

consists of a 3⁄4 inch think honeycomb Al substrate with 10 mil 

graphite epoxy face-sheets. A schematic diagram of the solar 

array structure is shown in Figure 1. 

OCI Coating 3 Mil

Coverglass 20 Mil

Adhesive 2 Mil

GaAs Solar Cell 8 Mil

Interconnect/Cell Adhesive 4 Mil

FM 73 Dielectric 10 Mil

Graphic Epoxy Face Sheet 10 Mil

FM73 Epoxy/Scrim Cloth 2 Mil

AL Honeycomb 750 Mil

FM 73 Epoxy 2 Mil

Graphic Epoxy Face Sheet 10 Mil

White Paint 5 Mil

OCI Coating 3 Mil

Coverglass 20 Mil

Adhesive 2 Mil

GaAs Solar Cell 8 Mil

Interconnect/Cell Adhesive 4 Mil

FM 73 Dielectric 10 Mil

Graphic Epoxy Face Sheet 10 Mil

FM73 Epoxy/Scrim Cloth 2 Mil

AL Honeycomb 750 Mil

FM 73 Epoxy 2 Mil

Graphic Epoxy Face Sheet 10 Mil

White Paint 5 Mil

Figure 1: A schematic diagram of the MPTB 

solar array. 

NASA/CP—2005-213431 31
The spacecraft telemetry data includes both the solar 

array temperature and current (Figure 2). To minimize 

earth albedo effects, only the data taken near apogee 

were analyzed. The backside solar array temperature 

varied from about 32 to 49oC. There is about an 11.1oC 

temperature increase from the rear to the front face of 

the array. Using temperature coefficient data from 

reference [1], the on-orbit current data was corrected to 

28o

C to match ground measurements. The solar array 

current data were also corrected for seasonal variations 

where it was assumed that the spacecraft was collocated 

with the Earth. The Sun-Earth separation was calculated 

as a function of time using Kepler’s equations [2]. Since 

the photocurrent of the solar array varies linearly with 

solar intensity and since the solar intensity varies with 

the inverse square of distance, the seasonal variations 

could be normalized to the data measured at the average 

Sun-Earth distance, 1 A.U.. 

As can be seen in Figure 2, the seasonal correction 

produced the desired effect of smoothing out the 

periodic oscillations in the measured data. The roughly 

saw-tooth pattern visible in the data occurs because the 

original data has a distinct stair-step pattern and the plateaus at each perihelion and aphelion are due to the limited 

resolution of the analog to digital (A/D) converter onboard the spacecraft. 

3. Ground Test Data 

The prediction of the performance of a solar cell in orbit 

is based on measurements of the cell output made on the 

ground. In this section, the ground test data reduction 

and analysis will be described. SPL supplied NRL with a 

set of ground test data consisting of the current vs. 

voltage (IV) parameters depicted in Figure 3 measured 

after irradiation with incrementally larger fluences of 1 

MeV electrons. A second dataset measured after 1 MeV 

electron irradiation was taken from [3]. In a separate 

project funded by SPL, NRL measured 0.4 and 1 MeV 

proton irradiation data on SPL SJ GaAs/Ge solar cells. 

The Isc data from these various sources are plotted in 

Figure 3. 

Under typical operational conditions, a solar array 

operates at a fixed voltage, so the individual solar cells 

operate at a fixed load point characterized by a load 

current and voltage. The data telemetered from the 

MPTB spacecraft is the solar array load current. 

Therefore, we must determine if the radiation response of the solar array load current is accurately modeled by the 

ground test parameters. The array consists of 27 strings of 54 solar cells in each string. Single junction GaAs 

solar cells have an open circuit voltage of about 1.02 V. The array open circuit voltage is therefore expected to be 

54 x 1.02 = 55.9 V. The MPTB solar array voltage is held constant at 31.2 V. Therefore, the array load point is 

24

25

26

27

28

29

30

31

9-Nov-97 28-May-98 14-Dec-98 2-Jul-99 18-Jan-00 5-Aug-00 21-Feb-01 9-Sep-01

Date

Current (A)

Array Current (A)

Array Current - Temp Corrected (A)

Array Current - Temp and Seasonal Variation Corrected (A)

Figure 2: The MPTB on-orbit solar array data. 

The raw telemetered data are shown along with the 

data after temperature correction and correction for 

seasonal variations. 

Figure 3: Short circuit data measured in SPL SJ 

GaAs/Ge solar cells after proton and electron 

irradiation. The data have been normalized to 

the pre-irradiation values.

0.5

0.6

0.7

0.8

0.9

1.0

1.1

1E+10 1E+11 1E+12 1E+13 1E+14 1E+15

Particle Fluence (cm-2)

Norm Isc

1 MeV e- data from SPL

1 MeV e- data from [3]

0.4 MeV p+ data from NRL

1 MeV p+ data from NRL

NASA/CP—2005-213431 32
approximately 56.8% of Voc. Using this value, we can 

estimate the load voltage for an individual cell to be 

56.8% of Voc or 0.578 V. With this information, we 

may now track the radiation degradation of the solar 

cell current at this load point (I-load). 

Using 1 MeV proton irradiation ground test data, we 

generated the graph in Figure 4. The maximum power 

point and the load point on each curve are highlighted 

in the graph. The two different points are seen to 

follow different tracks through the current-voltage 

space with increasing fluence, but since the slope of the 

IV curve is very small up to the maximum power point, 

the change in I-load with increasing fluence is 

essentially the same as that of the Imp. This is shown 

explicitly in where the I-load, Imp, and Isc have been 

normalized to their pre-irradiation values and plotted as 

a function of fluence. These data suggest that I-load 

tracks Imp closely, so we will use the data for Imp in 

the calculations that follow. 

4. Ground Test Data Reduction and Analysis 

The analysis of the ground test data performed here 

employs the displacement damage dose (Dd) 

methodology developed by NRL [4,5]. In the Dd 

methodology, the data measured after irradiation by 

different particles at different energies are reduced to a 

common degradation curve by converting the particle 

fluence to an equivalent value of Dd. The Dd is 

calculated through the following expression: 

The quantity Φ(E) is the particle fluence. The 

quantity, S(E), is the calculated nonionizing energy 

loss (NIEL) at incident particle energy E. Values of 

NIEL for electrons and protons incident on GaAs have 

been calculated and are tabulated in [5]. The symbol 

Eref refers to an arbitrary reference energy, and n is an 

empirically determined parameter. For proton 

irradiation of GaAs and a wide array of other 

semiconductor materials, n has been found to be unity 

indicating a linear dependence of the proton damage coefficients on NIEL [5]. In the case of electron irradiation, 

however, n has been found to vary between 1 and 2 depending on the material and dopant species. For p/n GaAs 

solar cells like those used on MPTB, n has been determined to be 1.7 [5]. To keep consistency with previous 

studies, the reference energy is set to 1 MeV for the electron irradiation data. The radiation degradation data for 

Imp corresponding to the data sets given in Figure 3 are plotted as a function of Dd in Figure 6. 

Figure 6 shows that the data measured after irradiation by different particles at different energies collapse to a 

single curve when plotted as a function of Dd. This curve is therefore characteristic for this technology and be 

1 MeV proton fluence (cm-2)

-140

-120

-100

-80

-60

-40

-20

0

0 0.2 0.4 0.6 0.8 1 1.2

Voltage (V)

Current (mA)

0

3.00E+10

1.00E+11

3.00E+11

1.00E+12

3.00E+12

Figure 4: IV curves measured on SJ GaAs solar 

cells that are representative of the cells on the 

MPTB array. Each curve represents the cell output 

after irradiation by 1 MeV protons up to the fluence 

level given in the legend. The squares indicate the 

maximum power point, and the diamonds indicate 

the load point. 

SPL SJ GaAs

1 MeV proton irrad.

0.5

0.6

0.7

0.8

0.9

1

1.1

1.E+10 1.E+11 1.E+12 1.E+13

1 MeV Proton Fluence (cm-2)

Normalized Current

I Load

Imp

Isc

Figure 5: Current data taken from Figure 4. These 

data show that the three parameters show similar 

degradation characteristics, but Imp is a better 

description of I-load than Isc. 

1

( )

( ) ( ) ( ) ( )

−















 = Φ •

n

Eref S

S E Dd E E S E Equation 1 

NASA/CP—2005-213431 33
used to describe or predict the cell response in any proton 

or electron radiation environment. This is the foundation 

of the NRL method. By analyzing the data in terms of 

Dd, data from a variety of different radiations can be 

readily reduced to a common curve. Moreover, since the 

NIEL is a calculated quantity, the characteristic curve can 

be determined from a single proton data set. In the case 

of electrons, at least two data sets are required to enable a 

determination of the n parameter. From a practical 

standpoint, this represents a significant advance because 

it greatly reduces the amount of radiation data required to 

fully characterize the radiation response of a solar cell 

technology. The Dd method frees resources that can be 

used to generate more data points at each specific energy, 

thereby specifying the characteristic curve more precisely 

and producing a more accurate predictive capability at a 

lower cost. Also, the Dd analysis allows a full 

characterization for new technologies even though there 

may be very little radiation data available. 

The next step in the data analysis is to parameterize the 

characteristic curve. This is typically done by fitting the 

data to the following semi-empirical equation: 

The terms C and Dx are the fitting parameters and Imp/Imp0 refers to the value of Imp measured after irradiation 

and normalized to the pre-irradiation value. This is essentially the equation presented in [3] except that the 

particle fluence has been replaced by Dd. Similar expressions can be written for the other PV parameters. 

Because the nonlinear dependence of the electron Dd on NIEL, the characteristic curve for the electron data does 

not typically align with that from the proton data without further manipulation. Therefore, in the analysis, the 

electron and proton data are typically fit separately. The separation of the proton and electron is not very evident 

in Figure 6 because the separation for the Imp data is typically small, while larger differences are often observed 

for the other PV parameters. The electron and proton data are fit such that each is described by a common value 

of C and a separate value of Dx. The proton Dx is designated by Dxp and the electron by Dxe. The ratio of Dxe 

to Dxp is referred to as the Rep factor, which is used to convert an electron Dd value to an equivalent proton Dd 

value. The fit parameters determined for each of the photovoltaic parameters are given in Table 1. 

Table 1: The parameters determined from fitting each of the GaAs PV parameter datasets to Equation 2. The 

columns identified as dC, dDxp, etc. are the uncertainty associated with the calculated parameter. The Dxp and 

Dxe parameters are in units of MeV/g, which is the unit of Dd. 

 C dC Dxp dDxp Dxe dDxe Rep 

Isc 0.153 0.002 2.36E+09 6E+07 3E+09 1E+09 1.3 

Voc 0.112 0.002 1.81E+09 9E+07 2E+09 1E+09 1.1 

Imp 0.236 0.004 5.2E+09 2E+08 6E+09 2E+09 1.2 

Vmp 0.128 0.002 1.9E+09 1E+08 3E+09 2E+09 1.6 

Pmp 0.261 0.003 1.57E+09 8E+07 2.7E+09 7E+08 1.7 

0.5

0.6

0.7

0.8

0.9

1.0

1.1

1E+07 1E+08 1E+09 1E+10 1E+11

Displacement Damage Dose (Dd)

Norm Imp

1 MeV electron data from SPL

1 MeV e- Isc data from [3]

1 MeV proton data from NRL

0.4 MeV proton data from NRL

Figure 6: GaAs/Ge solar cell Imp data plotted as 

a function of Dd. The data are seen to follow the 

same curve, which is characteristic for this 

technology and can be used to describe the cell 

response in any radiation environment. Imp data 

were not available in [3], so Isc data are shown. 





 



 = − +

Dx

Dd Imp/Imp 1 Clog 1 0 Equation 2 

NASA/CP—2005-213431 34
5. SAVANT CALCULATIONS 

In this section, we will describe the process by which 

the SAVANT code produces a prediction of the solar 

array output as a function of time in orbit. An image 

of the program interface is shown in Figure 7. The 

SAVANT program is being developed into a userfriendly, distributed program, and it is presently being 

distributed for beta-version testing [6]. The procedure 

begins with a determination and specification of the 

spacecraft orbit. However, the actual orbit for the 

MPTB spacecraft is not available for general use. 

Based on the published report of Dyer, et al. [7], an 

orbit consisting of 39,158 x 1,212 km 63.43o

inclination for 225 days, and then at 38,868 x 1,502 km 

63.43o inclination for the remainder of the mission was 

assumed. A detailed description of the actual 

ephemeris data was also obtained. The SAVANT code was run for both of these orbits. As will be seen in the 

results to be presented, there is only a small difference between these two calculations in terms of the calculated 

solar cell degradation. This result gives us confidence in our original choice of orbit. 

Once the orbit has been defined, the SAVANT code implements a model to account for geomagnetic shielding 

and then calls on the NASA AP8 and AE8 computer codes to calculate the proton and electron radiation 

environments, respectively. This is important because the degradation predictions will only be as accurate as the 

environment calculations. This research has given insight into the accuracy of the environment calculations [8]. 

Both the AP8 and AE8 programs "advertise" an inherent factor of two uncertainty due to scatter in the original 

space data. Also, AP8 and AE8 are static models that have not been updated since their release (1976 for AP-8, 

and 1991 for AE-8). They are based on satellite data taken between 1963 and 1970, and 1970 was the lowest 

solar max period of the space age. Therefore, the accuracy of the environment models is taken as a factor of two 

at best. 

With the radiation environment calculated, SAVANT performs an analysis of the shielding effects of materials in 

contact with the solar cells. In this procedure, the radiation spectrum incident directly upon the cell active region 

must be determined since the ground based solar cell measurements are performed on unshielded solar cells. This 

is done in SAVANT by the application of the continuous slowing-down approximation to the incident spectrum 

using slab geometry. The result is the slowed-down, omnidirectional, differential spectrum that emerges from the 

shielding material and is incident directly upon the solar cell active region [4, 9]. 

Typically, solar cells are shielded on the front by a coverglass and on the back by the solar array substrate. A 

schematic diagram of the MPTB solar array structure is shown in Figure 1. The shielding properties of each of 

the layers and the coverglass are included in SAVANT by reducing the multi-layer stack to an equivalent 

thickness of solar cell coverglass using the densities of the materials. This is summarized in Table 2 for the solar 

array itself, and it is seen that the rear-side shielding is equivalent to 31.25 mils of CMG coverglass. Examples of 

the calculated incident proton spectrum and associated slowed-down spectra are shown in Figure 8. 

As is evident from Figure 8 for the case of protons, the resultant radiation environment impinging on the solar 

cells is complex, consisting of protons and electrons with energies extending continuously over a large range. 

Because the ground test data are taken with unidirectional, monoenergetic radiation, a correlation method is 

required. With the NRL Dd method, the correlation is achieved by reducing the particle spectrum to an 

equivalent value of Dd. This is accomplished by integrating the product of slowed-down differential spectrum 

with the NIEL as a function of energy. The functional form of the integral comes from Equation 1 where the 

Figure 7: The input screen for the SAVANT GUI. 

NASA/CP—2005-213431 35
monoenergetic particle fluence is replaced by the differential fluence spectrum. The electron and proton 

contributions are calculated separately and then summed using the Rep factor as shown in Equation 3: 

In Equation 3, the “p” and “e” subscripts refer to the proton and electron cases, respectively. The value of n has 

been set to 1.7. The value of the Rep factor is given in Table 1. 

The SAVANT code performs this integral iteratively over time at a user defined time interval. For the MPTB 

mission, the accumulated Dd was calculated once per day. The results are shown in Figure 9. Results are shown 

for the two descriptions of the MPTB orbit given above. Up to about April 2000, the two orbits result in 

essentially the same accumulated Dd. After that date, the orbit determined from the orbital ephemeris data results 

in a lower accumulated Dd, but the difference is small. 

With the equivalent Dd values calculated, the performance of the MPTB solar cells as a function of time in orbit 

can be predicted. This is done by inserting the equivalent Dd values from Figure 9 into the solar cell 

characteristic degradation curve (Equation 1). The resultant data are shown in Fig. 11. In these calculations, the 

ephemeris data were used for the orbit, and the Imp parameters were used in the solar cell degradation 

calculations. 

The predictions are seen to underestimate the measured solar cell degradation. A second calculation is shown by 

the dashed curve assuming twice the estimated orbital value of Dd. Up to about August of 2000, the data 

0.7

e

e

e

e

p

p

Eqiv S (1MeV)

S (E) S (E) dEd (E) S (E)dE Re p dE

d (E) Dd 



 



 • Φ • +

Φ = ∫ ∫ Equation 3 

Figure 8: These are SAVANT 

calculations of the proton environment for 

the MPTB mission. Shown are the 

incident proton spectrum and the proton 

spectrum emerging from the front and rear 

shielding. The total spectrum is that 

incident upon the solar cells. 

1E+06

1E+07

1E+08

1E+09

1E+10

1E+11

1E+12

1E+13

1E+14

1E+15

1E+16

1.E-03 1.E-02 1.E-01 1.E+00 1.E+01 1.E+02 1.E+03

Energy (MeV)

Differential Proton Fluence (MeV -1cm -2)

Incident Spectrum

Rear Fluence

Front Spectrum

Total Spectrum

Table 2: Table of thickness and density of the layers 

comprising the MPTB solar array. The values are reduced to 

an equivalent thickness of coverglass material at the bottom of 

the table. These data are used to determine the radiation 

shielding of the MPTB solar cells on orbit.

Thickness 

(mils) 

Density 

(g/cm3) 

Equiv. thickness 

(mil g/cm3) 

Adhesive/Interconnect 4 0.64 2.56 

FM 73 dielectric 10 1.16 11.6 

Graphite Epoxy Face 

Sheet 10 1.65 16.5 

FM 73 dielectric 2 1.16 2.32 

Al honeycomb 750 0.032 24 

FM 73 dielectric 2 1.16 2.32 

Graphite Epoxy Face 

Sheet 10 1.65 16.5 

White Paint 5 0.8 4 

Total Equivalent 

thickness 79.8 

Density of Coverglass 

(g/cm3

) 2.554 

Total/CMG = equiv 

coverglass (mils) 31.25 

NASA/CP—2005-213431 36
calculated assuming twice the Dd seem to agree with 

the measured data very well. After that date, however, 

there appears to be a reduction in the degradation rate in 

the measured data. We have performed an investigation 

in an attempt to resolve this apparent discrepancy. This 

investigation made use of the onboard dosimeter data as 

well as an analysis of the effect of solar proton events, 

as will be described in the next sections. 

6. Onboard Dosimeter Analysis 

The MPTB experiment included the CREDO 

experiment [7] that measures the total fluence of 

protons with energy above 38 MeV per orbit. Some of 

the CREDO data are plotted in . As a comparison, 

SAVANT was used to predict the total fluence of 

protons with energies greater than 40 MeV received per 

orbit, and these data are also shown in Figure 12. Note 

that the data are shown as lines connecting the 

individual data points, and the lines appear to have a 

thickness. This is due to the fact that the even and odd 

orbits of MPTB alternatively pass through the South 

Atlantic Anomaly (SAA), so each orbit delivers 

different amounts of radiation to the spacecraft. 

Initially, SAVANT underpredicts the measured data by 

about a factor of 2. This is in agreement with the 

apparent factor of 2 under prediction of the solar array 

degradation seen in Figure 10. 

It is interesting to note that the CREDO data shows a 

decrease in the total proton fluence absorbed after about 

3/15/00 indicating a change in the radiation 

environment. This roughly coincides with the decrease 

in the degradation rate of the MPTB solar array output 

highlighted in Figure 10. The cause for this change is 

not currently known. 

7. Solar Event Analysis 

The CREDO data have been used to quantify the proton 

fluence incident on the MPTB spacecraft due to solar 

events. The CREDO data are shown in two forms in 

Figure 12. One curve in Figure 12 represents the 

total number of protons measured the CREDO instrument. The other curve is a subset of the CREDO data that 

represents the proton fluence measured while the spacecraft was outside of the trapped proton belts. The spikes in 

the data measured outside the proton belts correspond to solar proton events. Viewing the data in this fashion 

shows that while the solar events do produce large fluxes of protons for a short time, the vast majority of the 

protons incident on the spacecraft come from the trapped proton belts. The additional protons produced by solar 

proton events are relatively small in number. 

The data in Figure 12 suggest that the effect of the solar event protons will be minimal on the MPTB solar 

arrays. To quantify this result, a SAVANT calculation was made with a solar proton spectrum based on the July 

0.0.E+00

2.0.E+09

4.0.E+09

6.0.E+09

8.0.E+09

1.0.E+10

1.2.E+10

1.4.E+10

5/5/1997 11/21/1997 6/9/1998 12/26/1998 7/14/1999 1/30/2000 8/17/2000 3/5/2001 9/21/2001 4/9/2002 10/26/2002

Date

Accumulated Dd (MeV/g)

assuming orbit with single perigee change

using ephmeris data for orbital calculation

Figure 10: Equivalent Dd values accumulated 

by the MPTB solar cells calculated for each day 

in orbit. Calculations for the two descriptions of 

the MPTB orbits are shown. 

0.8

0.82

0.84

0.86

0.88

0.9

0.92

0.94

0.96

0.98

1

11/22/97 6/10/98 12/27/98 7/15/99 1/31/00 8/18/00 3/6/01 9/22/01 4/10/02 10/27/02

Date

Normlaized Imp

SAVANT calculation

Telemetered data

SAVANT calculation assuimng twice 

the calculated Data

Figure 10: These are the predicted solar array data 

from SAVANT compared to the measured data. 

The ephemeris data were used to generate the Dd 

values, and the degradation characteristics of Imp 

we used in the solar cell calculations. For 

discussion purposes, predictions assuming twice 

the amount of Dd appear to match the measured 

data more closely. 

NASA/CP—2005-213431 37
14, 2000, and the equivalent proton Dd was 

determined to be 1.9x106

 MeV/g. This is 

equivalent to about ten days in the MPTB orbit. 

Given that the mission spans thousands of days, 

this is not significant and will not noticeably 

change the slope of the solar cell damage curves. 

8. Discussion 

The data in Figure 10 represent the best effort at 

predicting the on-orbit solar array data. The 

predictions are accurate to within a factor of about 

two in terms of accumulated Dd. While this result 

is within the stated uncertainties of the 

environment models [8], previous analyses have 

produced better agreement between the SAVANT 

estimates and measured on-orbit data [9,10]. One 

difference with MPTB is that it is, by far, the 

longest mission to be considered. There are over 

four years of MPTB data as compared to less than 

2 years for the next longest mission. This means 

the predictions are especially sensitive to temporal 

variations in the radiation environment since such 

variations are not accounted for in the AP8 model. 

In addition, the solar cells on MPTB have a 20 mil 

thick coverglass, which is more than three times 

the thickness of the thickest coverglass analyzed in 

the past. This means that the portion of the proton 

spectrum under consideration is at much higher 

energies than in past studies. 

Another point to note is that in the MPTB orbit, 

the spacecraft experiences higher electron fluences 

than analyzed in the past. All of previous 

missions analyzed using SAVANT were almost 

exclusively proton environments, so the model has 

yet been fully validated in such a mixed 

environment. Research is currently underway to 

study solar cell response in a mixed electron and 

proton environment. 

While the points highlighted above suggest areas 

where the accuracy of the SAVANT calculations 

may be improved, the primary area of uncertainty 

in the present calculations lies in our knowledge of 

the radiation environment. Considering as an 

example only the higher energy proton of the incident proton spectrum, the CREDO data suggests that AP8 

underestimates the proton flux by a factor of 2. Increasing the SAVANT predicted Dd values by this amount 

brings the predicted solar array output into very close agreement with the measured data at least up to 3/00. 

Beyond this date, there appears to have been a change in the radiation environment of the spacecraft. This factor 

of 2 discrepancy between the measured data and AP8 calculations is in agreement with the CRRES dosimeter 

1.E+06

1.E+07

1.E+08

6/19/97 1/5/98 7/24/98 2/9/99 8/28/99 3/15/00 10/1/00 4/19/01 11/5/01 5/24/02 12/10/02 6/28/03

Date

Fluence per Orbit (4pi Str) (#/cm2)

CREDO data (protons with E>38 MeV)

SAVANT calculation (protons with E>40 MeV)

Figure 12: Values of total fluence of protons with 

energies above 38 MeV absorbed each orbit by the 

CREDO experiment on MPTB along with similar data 

predicted by the SAVANT code. 

1.E+04

1.E+05

1.E+06

1.E+07

1.E+08

1.E+09

1.E+10

6/19/97 1/5/98 7/24/98 2/9/99 8/28/99 3/15/00 10/1/00 4/19/01 11/5/01 5/24/02 12/10/02

orbit number

Fluence per Orbit (4pi Str) (#/cm2)

TOTAL CREDO Prot. Energy >38MeV

Belt Protons Subtracted CREDO Prot. Energy >38MeV

Figure 12: Another plot of the CREDO data. The solid 

curve represents that total CREDO data that shown in 

Figure 12. The dashed curve represents the amount of 

proton fluence measured by the CREDO instrument 

when the spacecraft was outside of the trapped proton 

belts. The spikes in the remaining data correspond to 

solar proton events. 

NASA/CP—2005-213431 38
measurements that indicate that AP8 under predicts the flux for energies > 10 MeV and overpredicts the flux for 

lower energies [11]. Since only protons with incident energies greater than ~8 MeV can pass through the 20 mil 

coverglass on the MPTB solar cells, the MPTB solar cell degradation is expected to track the higher energy 

portion of the spectrum. Therefore, it may be concluded that the calculated results presented here are in good 

agreement with the measured values, where the primary cause of uncertainty lies in the environment models. 

9. References 

[1] M. R. Brown, C. A. Garcia, G. S. Goodelle, J. S. Powe, and J. A. Schwartz, “Characterization Testing of 

Measat GaAs/Ge Solar Cell Assemblies”, Prog. In Photovoltaics: Res. and App., 4, 129 (1996) 

[2] J. B. Marion, Classical Dynamics, Academic Press, NY 1970 

[3] B. E. Anspaugh, “GaAs Solar Cell Radiation Handbook”, JPL Publication 96-9, July 1, 1966. 

[4] G. P. Summers, S. R. Messenger, E. A. Burke, M. A. Xapsos, and R. J. Walters, “Low Energy ProtonInduced Displacement Damage in Shielded GaAs Solar Cells in Space”, Appl. Phys. Lett., 71, 832 (1997) 

[5] 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. Nucl. Sci., vol. 40, pp. 

1372-1379, December 1993. 

[6] For a copy of the Beta-version, contact the authors: phone (202) 767-2533, Email: 

Robert.Walters@nrl.navy.mil 

[7] C. S. Dyer, P. R. Truscott, C. Sanderson, C. Watson, C. L. Peerless, P. Knight, R. Mugford, T. Cousins, and 

R. Noulty, IEEE Trans. Nuc. Sci., 47, 2208, Dec 2000 

[8] T. L. Morton, Proc. 5th International Workshop on Radiation Effects on Semiconductor Devices for Space 

Applications, page 7, Oct 9-11, 2002, 

[9] S. R. Messenger, R. J. Walters, G. P. Summers, T. Morton, G. La Roche, C. Signorini, O. Anzawa, and S. 

Matsuda, “A Displacement Damage Dose Analysis of the COMETS AND EQUATOR-S Space Solar Cell 

Flight Experiments”, Proc. 16th European Photovoltaic Solar Energy Conference and Exhibition, Glasgow, 

Scotland, May 2000 

[10] T. L. Morton, R. Chock, K. Long, S. Bailey, S. R. Messenger, R. J. Walters, and G. P. Summers, Technical 

Digest of the International PVSEC-11, Sapporo Hokkaido, Japan, 1999. 

[11] M. S. Gussenhoven, E. G. Mullen, and D. H. Brautigam, Trans. Nuc. Sci., 46, 353 (1995) 

NASA/CP—2005-213431 39
RECENT PROGRESS ON THE STRETCHED LENS ARRAY (SLA) 

Mark O’Neill and A.J. McDanal, ENTECH, Inc., Keller, TX 76248 

Michael Piszczor and Patrick George, NASA Glenn Research Center, Cleveland, OH 44135 

 Michael Eskenazi and Matthew Botke, ABLE Engineering Company, Goleta, CA 93117 

David Edwards and David Hoppe, NASA Marshall Space Flight Center, Huntsville, AL 35812 

Henry Brandhorst, Auburn University, Auburn, AL 36849 

ABSTRACT 

At the last Space Photovoltaic Research and Technology Conference, SPRAT XVII, held during the fateful week 

of 9/11/01, our team presented a paper on the early developments related to the new Stretched Lens Array (SLA), 

including its evolution from the successful SCARLET array on the NASA/JPL Deep Space 1 spacecraft. Within 

the past two years, the SLA team has made significant progress in the SLA technology, including the successful 

fabrication and testing of a complete four-panel prototype solar array wing (Fig. 1). The prototype wing verified 

the mechanical and structural design of the rigid-panel SLA approach, including multiple successful 

demonstrations of automatic wing deployment. One panel in the prototype wing included four fully functional 

photovoltaic receivers, employing triple-junction solar cells. These receivers were fully encapsulated to enable 

high-voltage operation in space plasma, and the receivers all passed 500 V wet hi-pot testing. Complete 

lens/receiver units were accurately tested for performance using a large-area pulsed solar simulator (LAPSS), 

calibrated with reference cells flown by NASA Glenn on their Lear Jet photovoltaic test facility. The best 

lens/receiver unit achieved 27.5% net efficiency at 28C cell temperature under AM0 sunlight. The measured 

mass and performance of the prototype wing accurately matched predictions. The same performance and mass 

model shows that a 7 kW wing, using the same rigid-panel technology demonstrated on the prototype wing, will 

achieve these unprecedented performance metrics at beginning of life (BOL) on geostationary orbit (GEO, with 

75C cell temperature): 

z >180 W/kg specific power 

z >300 W/m2

 areal power density 

z >300 V operational voltage 

z >9 kW/m3 stowed power at launch 

z >85% savings in cell area (cm2

/W), 

cell mass (g/W), and cell-related cost 

($/W) compared to planar arrays 

In addition to making rapid progress on 

the rigid-panel version of SLA, the SLA 

team has also made significant advances 

over the past two years in the flexibleblanket version of SLA. By integrating 

SLA with ABLE’s new SquareRigger

deployment and support platform, truly 

transformational improvements in 

performance metrics are achievable for 

very large arrays (50 kW to MW class). 

The following paragraphs present further details related to these recent developments on both versions of SLA. 

Figure 1 – Stretched Lens Array (SLA) 

NASA/CP—2005-213431 40
INTRODUCTION 

Fig. 2 shows the basic concept of the Stretched Lens Array (SLA) in an early functional prototype. Thin 

(140 micron), flexible, line-focus Fresnel lenses, made from a space-qualified silicone polymer (Dow Corning 

DC93-500), are deployed and supported by end arches, which tension the lenses in their lengthwise direction 

forming a stressed membrane optical element. These stretched lenses collect space sunlight and focus it onto 

narrow state-of-the-art multi-junction photovoltaic (PV) cell receivers, which are mounted to thin (125 micron) 

carbon composite sheet radiators for waste heat rejection to deep space. 

The arched-shaped lenses are each 8.5 cm 

wide, and focus sunlight onto PV cells which 

are 1.0 cm wide, for a geometric concentration 

ratio of 8.5X. This 8.5X concentration ratio was 

selected to provide + 2 degrees of sun-pointing 

tolerance without appreciable power loss. 

Compared to conventional planar one-sun 

photovoltaic arrays, SLA’s principal and 

inherent advantages include a significantly 

lower mass (kg/m2

) and a substantially lower 

cost ($/W). SLA’s weight advantage is due to 

the simple fact that the lens, radiator, and 

narrow photovoltaic receiver assembly, all 

taken together, weigh about half as much per 

square meter as a one-sun solar cell assembly 

by itself. SLA’s cost advantage is due to the use of 85% less of the expensive photovoltaic cell material per Watt 

of power produced. In addition, SLA offers better electrical performance (W/m2

) than conventional planar one-sun 

photovoltaic arrays, due to the normal gain in cell conversion efficiency with concentration. The small size of the 

photovoltaic receiver in SLA enables super-encapsulation of the photovoltaic circuit with very little mass penalty. 

Such super-encapsulation can be tailored to enable high-voltage operation in space plasma and/or to provide 

radiation hardness for space missions in high charged-particle radiation environments. 

Since a stressed 

membrane support 

method will never provide 

a near-perfect shape for 

the optical element, the 

lens is engineered to be 

extremely shape-error 

tolerant. This shape error 

tolerance is obtained by 

configuring the lens to a 

unique arch shape, 

wherein each prism 

comprising the lens 

symmetrically refracts the 

solar ray passing through 

it. As shown in Fig. 3, for 

each prism in the lens, the 

solar ray angle of 

incidence at the smooth 

outer surface is equal to 

the solar ray angle of 

emergence at the 

prismatic inner surface. 

This symmetrical 

refraction condition, 

Figure 2 – Stretched Lens Array (SLA) 

with Four Lenses and Receivers 

Enlarged

View of Lens

Enlarged View of Focus

on Photovoltaic Cell

Incidence

Angle

Emergence

Angle

Enlarged

View of Lens

Enlarged View of Focus

on Photovoltaic Cell

Incidence

Angle

Emergence

Angle

Figure 3 – Symmetrical Refraction Lens with False-Color Rays 

Showing Wavelengths in the Photovoltaic Cell Response Range 

(0.36 μm to 1.80 μm for All Three Junctions 

of a Triple-Junction GaInP/GaAs/Ge Cell) 

NASA/CP—2005-213431 41
combined with the refractive index of the lens material, fully defines the lens shape and prismatic pattern, and 

implies that each prism is oriented at its minimum-deviation condition. The symmetrical refraction lens has two 

key optical benefits: minimum total reflection loss at the two lens/vacuum interfaces (thereby maximizing 

throughput transmittance) and a unique and remarkable slope error tolerance. Compared to a reflective 

concentrator of any kind or to a conventional flat Fresnel lens, the slope error tolerance of the symmetrical 

refraction lens is more than 100 times better, as further discussed in later sections of this paper. This shape error 

tolerance is critical to the excellent optical performance of the SLA and of its predecessor solar concentrator 

arrays which also used the symmetrical refraction lens approach. The following section compares the 

symmetrical refraction lens approach with reflective concentrator arrays which have recently experienced highly 

publicized problems on communication satellites. 

FRESNEL LENS VERSUS MIRROR SOLAR CONCENTRATORS: VASTLY DIFFERENT 

In the past two years, significant problems with large reflective concentrator arrays on six GEO communication 

satellites have been highly publicized [1-3]. These reflective concentrator arrays use vastly different technology 

than the stretched lens array (SLA), which is the subject of this paper. The reflective concentrator arrays use 60-

degree tilted mirrors on both sides of planar photovoltaic panels to increase the solar flux on these panels, an 

approach which has also been attempted in large terrestrial solar arrays, which also suffered from significant 

problems [4]. The basic concept of using tilted mirrors to augment the solar flux on planar photovoltaic panels is 

shown in the left sketch of Fig. 4. If such mirrors could be made and maintained in a perfect optically flat 

configuration, the reflected solar flux would be uniform over the planar photovoltaic panel, nearly doubling the 

total solar flux on the panel. Unfortunately, shape errors in these mirrors can cause significant losses and nonuniformity of the reflected solar flux on the planar photovoltaic panel, as shown in the right sketch of Fig. 4. If the 

mirror errors lead to concavity, the reflected flux can have spikes and voids. If the mirror errors lead to convexity, 

the reflected flux can be reduced by significant reflected ray losses. Both types of mirror errors are shown in the 

right sketch of Fig. 4, for an example value of the maximum slope error of 10 degrees at the edges of the mirrors. 

In contrast to reflective optics, the stretched lens array (SLA) uses a symmetrical-refraction Fresnel lens, 

described in the previous section, which is by far the most error-tolerant optical concentrator yet developed [5,6]. 

The symmetrical-refraction lens concentrator approach is shown in the left sketch of Fig. 5. The right sketch 

shows a ray trace for the same symmetrical-refraction lens with its shape distorted by a similar amount as for the 

mirrors in Fig. 4, i.e., with 10 degree slope errors at both edges. Note that the lens still focuses almost perfectly 

despite these large slope errors. This unique error tolerance has been fully described in previous publications 

about the symmetrical-refraction lens, which is more than 100 times as tolerant of shape errors as any reflective 

concentrator [6]. A direct comparison of the right sketches of Figs. 4 and 5 shows why shape errors are the most 

critical optical problem for reflective concentrators of all types, while shape errors are not even a significant 

concern for the symmetrical refraction lens. 

Perfect Mirror Shapes –

Uniform Flux and No Lost Rays

Mirrors with Same End Points but 

with 10 Degree Slope Errors at Both 

Edges – Concave on Left (Flux 

Spikes and Voids) and Convex on 

Right (Many Lost Rays)

Incident 

Solar Rays

Reflected 

Solar Rays

Incident 

Solar Rays

Reflected 

Solar Rays

Mirrors

Planar PV 

Panels

Figure 4 – Ray Traces for 60° Tilted Mirrors 

for Perfect Mirrors (Left) and for Mirrors 

with Shape Errors (Right) 

Perfect Lens Shape Same Lens with 10 Degree Slope 

Errors at Both Edges

Refracted 

Solar Rays

Stretched 

Fresnel 

Lenses

Solar Cells

Radiators

Perfect Lens Shape Same Lens with 10 Degree Slope 

Errors at Both Edges

Refracted 

Solar Rays

Stretched 

Fresnel 

Lenses

Solar Cells

Radiators

Figure 5 – Ray Traces for Symmetrical-Refraction 

Lens for a Perfect Lens Shape (Left) and for 

a Lens with Shape Errors (Right) 

NASA/CP—2005-213431 42
While the shape error tolerance of the symmetrical-refraction lens (Fig. 5) is its most dramatic advantage over 

reflective concentrators (Fig. 4), a more complete comparison of these two concentrator approaches is provided 

in Table 1 (on the following page). A conventional planar array is also included in this comparison as the point of 

departure for both concentrator designs. The primary reason for using concentrators instead of planar arrays is to 

save cell area, mass, and cost, and the higher concentration of SLA leads to much higher savings than for the 

reflective concentrator. The space flight heritage of the Fresnel lens concentrators has been excellent, in contrast 

to the reflective approach. Boeing has diagnosed the problems on the 702 communication satellite arrays, and 

believes that outgassing products from the hot photovoltaic panels caused deposits on the mirrors, which were 

within the line-of-sight of the panel outgassing source. This deposition of outgassing products on the mirrors was 

enabled by a “photofixing” process made possible by the direct solar flux on the mirrors [3]. The material 

deposited on the mirrors could then itself outgas and cause deposits on the photovoltaic panels, which were 

within the line-of-sight of the mirror outgassing source. The deposition of outgassing products on the photovoltaic 

panels was enabled by the “photofixing” process made possible by the direct solar flux on the cells [3]. SLA’s 

geometry inherently discourages this complex process, since only the outer lens surface is exposed to direct solar 

flux. SLA’s predecessor, the SCARLET (Solar Concentrator Array using Linear Element Technology) array, 

which performed flawlessly for 38 months on Deep Space 1, did not experience any of the unexpected power 

degradation problems experienced by the Boeing 702 communication satellite arrays. 

As summarized in Table 1, the terrestrial experience for both reflective and refractive concentrator approaches 

has also been problematic for the mirrors, and successful for the lenses. In fact, a 5 MW installation of 2X mirroraugmented silicon photovoltaic panels was installed by ARCO Solar in the middle 1980’s, and degraded rapidly in 

power output [4]. In fact, when the system was being dismantled in the early 1990’s, tests were run on array 

segments with and then without the mirrors in place, and more power was measured with the mirrors removed 

than with them in place [4]. In stark contrast, the line-focus Fresnel lens array at the U.S. Department of Energysponsored Photovoltaics for Utility Scale Applications (PVUSA) test site in Davis, California, outperformed all of 

the other photovoltaic technologies in an independent, side-by-side field test throughout the whole decade of the 

1990’s [7]. 

50-60C 120-130C 75-80C Cell Operating Temperature on 

GEO

Yes – ENTECH’s ErrorTolerant Refractive Optics

No – Reflective Optics Are 

Inherently Sensitive to Minuscule 

Shape Errors (e.g., Hubble 

Telescope Primary)

Yes

Predictable, Same, Unchanging 

Photon Flux on Each Solar Cell 

in Each Source Circuit

Several Successful 

Programs (e.g., Best 

Performance of All 20 kW 

PVUSA Arrays for 10 Years)

Significant Performance Problems 

for Reflective Systems (e.g., ARCO 

Solar 5 MW System at Carissa

Plains – Mirrors Reduced Power)

Numerous 

Successful 

Programs

Terrestrial Experience for Similar 

Technology

Super-Encapsulation of Cells for 

High-Voltage and/or HighRadiation Missions

Optics – Inherent Shape Error 

Problems

Space Flight Heritage and 

Experience

Cell Savings vs. Planar Array:

Area, Mass, and Cost per Watt

Item

Simple and Lightweight –

Auburn Micrometeoroid 

Tests with Cells at 1,000 V

Relatively Difficult and Relatively 

Heavy

Difficult and 

Heavy

Negligible Problem (200X 

Advantage Over Mirrors)

Significant Problem (Efficiency 

and/or Flux Uniformity) N/A

Extremely Successful: 

PASP+ and SCARLET on 

Deep Space 1 (and Geometry 

Discourages Photofixing)

Significant Performance Problems 

on 6 Boeing 702 Comsats (and 

Geometry Is Conducive to 

Outgassing/Photofixing)

Numerous 

Successful 

Programs

0% 50% 85%

8.5X Stretched Lens Array

(SLA)

2X Mirror-Augmented Array

(Reflective Trough Array)

Planar 

Array

50-60C 120-130C 75-80C Cell Operating Temperature on 

GEO

Yes – ENTECH’s ErrorTolerant Refractive Optics

No – Reflective Optics Are 

Inherently Sensitive to Minuscule 

Shape Errors (e.g., Hubble 

Telescope Primary)

Yes

Predictable, Same, Unchanging 

Photon Flux on Each Solar Cell 

in Each Source Circuit

Several Successful 

Programs (e.g., Best 

Performance of All 20 kW 

PVUSA Arrays for 10 Years)

Significant Performance Problems 

for Reflective Systems (e.g., ARCO 

Solar 5 MW System at Carissa

Plains – Mirrors Reduced Power)

Numerous 

Successful 

Programs

Terrestrial Experience for Similar 

Technology

Super-Encapsulation of Cells for 

High-Voltage and/or HighRadiation Missions

Optics – Inherent Shape Error 

Problems

Space Flight Heritage and 

Experience

Cell Savings vs. Planar Array:

Area, Mass, and Cost per Watt

Item

Simple and Lightweight –

Auburn Micrometeoroid 

Tests with Cells at 1,000 V

Relatively Difficult and Relatively 

Heavy

Difficult and 

Heavy

Negligible Problem (200X 

Advantage Over Mirrors)

Significant Problem (Efficiency 

and/or Flux Uniformity) N/A

Extremely Successful: 

PASP+ and SCARLET on 

Deep Space 1 (and Geometry 

Discourages Photofixing)

Significant Performance Problems 

on 6 Boeing 702 Comsats (and 

Geometry Is Conducive to 

Outgassing/Photofixing)

Numerous 

Successful 

Programs

0% 50% 85%

8.5X Stretched Lens Array

(SLA)

2X Mirror-Augmented Array

(Reflective Trough Array)

Planar 

Array

Table 1 – Fundamental Comparison of Different High-Efficiency Space Solar Arrays 

NASA/CP—2005-213431 43
The inherent shape-error tolerance of the symmetrical-refraction lens enables the concentrated solar flux profile 

over the solar cell to be tailored by design and then accurately produced and maintained over the operational life 

of the system. In contrast, minute shape errors in reflective concentrators cause substantial variations in the 

concentrated solar flux profile, not only over an individual solar cell, but also from cell-to-cell in a series string of 

cells forming a source circuit. These photon flux variations from cell-to-cell in a source circuit, and the cell current 

variations they produce, are the Achilles’ heel of reflective concentrator optics, since they can substantially 

degrade the source circuit power output. 

As summarized in Table 1, an additional problem for the mirror-augmented panel (reflective trough) is the high 

cell operating temperature caused by the increased solar flux without a corresponding increase in heat rejection 

area. SLA does not suffer from this problem, since its heat rejection area equals its sunlight-collecting aperture 

area, just like a one-sun planar array. Furthermore, the SLA’s small individual lens size was selected based on 

radiator heat conduction from the line-focus solar cell receiver laterally outward into the surrounding radiator area. 

For the selected 8.5 cm aperture width, only a 125 micron thick graphite cloth radiator is needed to efficiently 

spread the waste heat laterally into the radiator. Thus, the operating cell temperature for SLA is about 50°C 

cooler than for the reflective trough. 

Finally, the SLA enables super-encapsulation of the photovoltaic cell circuit with very little mass penalty, due to 

the small size of the cells compared to the aperture area of the array. Such super-encapsulation enables highvoltage operation of the SLA in the space plasma, reducing cabling size and mass, and minimizing the need for 

voltage-boosting electronics for high-voltage loads such as electronic thrusters. Such super-encapsulation can 

also provide radiation hardness for the solar cells, especially for high-radiation missions or military missions. 

Such super-encapsulation is heavy and expensive for lower concentration arrays or planar arrays, due to the 

much larger cell area per unit power output. 

In summary, the technologies for mirror-augmented photovoltaic panels and Fresnel lens photovoltaic 

concentrators are vastly different. The track records of these two competing concentrator approaches have also 

been completely different, both in space and on the ground. Clearly, experiences, good or bad, for either one of 

these two technologies do not apply to the other. 

The following paragraphs further describe the Fresnel lens photovoltaic concentrator array technology. 

BACKGROUND 

Since 1986, ENTECH and NASA have been developing and refining space photovoltaic arrays using refractive 

concentrator technology [8]. As discussed above, unlike reflective concentrators, these refractive Fresnel lens 

concentrators are configured to minimize the effects of shape errors, enabling straightforward manufacture, 

assembly, and operation on orbit. By using a unique arch shape, these Fresnel lenses provide more than two 

orders of magnitude better shape error tolerance than reflective concentrators or conventional flat Fresnel lens 

concentrators [6]. 

In the early 1990’s, the first refractive concentrator array was developed and flown on the PASP Plus mission, 

which included a number of small advanced arrays [9]. The refractive concentrator array used ENTECH minidome lenses over Boeing mechanically stacked multi-junction (MJ) cells (GaAs over GaSb). The mini-dome 

lenses were made by ENTECH from space-qualified silicone (DC 93-500), and coated by Boeing and OCLI to 

provide protection against space ultraviolet (UV) radiation and atomic oxygen (AO). This array performed 

extremely well throughout the year-long mission in a high-radiation, 70-degree inclination, 363 km by 2550 km 

elliptical orbit, validating both the high performance and radiation hardness of the refractive concentrator 

approach [9]. In addition, in high-voltage space plasma interaction experiments, the refractive concentrator array 

was able to withstand cell voltage excursions to 500 V relative to the plasma with minimal environmental 

interaction [9]. 

NASA/CP—2005-213431 44
In the middle 1990’s, ENTECH and NASA developed a new line-focus Fresnel lens concentrator, which is easier 

to make and more cost-effective than the mini-dome lens concentrator. Using a continuous roll-to-roll process, 

3M can now rapidly mass-produce the line-focus silicone lens material in any desired quantity. In 1994, ABLE 

joined the refractive concentrator team and led the development of the SCARLET solar array [10]. SCARLET 

(Fig. 6) used a small (8.5 cm wide aperture) silicone Fresnel lens to focus sunlight at 8X concentration onto 

radiatively cooled triple-junction cells. Launched in October 1998, the 2.5 kW SCARLET array powered both the 

spacecraft and the ion engine on the NASA/JPL Deep Space 1 probe, shown in Fig. 6. 

SCARLET achieved over 200 W/m2

 areal power density and over 45 W/kg specific power, the best performance 

metrics up to that time [11]. With SCARLET working flawlessly, Deep Space 1 had a spectacularly successful 

rendezvous with the comet, Borrelly, in September 2001, capturing the highest-resolution images of a comet to 

date and other unprecedented comet data. At the end of the 38-month extended mission, in December 2001, 

SCARLET’s power was still within + 2% of predictions. 

Over the past four years, the team, now including Auburn University, has developed an evolved version of the 

flight-proven SCARLET array, called the Stretched Lens Array (SLA), with much better performance metrics [12-

17]. A prototype SLA wing is shown in Fig. 7, and the new SLA approach is described in the following section. 

STRETCHED LENS ARRAY (SLA) DESCRIPTION 

The Stretched Lens Array (SLA) is an evolved version of SCARLET, retaining the essential power-generating 

elements (the silicone Fresnel lens, the multi-junction solar cells, and the composite radiator sheet) while 

discarding the non-power-generating elements (the lens glass arch superstrates, the lens support frames, the 

photovoltaic receiver support bars, and most 

of the honeycomb and back face sheet 

material in the panels). The defining feature 

of SLA (Fig. 7) that enables the elimination of 

so many elements of the SCARLET array is 

the stretched lens optical concentrator (Fig. 

8). By using end arches to stretch the 

silicone Fresnel lens in the lengthwise 

direction only, these lenses become selfsupporting stressed membranes. 

SCARLET’s glass arches are thus no longer 

needed, eliminating their complexity, fragility, 

expense, and mass in the new, patented SLA 

[13]. With this substantial lens-related mass 

reduction, the supporting panel structural 

Figure 8 – Model Showing 

Basic Stretched Lens Approach 

Figure 6 – SCARLET Array on Deep Space 1 Figure 7 – Stretched Lens Array (SLA) Wing 

NASA/CP—2005-213431 45
loads are reduced, making ultra-light panels practical for 

SLA. This cascading mass-reducing effect of the 

stretched lenses continues throughout the SLA wing 

structure, resulting in unprecedented performance 

metrics. 

HARDWARE DEVELOPMENT AND TESTING 

In the past year, the SLA team has fabricated and 

successfully evaluated a subscale four-panel array 

(Fig. 7), including 48 stretched lenses and 4 fully 

functional photovoltaic receivers, each containing 14 

series-connected solar cells (Fig. 9). The four receivers 

included two using triple-junction (GaInP/GaAs/Ge) cells 

from Spectrolab and two using triple-junction cells from 

EMCORE. Two of the receivers used prism covers over 

the cells to eliminate gridline shadowing losses and two 

used more conventional ceria-doped microsheet covers. 

The receivers were assembled as flex circuits, including 

both cells and bypass diodes, and were fully 

encapsulated to enable high-voltage operation. To verify 

high-voltage operation, each of the receivers was wet hipot tested, with 500 V applied between the cell string and 

the composite panel, while distilled water was sprayed 

onto the receiver and panel. The water simulated space 

plasma, which can lead to leakage currents or arcs from 

photovoltaic circuits to the panel structure for 

conventional cell circuits exposed to high voltage 

operation. None of the four circuits had more than 1 

micro-Amp leakage current during the test. Similar wet 

hi-pot testing has been used in the past to verify the 

encapsulation on SLA photovoltaic receiver coupons 

which were successfully tested at 1,000 V in simulated 

space plasma while being subjected to simulated 

micrometeoroid impacts [15]. These latest results extend 

the earlier single-cell coupon high-voltage test results to 

more flight-like 0.5-meter-long SLA receivers. 

The four SLA receivers were performance-tested for onesun performance in ABLE’s Large Area Pulsed Solar 

Simulator (LAPSS), using NASA-Glenn-Lear-Jet-flown 

triple-junction reference cells from the same production 

lots, in addition to individual reference cells of each of the 

three junction types. Then the panel was equipped with 

stretched lenses, and the receivers were again LAPSStested for SLA performance. Fig. 10 shows the measured 

results (AM0, 20°C) for the two receivers using 

Spectrolab cells with and without the lenses installed. 

The solid curves represent the receiver with prism covers 

over the cells, while the dashed curves represent the 

receiver with ceria-doped microsheet over the cells. Note 

that the one-sun efficiency of the prism-covered-cell receiver was approximately 29%, while the net lens/receiver 

efficiency for the same prism-covered-cell receiver was approximately 28%. When these results were corrected 

to the standard reporting temperature of 28°C, the net lens/receiver efficiency for the prism-covered-cell receiver 

was 27.5%. As expected, the net lens/receiver efficiency for the microsheet-covered-cell receiver was slightly 

lower at 26.1% at 28C. 

Figure 9 – Photovoltaic Receivers Undergoing 

500 V Wet Hi-Pot Testing

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0 10 20 30 40

Volts

Amps with Lens

0.00

0.01

0.02

0.03

0.04

0.05

0.06

0.07

Amps Without Lens

Lens LAPSS (Prism-Covered Cells)

Lens LAPSS (Glass-Covered Cells)

One-Sun LAPSS (Prism-Covered Cells)

One-Sun LAPSS (Glass-Covered Cells)

~ 29%

~ 28% Net

Figure 10 – Measured Performance Results 

(Air Mass Zero, 20°C) 

NASA/CP—2005-213431 46
The measured performance results of the 0.5-meter-long lenses and receivers extend the earlier results for single 

lens/cell units which were flown by NASA Glenn on their Lear jet test platform for AM0 current calibration, and 

then LAPSS-tested for AM0 performance. These lens/cell units also had over 27% net lens/cell efficiency (AM0, 

25°C), using prism-covered cells from either Spectrolab or EMCORE [14,15]. All of these results are also in close 

agreement with individual lens optical efficiency measurements and prism-covered solar cell conversion efficiency 

measurements. The stretched lens provides 90% + 2% net optical efficiency in collecting photons and placing 

them onto the triple-junction solar cells. This excellent lens performance is due in large part to the unique colormixing design of the lens, which eliminates chromatic aberration losses in the multi-junction cells [17]. The prismcovered multi-junction cells are over 30% efficient (8 suns AM0, 28°C) in converting sunlight into electricity 

[16,18]. This excellent cell performance is due in large part to the gain in efficiency with concentration, an 11% 

relative improvement for the Spectrolab triple-junction cells [16,18]. 

The mass of every element in the prototype fourpanel wing of Fig. 7 was carefully measured, and 

ABLE used these results to accurately estimate the 

mass of the prototype wing if it had been fully 

populated with 48 photovoltaic receivers under 48 

stretched lenses. The result of this mass estimate 

is 6.46 kg for a fully populated four-panel wing of 

the prototype’s size (2.06 m2

 total wing area). 

ABLE also estimated the on-orbit beginning-of-life 

(BOL) power output of this prototype wing, 

including the power reduction at the GEO 

operating temperature of 75°C compared to the 

standard test temperature of 28°C and all other 

normal array losses (e.g., packing factors, cabling 

and blocking diode losses, etc.). The result of this 

performance estimate is 629 W for a fully 

populated wing of the prototype’s size. Ratios of 

these values provide the prototype wing’s key 

performance metrics of 305 W/m2

 areal power density and 97 W/kg specific power. The areal power density goal 

for a near-term SLA was 300 W/m2

, which has been met. The wing-level specific power value also falls right on 

the predicted curve of SLA specific power versus wing power level (Fig. 11). Note also that NASA’s target value 

for future lightweight solar arrays of 175 W/kg for a 7 kW wing, as detailed in the recent New Millennium Program 

Space Technology 8 (NMP ST8) procurement is exceeded by SLA. 

In addition to verifying the performance and mass estimates for SLA, the prototype four-panel wing of Fig. 7 also 

validated, via numerous deployments, the basic SLA mechanical approaches for deploying the four panels and 

the 12 lenses on each panel (Fig. 12). 

The following section describes a point design of an SLA solar array wing rated at 7 kW of output power, a wing 

size typical of today’s communication satellites. 

0

50

100

150

200

250

0 2,000 4,000 6,000 8,000 10,000 12,000 14,000 16,000

BOL On-Orbit Wing Power (W)

Specific Power (W/kg)

Calculated Rigid Panel SLA

NMP ST8 Target

Fully Populated SLA Demo Wing

Figure 11 – SLA Specific Power vs. Wing Power on 

Geostationary (GEO) Orbit at Beginning of Life (BOL)

600 s 240 s 30 s 10 s 5 s 0.5 s 0 s

Figure 12 – SLA Prototype Wing Deployment

NASA/CP—2005-213431 47
POINT DESIGN OF A 7 KW SLA WING 

As shown in Fig. 11, the performance advantages 

of SLA increase rapidly with increasing wing size, 

with the specific power reaching more than 180 

W/kg at the important wing size of 7 kW. Since 

this wing size is typical of the latest generation of 

GEO communication satellites, the SLA team has 

generated a detailed point design for a 7 kW wing, 

as shown in the sketch and table below (Fig. 13). 

This wing design is conventional and conservative 

in configuration and components, with robust 

structural stiffness parameters. The beginning-oflife (BOL) areal power density is slightly lower 

than already demonstrated on the prototype 

hardware described above. Future multi-junction 

cells will no doubt eclipse the demonstrated 

values, providing even higher performance 

metrics than those shown in Fig. 13. While cost 

values are not presented, the SLA team has 

analyzed costs in detail, and due to more than 

85% savings in cell area and cost compared to a 

planar multi-junction array, the SLA wing will cost 

50% less than a planar wing of equal power. 

RELATED WORK 

The SLA team is also performing space 

environmental effects testing of key SLA 

components [19,20]. This ongoing testing 

includes solar ultraviolet and charged particle 

radiation exposure of stretched lenses, and 

micrometeoroid impact testing of lenses and 

photovoltaic receivers at high voltage in simulated 

space plasma. In addition, the SLA team has 

analyzed additional SLA performance metrics for 

a wide variety of different missions, and in each 

case, SLA provides excellent advantages over the 

planar array competition [21]. 

ALTERNATE VERSIONS OF SLA

While the rigid-panel wing, described in previous paragraphs, is the near-term, conservative embodiment of SLA 

for near-term missions, alternate flexible-blanket versions of SLA are also under development. These versions 

use lenses and photovoltaic receiver/radiator elements that accordion fold for compact launch stow volume, and 

then deploy as an end-tensioned dual-blanket (stretched lenses on top and stretched radiator sheets with 

photovoltaic receivers on bottom) array on orbit. In fact, Fig. 2 shows a small model of the flexible-blanket SLA. 

One of the most attractive approaches for deploying and supporting the flexible-blanket version of SLA on orbit is 

ABLE Engineering’s SquareRigger platform, originally developed for the Air Force Research Laboratory. SLA on 

SquareRigger could provide very large power arrays (50 kW to MW class) for a variety of future space missions. 

Fig. 14 (on the following page) shows a schematic of the SLA/SquareRigger array. The stowed package of 

SquareRigger is a tight bundle of structural tubes with the flexible blanket elements (flexible silicone lenses and 

composite radiator sheets with photovoltaic receivers mounted to the sheets) folded tightly between tubes. The 

tubes deploy first to form initially empty rectangular bays. 

7,129 W, 23.78 m2, 38.9 kg

Rigid-Panel Stretched Lens Array

Low-Mass “Picture Frame” Panels

Inboard and Outboard Panels:

Standard Construction Graphite

Composite Honeycomb Sandwich Panels

7,129 W, 23.78 m2, 38.9 kg

Rigid-Panel Stretched Lens Array

Low-Mass “Picture Frame” Panels

Inboard and Outboard Panels:

Standard Construction Graphite

Composite Honeycomb Sandwich Panels

Stowed Power Easily implemented on outer panel

Ease of Adding Planar Panel Easily implemented on outer panel

Flatness & Warping Well understood flat stable platform

Deployment Testing Can use existing off-loaders

Pop-up lenses allow each panel to be

tested as a complete assembly before

wing integration

Power Testing

Easier to integrate on commercial

spacecraft. Readily accepted

configuration.

Commercial Appeal

Stowed Stiffness 40 Hz

Deployed Stiffness 0.1 Hz

Specific Power 183 W/kg

0.11 m3 Stowed Volume /kW

Base Platform Design Maturity Most components flight proven on DS1

SLA Implementation Pop-up lenses

Self Shadowing No self shadowing

Point Design Basis 7,129 Watts (BOL)

Feature Value or Characteristic

Stowed Power Easily implemented on outer panel

Ease of Adding Planar Panel Easily implemented on outer panel

Flatness & Warping Well understood flat stable platform

Deployment Testing Can use existing off-loaders

Pop-up lenses allow each panel to be

tested as a complete assembly before

wing integration

Power Testing

Easier to integrate on commercial

spacecraft. Readily accepted

configuration.

Commercial Appeal

Stowed Stiffness 40 Hz

Deployed Stiffness 0.1 Hz

Specific Power 183 W/kg

0.11 m3 Stowed Volume /kW

Base Platform Design Maturity Most components flight proven on DS1

SLA Implementation Pop-up lenses

Self Shadowing No self shadowing

Point Design Basis 7,129 Watts (BOL)

Feature Value or Characteristic

Figure 13 – Point Design for 7 kW SLA Wing

NASA/CP—2005-213431 48
As shown in Fig. 14, after tube deployment is complete, a motor automatically unfolds the flexible blanket 

photovoltaic array from one end to the other until the bay is fully populated. While SquareRigger was initially 

envisioned as a thin-film photovoltaic blanket 

platform, it is ideally suited to the more efficient 

SLA technology. The combination of SLA and 

SquareRigger provides an unmatched set of 

performance metrics for large-capacity space 

power arrays, as summarized in Table 2. 

ABLE Engineering has recently completed a 

NASA Phase I SBIR contract related to the 

integration of SLA with SquareRigger, including 

fabrication of a small demonstration unit. The 

results of this contract have confirmed the expected advantages of marrying SLA to SquareRigger, as discussed 

in another paper at this SPRAT XVIII Conference. 

In addition to space applications, the combination of color-mixing symmetrical-refraction Fresnel lenses and multijunction photovoltaic cells should have spin-off applications in the terrestrial solar energy marketplace. 

Developments are already underway on terrestrial versions of these high-performance photovoltaic concentrators. 

Indeed, numerous outdoor performance tests of a space-optimized SLA mini-concentrator module have 

demonstrated over 30% net-aperture-area solar-to-electric conversion efficiency. These results are believed to 

represent the first time that the 30% efficiency threshold has ever been broken by any type of solar energy 

converter tested outdoors under natural terrestrial sunlight [15]. 

Time Frame < 5 Years 5-10 Years

Power Capability (kW) 100 1,000

BOL Specific Power (W/kg) 330 500

Stowed Power (kW/m3

) 80 120

Voltage 1,000 TBD

Table 2 – Performance Attributes of SLA

on ABLE’s SquareRigger Platform 

Folded blanket 

between stowed 

tubes (typical)

Stowed Array Cross-Section

All bays deploy as 

shown in detail for 

this one

Bays finish deployment & latch 

prior to blanket deployment 

Folded SLA 

blanket 

Motor starts blanket 

deployment 

1 2 3 4

5 6

Dual-deck 

lens/cell panel 

blanket 

Deployment complete

Folded blanket 

between stowed 

tubes (typical)

Stowed Array Cross-Section

All bays deploy as 

shown in detail for 

this one

Bays finish deployment & latch 

prior to blanket deployment 

Folded SLA 

blanket 

Motor starts blanket 

deployment 

1 2 3 4

5 6

Dual-deck 

lens/cell panel 

blanket 

Deployment complete

Figure 14 – Stretched Lens Array (SLA) on SquareRigger Platform 

NASA/CP—2005-213431 49
CONCLUSIONS 

A new type of space solar power system is being developed with unprecedented performance and mass 

properties. The Stretched Lens Array (SLA) uses ultra-thin refractive optical elements to collect and focus 

sunlight onto narrow state-of-the-art multi-junction photovoltaic cells, which are mounted to ultra-thin composite 

radiator sheets. SLA is being developed in two versions: 

z A rigid-panel SLA, which uses a conventional and low-risk approach to deployment and support on orbit. This 

version of SLA offers substantially improved performance metrics for near-term solar array wings up to 20 

kW. 

z A flexible-blanket SLA, which uses more advanced and higher performance approaches to deployment and 

support on orbit. This version offers transformational performance metrics for longer-term solar array wings in 

the 50 kW to MW class. 

Hardware developments to date have verified the performance, mass, and cost advantages of the new SLA 

technology. 

ACKNOWLEDGMENT 

The authors gratefully acknowledge NASA’s support of the work presented in this paper. 

REFERENCES 

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2. P.B. deSelding, “Space Insurance Continues to Be Seen as Risky Market,” Space News, January 14, 2003. 

3. M. Bodeau, “Root-Cause of the 702 Concentrator Array Anomaly,” Presentation at the 2003 Space Power Workshop, 

Redondo Beach, April 2003. 

4. H.J. Wenger et al., “Decline of the Carissa Plains PV Power Plant: The Impact of Concentrating Sunlight on Flat Plates,” 

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Applications, John Wiley & Sons, 1995. 

7. PVUSA Project Team, "1998 PVUSA Progress Report," USDOE Contract No. DE-FC04-96AL89774, 1999. 

8. M.F. Piszczor and M.J. O’Neill, “Development of a Dome Fresnel Lens/GaAs Photovoltaic Concentrator for Space 

Applications,” 19th IEEE-PVSC, 1987. 

9. H. Curtis and D. Marvin, “Final Results from the PASP Plus Flight Experiment,” 25th IEEE-PVSC, 1996. 

10. P.A. Jones et al., “The SCARLET Light Concentrating Solar Array,” 25th IEEE-PVSC, 1996. 

11. D.M. Murphy, “The SCARLET Solar Array: Technology Validation and Flight Results,” Deep Space 1 Technology 

Validation Symposium, Pasadena, 2000. 

12. M.J. O’Neill, “Stretched Fresnel Lens Solar Concentrator for Space Power,” U.S. Patent 6,075,200, 2000. 

13. M.J. O’Neill et al., “Development of the Ultra-Light Stretched Lens Array,” 29th IEEE-PVSC, 2002. 

14. M.J. O’Neill et al., “The Stretched Lens Array (SLA): A Low-Risk, Cost-Effective Concentrator Array Offering Wing-Level 

Performance of 180 W/kg and 300 W/m2

 at 300 VDC,” 37th IECEC, 2002. 

15. M.J. O’Neill et al., “The Stretched Lens Array (SLA),” IEEE Aerospace and Electronic Systems Magazine, Vol. 18, No. 1, 

January 2003. 

16. M.F. Piszczor et al., “Stretched Lens Array (SLA) Photovoltaic Concentrator Hardware Development & Testing,” Third 

World Conference on Photovoltaic Energy Conversion, Osaka, 2003. 

17. M.J. O’Neill, “Color-Mixing Lens for Solar Concentrator System and Methods of Manufacture and Operation Thereof,” U.S. 

Patent 6,031,179, 2000. 

18. A. Stavrides et al., “Fabrication of High-Efficiency III-V Multi-Junction Solar Cells for Space Concentrators,” 29th IEEEPVSC, 2002. 

19. H.W. Brandhorst et al., “Hypervelocity Impact Testing of Stretched Lens Array Modules,” Sixth European Space Power 

Conference, Oporto, Portugal, 2002. 

20. D.L. Edwards et al., “Optical Analysis of Transparent Polymeric Material Exposed to Ultraviolet Radiation,” 8th International 

Symposium on Materials in a Space Environment, Arcachon, France, 2000. 

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Photovoltaic Energy Conversion, Osaka, 2003. 

NASA/CP—2005-213431 50
Materials on the International Space Station - Forward Technology Solar Cell Experiment 

R. J. Walters, J. C. Garner, S. N. Lam, J. A. Vazquez, W. R. Braun, and R. E. Ruth 

US Naval Research Laboratory, Washington, DC 20375 

J. R. Lorentzen 

SFA Inc. 

Cdr R. Bruninga (Ret.) 

US Naval Academy, Annapolis, MD 21402 

P. P. Jenkins 

Ohio Aerospace Institute, Cleveland, OH 44135 

J. M. Flatico 

QSS, Cleveland, OH 44135 

D. M. Wilt, M. F. Piszczor, L. C. Greer, and M. J. Krasowski 

NASA Glenn Research Center, Cleveland, OH 44135 

INTRODUCTION 

This paper describes a space solar cell experiment currently being built by the Naval Research Laboratory (NRL) 

in collaboration with NASA Glenn Research Center (GRC), and the US Naval Academy (USNA). The 

experiment has been named the Forward Technology Solar Cell Experiment (FTSCE), and the purpose is to 

rapidly put current and future generation space solar cells on orbit and provide validation data for these 

technologies. The FTSCE is being fielded in response to recent on-orbit and ground test anomalies associated 

with space solar arrays that have raised concern over the survivability of new solar technologies in the space 

environment and the validity of present ground test protocols. The FTSCE is being built as part of the Fifth 

Materials on the International Space Station (MISSE) Experiment (MISSE-5), which is a NASA program to 

characterize the performance of new prospective spacecraft materials when subjected to the synergistic effects of 

the space environment. Telemetry, command, control, and communication (TNC) for the FTSCE will be 

achieved through the Amateur Satellite Service using the PCSat2 system, which is an Amateur Radio system 

designed and built by the USNA. In addition to providing an off-the-shelf solution for FTSCE TNC, PCSat2 will 

provide a communications node for the Amateur Radio satellite system. The FTSCE and PCSat2 will be housed 

within the passive experiment container (PEC), which is an approximately 2ft x2ft x 4in metal container built by 

NASA Langley Research Center (NASA LaRC) as part of the MISSE-5 program. NASA LaRC has also supplied 

a thin film materials experiment that will fly on the exterior of the thermal blanket covering the PCSat2. The PEC 

is planned to be transported to the ISS on a Shuttle flight. The PEC will be mounted on the exterior of the ISS by 

an astronaut during an extravehicular activity (EVA). After nominally one year, the PEC will be retrieved and 

returned to Earth. At the time of writing this paper, the subsystems of the experiment are being integrated at 

NRL, and we are preparing to commence environmental testing. 

NASA/CP—2005-213431 51
The Space Test Program (STP) is responsible for the integration of MISSE-5 with the Shuttle and the launch of 

MISSE-5. All public release of information concerning the Spaceflight of the MISSE5 payload will be 

coordinated and approved by STP and NRL. This includes the release of any information associated with 

MISSE-5 space flight experiment integration or operations. 

MATERIALS ON THE INTERNATIONAL SPACE STATION EXPERIMENTS 

The Materials on the International Space Station 

Experiments (MISSE) is a NASA program 

designed to provide access to space for new 

materials and devices being considered for use in 

space (http://misse5.larc.nasa.gov/index.html). 

This is accomplished by placing experiments into 

the Passive Experiment Container (PEC) which is a 

metal box approximately 2 ft x 2ft x 4 inches 

fabricated by NASA Langley Research Center 

(Figure 1). As shown in Figure 1, the 

experiments are mounted on custom designed trays 

that mount within the PEC. When closed, the PEC 

provides the container for the experiments for 

launch and transfer to the ISS on the Shuttle. For 

deployment, the PEC is clamped to a handrail on 

the exterior of the ISS by an astronaut who then 

opens the PEC to expose the experiments (Figure 

2). The FTSCE will be the primary experiment on 

the 5th MISSE project. 

DESCRIPTION OF THE FTSCE EXPERIMENTS 

A total of 39 solar cells are included in the FTSCE. The 

technologies include state-of-the-art and next generation 

multijunction InGaP/GaAs/Ge, heteroepitaxial 

GaAs/GeSi/Ge, and amorphous Si and CuIn(Ga)Se2 thin 

film solar cells. A photograph of the deck that holds the 

experiments is shown in Figure 3. This deck is designed 

as a heat shield to help keep the interior of the PEC warm 

and is referred to as the Thermal Deck. The experiments 

are identified in Table 1. The primary experiments are 

the triple-junction (3J) InGaP/GaAs/Ge based 

technologies from Spectrolab (SPL) and Emcore. In each case, the current state-of-the-art technology (ITJ for 

SPL and ATJM for Emcore) and the next generation technology (UTJ for SPL and BTJ for Emcore) are included. 

In addition, the Emcore ATJM devices include the new monolithic bypass diode. The SPL panel includes two DJ 

InGaP/GaAs/Ge solar cells that serve as control cells. The Emcore and SPL experiments were assembled by the 

manufacturer according to their standard practices on aluminum honeycomb rigid array substrates. 

Figure 1: This is a photograph of a Passive Experiment 

Container (PEC). The PEC is built by NASA and is designed 

to hold experiments mounted on trays as shown. The PEC is 

closed to protect the experimental samples during transport on 

the Shuttle and deployment on the ISS. Once deployed, the 

PEC is opened to expose the experiments to the space 

environment. At the end of the mission, the PEC is closed and 

returned to Earth. 

Figure 2: This is a photograph of an astronaut 

deploying the second PEC onto the exterior of the ISS. 

NASA/CP—2005-213431 52
Laboratory development technologies are also included in the FTSCE. There are several single-junction (SJ) 

GaAs solar cells grown on Si substrates that are representative of the GaAs/SiGe/Si technology being developed 

jointly by NASA GRC, Ohio State University (OSU), and Massachusetts Institute of Technology (MIT). These 

cells, along with GaAs/Ge control cells, are mounted on a rigid Al honeycomb substrate. There are four 3J 

InGaP/InGaAs/Ge solar cells supplied by SPL. These are the metamorphic cells that employ a stoichometry that 

results in a slightly lattice mismatched semiconductor 

stack, which, in turn, results in a bandgap combination 

more closely optimized for the air mass zero (AM0) 

spectrum. These solar cells are mounted on the SPL panel. 

The FTSCE also includes flexible, thin-film solar cell 

(TFSC) technologies. Because the flexible TFSC 

experiments represent not only a shift in solar cell 

technology but also a potential dramatic change in future 

solar blanket and array technology designs, the FTSCE 

TFSC experiments were fabricated to mimic as closely as 

possible functional “first-generation” thin-film blanket 

technology. Each experiment includes multiple, 

interconnected cells mounted on a light weight array 

substrate. Two of the experiments incorporate amorphous 

silicon (a-Si) photovoltaic technology. One a-Si sample 

utilizes monolithically-interconnected cells grown by Iowa 

Thin Film (ITF) on a Kapton substrate. This unit was 

provided by Lockheed Martin and the Aerospace 

Corporation and represents a variety of interconnect, 

laydown and coating techniques developed for their thinfilm program. The other a-Si experiment consists of a-Si 

material from UniSolar grown on a stainless steel (SS) 

substrate using their commercial production process. The 

cells for this experiment were integrated by AEC-Able Engineering using an adaptation of their UltraFlex blanket 

design and represents a near-term attempt to “space-qualify” thin-film cell technology when integrated into a 

viable light weight solar array design. This sample consists of two interconnected a-Si on SS cells affixed to a 

Vectran gore weave, which simulates the deployed conditions on the UltraFlex array design. It will test cell-tocell interconnects, cell-to-array attachments and coating technology under long-term space environmental 

conditions. AEC-Able/NASA Glenn also supplied a CuIn(Ga)Se2 (CIGS) experiment integrated into the same 

UltraFlex blanket design as described above. This experiment consists of five CIGS cells interconnected in series 

using a “shingled” approach. The CIGS cells were provided by ITN/Global Solar. It is important to note that the 

three FTSCE thin film experiments are primarily thin-film blanket technology durability tests. As noted in Table 

1, the efficiency of some of the thin-film devices being flown do not represent the current achievable performance 

of that technology and were selected because of availability or adaptability to specific thin-film blanket 

technologies. 

The FTSCE includes an experiment to test the environmental durability and long-term transmittance of silicone 

materials. Certain silicones, such as DC 93-500 that is currently used as a transparent adhesive to affix a 

coverglass to crystalline solar cells, are being considered for use as concentrator lens materials or coatings for 

advanced photovoltaic devices. It is critical to quantitatively understand the long-term performance of these 

materials, specifically optical transmittance degradation due to UV darkening, for such applications. The Silicone 

Degradation Experiment consists of films of silicone (DC 93-500) attached to two different solar cell 

coverglasses. These two samples are placed over multijunction solar cells so that optical transmission (relative to 

the response of the MJ cells) can be monitored throughout the duration of the flight experiment. Two other bare 

2

3 

4

5 

6 

7 

8 

10 9 

1 

Figure 3: This is a photograph of the FTSCE 

experiment deck to be placed into the MISSE5 PEC. 

The experiments are numbered in the picture and 

described in Table 1. Note that the contamination 

monitor (6) and sun sensors (7) are not in this picture.

NASA/CP—2005-213431 53
silicone samples are being flown as passive experiments and will be evaluated upon return to earth at the 

completion of the mission. 

A passive contamination monitor will fly on the FTSCE. The monitor consists of a piece of CMX coverglass 

mounted such that a large surface area of the glass is exposed to the space environment. Transmission and 

reflectance measurements were made on the glass prior to integration, and these measurements will be repeated 

upon return to Earth. Comparison of the pre- and post-flight data will enable an evaluation of the effect of 

contamination on the output of the experimental solar cells. The theory is that any loss in transmission due to 

contamination will be observed in the coverglass data, and assuming uniform contamination of the FTSCE 

surface, this will translate into a decrease in photocurrent of all of the experimental cells. 

DESCRIPTION OF THE THIN FILM MATERIALS EXPERIMENT 

As described more fully below, one half of the 

MISSE 5 PEC will hold the PCSat2 system 

which is the RF communications subsystem. 

This half of the PEC is covered by a thermal 

blanket, and the PEC will be oriented on orbit 

so that this side is always facing away from the 

sun. The NASA Langley team has transformed 

the outer layer of the thermal blanket into a 

three and one half ounce experiment to evaluate 

the in space survivability of 200 advanced 

materials that are being developed to enable 

future US space missions. The survivability of 

these materials will be established by 

comparing pre and post flight characterization 

test data. The exposure conditions anticipated 

for this experiment are unprecedented in 

previous space exposure experiments. Since 

these specimens will be facing the anti-solar 

direction when mounted on the ISS, they will 

receive no UV radiation that can rupture 

chemical bonds and provide reactive free-radical sites. They will be subjected to thermal cycling, particulate 

radiation and atomic oxygen bombardment. With no UV exposure, all reactions that do occur will be essentially 

chemically driven oxidation. A photograph of the experiment package is shown in Figure 4. 

DATA ACQUISITION ELECTRONICS AND SOFTWARE 

The data acquisition electronics were designed and built by the NASA GRC team. A photograph of the 

electronics is shown in Figure 6. In Figure 6, the electronics are shown mounted to the deck (named the 

Electronics Deck) that will hold the boards within the PEC. The Thermal Deck holding the solar cell experiments 

(Figure 3) is mounted on stand-offs on the opposite side of the Electronics Deck. The two layer unit will be 

mounted in one half of the PEC. The electronics consists of one “main” microprocessor board and nine data 

acquisition (DAC) boards. The main microprocessor board provides the communications link with the TNC of 

PCSat2, serves as the command interpreter, and controls the DAC boards. In addition, the main microprocessor 

includes dual redundant flash memory so that not only is data transmitted to ground, but is also archived on board. 

Upon return to Earth, and in the event of communication downlink failure, the mission will still have data 

available. 

Figure 4: This is the Thin Film Materials Experiment fielded by 

NASA Langley Research Center. This 3 & 1/2 ounce experiment 

will evaluate in-space survivability of 200 advanced space 

materials. Specimens are attached to thermal blanket outer layer. 

Specimen survivability determined from comparisons of pre and 

post flight laboratory characterization tests data 

NASA/CP—2005-213431 54
Each DAC board is approximately 4x6 in2

and is capable of measuring a 32 point IV 

curve on four individual solar cells, making 

two temperature measurements using AD590 

temperature sensors, and taking data from one 

sun angle sensor. The two temperature 

channels on a board can be combined to 

measure temperature using a resistance 

temperature device (RTD), which provides a 

wider operating range than the AD590. A 

single temperature channel on a board can 

also be configured to make a single IV point 

measurement on a cell. The IV curve is 

created by using a FET as a variable resistor 

and thereby sweeping the load resistance 

while measuring the solar cell current and cell 

voltage. Considering that most solar cell 

experiments employ a bank of switched load 

resistors for making the IV measurement, 

which requires much more space and weight, 

these measurement boards are a significant 

improvement. A comparison of data 

measured by one of the DAC boards on a 3J 

InGaP/GaAs/Ge solar cell under illumination 

by the X-25 solar simulator in the NRL Solar 

Cell Characterization Laboratory with data 

measured by laboratory equipment under 

illumination by the same simulator is shown 

in Figure 6, and the agreement can be seen to 

be excellent. 

The data acquisition software was also 

designed and written by the NASA GRC 

team. Portions of the software reside on each 

of the DAC boards and the main 

microprocessor. The DAC board software is 

responsible for taking commands from the 

main microprocessor, returning data to the 

main microprocessor, and performing the IV 

curve, temperature, and sun angle 

measurements. The software is designed to 

autonomously take data when user defined 

conditions of sun angle and temperature are 

met. Also, the experiment can be 

commanded to measure on demand. In 

addition, each DAC board can be commanded 

individually with a specific set of 

measurement criteria. This grants the scientists on the ground flexibility in creating experimental data sets. For 

example, the DAC boards can be commanded to measure data once the sun angle is below a set threshold, which 

allows IV data vs angle of incidence data to be generated. Alternatively, the DAC boards can be commanded to 

Figure 6: This is a photograph of the data acquisition 

electronics mounted on the electronics deck. There is a single 

“main” microprocessor board (0) that controls nine “daughter” 

boards (1-9), which measure the IV curve, temperature, and sun 

angle data. The metal box in the upper right of the photograph 

is the power control unit (PCU). The metal boxes in the center 

are feed-throughs that route the wires from the solar cells that 

are mounted on the opposite side. These boxes serve to 

maintain a Faraday cage around the measurement electronics. 

(0)

(1)

(2)

(3)

(4)

(5)

(6)

(7)

(8)

(9)

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0.45

0.5

0 0.5 1 1.5 2 2.5 3

Volts

Amps

GRC DAC 1

GRC DAC 2

GRC DAC 3

NRL 1

NRL 2

NRL 3

Average Isc: DAC=443mA, NRL=439mA

AverageVoc: DAC=2.613V, NRL=2.612V

Figure 6: This graph compares IV data measured on one of the 

FTSCE 3J InGaP/GaAs/Ge solar cells by a flight DAC board and 

flight software compared to data measured on the same cell by 

laboratory equipment. In each case, three data sets were measured 

in succession, and the three curves are shown. All measurements 

were made under the same X-25 solar simulator in the NRL Solar 

Cell Characterization Laboratory. 

NASA/CP—2005-213431 55
measure once the temperature has exceeded a set threshold value, which allows IV data vs. temperature to be 

generated. 

The software resident on the main microprocessor board is responsible for receiving command from the TNC of 

PCSat2, decoding the commands, and passing the appropriate commands to the DAC boards. The main 

microprocessor software must also take the data from the 9 DAC boards, translate it into printable ASCII 

characters and pass it to the TNC for down-linking by PCSat2. The main microprocessor software must also 

“oversee” the autonomous operation of the experiment, which consists of periodically recording readings from all 

of the temperature sensors to give a temperature profile for the PEC during each orbit and monitoring the 

temperature and sun angle data and determining if the measurement conditions have been reached. 

POWER SUBSYSTEM 

In sunlight the experiment is powered by an array of 4 strings of Emcore ATJ solar cells (the power array is 

pictured in Figure 3). Each string has 9 cells wired electrically in series to produce an open circuit voltage (Voc) 

of approximately 23.09 Volts. Power from the 4 strings is fed into the Power Control Unit (PCU). The PCU 

distributes power to the GRC data acquisition electronics and the PCSat2 communications system and regulates 

charge current to the battery. When the experiment power array is shadowed, by either the ISS itself or normal 

orbit eclipses, the experiment is powered by 4 high capacity (55 A-hr) prismatic Lithium Ion (Li-ion) batteries. 

This will be the first flight of this battery technology and one of the first flights of a Li-Ion battery system in a low 

Earth orbit (LEO) space environment. Lithion of Pawcatuck, Connecticut, manufactured the battery cells. The 

cells are wired together in series to provide a bus voltage of 12.0 to 16.0 volts. The PCU uses a shunt regulator to 

reduce the charge current to the battery when the battery voltage reaches 16.0 volts. The thin film shunt resistors 

are taped to the inside of the PEC. An under-voltage detection circuit sheds all non-critical loads when the battery 

voltage drops below 11.5 volts. 

It was determined that RF emissions from the PCSat2 transmitters could cause electromagnetic interference with 

the astronauts’ EVA suit, so a multiple inhibit scheme has been implemented to prevent RF transmission when the 

astronauts are in the vicinity of the PEC. The power system is inactive for launch and EVA transportation to the 

experiment site on the ISS. Solar array power and battery connection to the experiment are only enabled after an 

astronaut installs a turn-on plug and toggles a power switch. Power to the transmitters, however, remains delayed 

by an 8 hour timer circuit to allow sufficient time for the astronauts to exit the area around the PEC. 

PCSAT2 – RF COMMUNICATIONS SUBSYSTEM 

The RF communications subsystem is called PCSat2. PCSat2 is an Amateur Satellite Communications system 

similar to what it is flying on PCsat as an external Amateur Radio on the ISS (ARISS) payload. The PCSat2

subsystem will operate in the ITU Amateur Satellite Service in cooperation with ARISS and provide a PSK-31 

multi-user transponder, an FM voice repeater for possible use with ISS Crew communications and an AX.25 

packet system for use as a UI digipeater and for telemetry, command, control. PCSAT2 will use the same dual 

redundant AX.25 command and control system as used on PCsat (NO-44) offering 8 on/off commands, 5 

telemetry channels and a serial port for the solar cell experiment telemetry. It also supports the Digital Comms 

Relay support of the PCsat2/APRS mission. The packet uplink is on 145.825 MHz and the default downlinks are 

in the 435 MHz band to avoid any possible interference with existing ARISS missions. PCSAT2 will have quad 

redundant transmit inhibits for EVA safety issues, thus, it is also easy to turn off to avoid any issues with other 

UHF ARISS experiments that may be activated in the future.

One of the key issues with any Amateur Radio experiments on ISS is the requirement to avoid any mutual 

interference between systems. For this reason, ARISS will need to eventually move all uplinks and downlinks 

into separate bands. This is so that multiple uplinks and multiple downlinks can be going on simultaneously. As 

it is, with both uplinks and downlinks on 2m, that band cannot be shared without mutual interference. Thus 

NASA/CP—2005-213431 56
PCSat2 is designed as mode J to avoid transmitting on 2m, even though mode B (downlinks on 2m) is far superior 

to reaching schools and low-tech stations and meeting our mission objectives. There is much UHF equipment 

being planned for ISS, but until it is operational and ARISS has a long range plan, mode J must be used. 

DEPLOYMENT ON ISS 

In the MISSE Program, experiments are packaged inside a PEC which is then transported to the ISS via the 

Shuttle. Once aboard the ISS, the PEC is taken outside the Station by an astronaut during an EVA and clamped 

onto an external handrail. For MISSE-5, the requirement for favorable sun exposure for the experiments and for 

power generation made the choice of the location of the PEC critical. After several iterations of lighting, thermal, 

and power budget analysis, the P4 trunion handrail was chosen. This handrail is located on the port-side solar 

array truss section between the solar arrays. This section of the ISS is scheduled to be installed on ISS Mission 

12A.1/STS Flight 116, and MISSE-5 was scheduled to be installed at the same time. The original STS-116 

launch date was July 14, 2003, but the loss of the Space Shuttle Columbia has delayed this date. In an effort to be 

deployed sooner, alternate locations are being investigated. In particular, a handrail on the crew airlock, which is 

the current site of MISSE-2, has been identified as a viable alternative. Another possible alternative is on the P6 

truss section between the existing solar arrays. The important point gained from the analysis of possible locations 

is that the amount of sun that the experiment will see during a given orbit and the incident solar angle will vary 

significantly over the duration of the mission depending on the ISS orbital attitude. Furthermore, certain locations 

may experience extended periods of shadow during Shuttle docking. This proved to be the major driver in the 

design of our thermal management system, power subsystem, and experimental data acquisition protocols. 

CONCEPT OF OPERATIONS 

The plan is for MISSE-5 to remain on orbit for one year. At the end of the mission, the PEC will be closed and 

returned to Earth. Once returned, the FTSCE will be disassembled to allow for full characterization of the 

individual experiments. While on orbit, the FTSCE will experience the LEO orbital environment of the ISS. The 

orbit has a 90 minute period with typically 60 minutes of sunlight and a thirty minute eclipse. The experiment is 

designed to track the illumination conditions on the face of the FTSCE and the temperature. As the illumination 

and temperature approach the user-defined set points, the main microprocessor board energizes the DAC boards. 

When the measurement conditions have been met, the main microprocessor initiates the data measurement, and 

the DAC boards measure all 36 IV curves within one minute. When complete, the DAC boards pass the data to 

the main microprocessor, which, in turn converts the data into Automatic Position Reporting System (APRS) 

format and passes it to the TNC via a serial port. PCSat2 telemeters the data to the ground as AX.25 packets at 

9600 bps. To maximize probability of reception, the data is transmitted as a beacon, repeating approximately 

every 3 minutes. The primary ground station is the USNA where the data will be received and transferred to NRL 

for analysis. The back-up ground station is NASA GRC. In addition, the data can be received by any amateur 

radio station operating in AX.25 mode. Once received, the data is placed on the Internet from which it will be 

received by NRL. 

NASA/CP—2005-213431 57
Table 1: This table identifies and gives a brief description of the individual solar cell technologies included in the FTSCE . 

The approximate solar cell size is given and the number of individual cells is given in parentheses. The beginning of life 

(BOL) efficiency is the average of the data measured by the responsible organization and given to NRL with the experiments.

* - Denotes atypical cell efficiency due to available selection/modification of cells for flight experiment. 

# Experiment Name Responsible 

Organization 

Description BOL 

Eff (%) 

ITJ Spectrolab 3J commercial ITJ cells, 26.6 cm2 (10) 27.0 

UTJ Spectrolab 3J commercial UTJ cells, 26.6 cm2 (5) 28.6 

Metamorphic Spectrolab 3J lattice-mismatched cells, 4 cm2 (4) 28.5 1 

Control Spectrolab DJ InGaP/GaAs/Ge control cells, 27 cm2

 (2) 22.9 

Si Degradation NASA 

GRC/Entech 

Films of DC 93-500 silicone mounted on coverglass 

superstrate over MJ cell to measure the transmittance 

degradation of silicone material, 4 cm2

, (2 active, 2 passive) 

N/A 

2 

USNA Solar Cells USNA/NASA 

GRC 

Commercial-off-the-shelf terrestrial solar cells, passive 

experiment, 2 are ~3.5 cm2

 and the third is ~7.75cm2, (3) 7.0* 

3 Amorphous Si on 

Kapton 

NASA 

GRC/LM/Iowa 

Thin Films 

Blanket-level test of interconnected thin film amorphous 

silicon cells on Kapton substrate (3 monolithicallyinterconnected cells for 1 active sample, 2 passive samples) 0.24* 

4 Power Panel Emcore Primary power panel for mission, ATJ cells N/A 

5 GaAs on Si NASA 

GRC/OSU/MIT 

SJ GaAs cells grown on SiGe/Si substrates with one 

GaAs/GaAs control, 1 & 4 cm2

, (7) 12.2 

6 

Amorphous Si on 

Stainless Steel 

NASA 

GRC/AEC 

Able/UniSolar 

Blanket-level test of interconnected thin film amorphous 

silicon cells on stainless steel substrate affixed to a Vectran 

gore sheet (simulating UltraFlex array design), ~20.5cm2

 (2 

series-interconnected cells) 

10.4 

7 Contamination 

Monitor NRL Thales 5 mil CMX coverglass, passive witness plate for 

contamination effects, Not shown in figure N/A 

8 Sun Angle Sensors NRL Two orthogonal sun angle sensors. Not shown in figure N/A 

9 

CIGS 

NASA 

GRC/AECAble/ITN 

Blanket-level test of interconnected CIGS cells on stainless 

steel affixed to a Vectran gore sheet simulating UltraFlex 

array, ~20cm2

 (5 series-connected cells) 

~3.5* 

ATJM Emcore 3J commercial ATJM cells, includes monolithic bypass diode, 

26.6cm2

 (5) 27.1 10 

BTJ Emcore 3J commercial BTJ cells, 26.6cm2

 (5) 27.8 

NASA/CP—2005-213431 58
PROMISING RESULTS FROM THREE NASA SBIR 

SOLAR ARRAY TECHNOLOGY DEVELOPMENT PROGRAMS 

Mike Eskenazi, Steve White, Brian Spence, Mark Douglas, Mike Glick, Ariel Pavlick, and David Murphy 

ABLE Engineering, Goleta, CA 93117 

Mark O’Neill and A.J. McDanal 

ENTECH Inc., Keller, TX 76248 

Michael Piszczor 

NASA Glenn Research Center, Cleveland, OH 44135 

ABSTRACT 

Results from three NASA SBIR solar array technology programs are presented. The programs discussed are the 

1) Thin Film Photovoltaic UltraFlex Solar Array; 2) Low Cost/Mass Electrostatically Clean Solar Array (ESCA); 

and 3) Stretched Lens Array SquareRigger (SLASR). The purpose of the Thin Film UltraFlex (TFUF) Program is 

to mature and validate the use of advanced flexible thin film photovoltaics blankets as the electrical subsystem 

element within an UltraFlex solar array structural system. In this program operational prototype flexible array 

segments, using United Solar amorphous silicon cells, are being manufactured and tested for the flight qualified 

UltraFlex structure. In addition, large size (e.g. 10 kW GEO) TFUF wing systems are being designed and 

analyzed. Thermal cycle and electrical test and analysis results from the TFUF program are presented. The 

purpose of the second program entitled, “Low Cost/Mass Electrostatically Clean Solar Array (ESCA) System,” is 

to develop an Electrostatically Clean Solar Array meeting NASA’s design requirements and ready this technology 

for commercialization and use on the NASA MMS and GED missions. The ESCA designs developed use flight 

proven materials and processes to create a ESCA system that yields low cost, low mass, high reliability, high 

power density, and is adaptable to any cell type and coverglass thickness. All program objectives, which included 

developing specifications, creating ESCA concepts, concept analysis and trade studies, producing detailed 

designs of the most promising ESCA treatments, manufacturing ESCA demonstration panels, and LEO (2,000 

cycles) and GEO (1,350 cycles) thermal cycling testing of the down-selected designs were successfully achieved. 

The purpose of the third program entitled, “High Power Platform for the Stretched Lens Array,” is to develop an 

extremely lightweight, high efficiency, high power, high voltage, and low stowed volume solar array suitable for 

very high power (multi-kW to MW) applications. These objectives are achieved by combining two cutting edge 

technologies, the SquareRigger solar array structure and the Stretched Lens Array (SLA). The SLA 

SquareRigger solar array is termed SLASR. All program objectives, which included developing specifications, 

creating preliminary designs for a near-term SLASR, detailed structural, mass, power, and sizing analyses, 

fabrication and power testing of a functional flight-like SLASR solar blanket, were successfully achieved. 

1. NASA SBIR PHASE 2 THIN FILM PV ULTRAFLEX PROGRAM 

1.1 INTRODUCTION

As flexible thin film photovoltaic (FTFPV) cell technology is developed for space applications, integration into a 

viable solar array structure that optimizes the attributes of this cell technology is critical. An advanced version of 

the ABLE Engineering, Inc. (ABLE) UltraFlex solar array platform represents a near-term, low-risk approach to 

demonstrating outstanding array performance with the implementation of FTFPV technology. Recent studies 

indicate that an advanced UltraFlex solar array populated with 15% efficient thin film cells can achieve over 250 

W/kg BOL. 

NASA/CP—2005-213431 59
ABLE, in collaboration with the NASA Glenn Research Center (GRC), is currently executing a NASA Phase 2 

Small Business Innovation Research (SBIR) program to implement and validate the integration of FTFPV 

technology within an advanced UltraFlex solar array [1]. The primary objectives of the Phase 2 SBIR program are 

to mature and validate the use of advanced FTFPV blankets as the electrical subsystem element within an 

advanced UltraFlex solar array structural system. Program objectives to mature and validate the use of FTFPV 

blankets in the UltraFlex solar array platform are being met though the implementation of a hardware oriented 

development and test approach (at the coupon and engineering model level). 

As has been recognized in recent array system level studies [2], integration of a FTFPV electrical subsystem into 

an advanced UltraFlex solar array structural system allows realization of the full potential of FTFPV technologies 

to increase solar array specific power and significantly reduce cost. While the prior research focus has been 

aimed at the FTFPV device level in an effort to improve BOL/EOL efficiencies, reduce weight, and enhance space 

environmental survivability [2, 3], the development, implementation and validation of these technologies within a 

viable lightweight array system (including the “real world” effects on specific power performance from array wiring, 

structures and deployment mechanisms) has been only superficially addressed prior to this Thin-Film UltraFlex 

SBIR study. Validating the successful combination of a structurally/mechanically efficient advanced UltraFlex 

solar array platform with the unique properties of FTFPV technology is a key step in maturing these technologies 

and readying their space commercialization. 

During this NASA Phase 2 SBIR program preliminary design/analysis activities have been performed for a high 

performance thin film PV UltraFlex solar array system configured for a 10kW EOL GEO application. Feasible 

design solutions for high voltage operability, scalability, optical characteristics, high thermal performance, and 

radiation survivability have been identified. Coupon development and hardware tests have been performed on 

flight-like FTFPV gore/blanket assemblies to validate concept/design feasibly. Larger scale flight-like engineering 

model hardware has been built to further validate design manufacturing, packaging and deployment feasibility at 

the wing system level. 

1.2 TECHNOLOGY OVERVIEW

The UltraFlex solar array, shown in Figure 1, combines structural performance and the highest available specific 

power with a very low stowed volume and footprint [3]. The UltraFlex achieves its deployed strength and stiffness 

from lightweight composite radial spar members that allow tensioning of a flexible blanket populated with 

photovoltaic. This unique structural system allows the use of a flexible blanket without requiring massive 

secondary structure to deploy and tension the wing as is common in other flexible substrate solar array systems. 

UltraFlex is an excellent platform to demonstrate maximum specific power performance gains from high areal 

mass-efficient thin-film photovoltaics because of its superior areal structural mass efficiency that is up to 3.5X 

better than a conventional rigid planar array. 

Figure 1. UltraFlex Solar Array for the Mars 01-Lander Program

The UltraFlex is an accordion-fan-fold flexible solar array composed of ten interconnected isosceles triangular 

shaped lightweight substrates (gores). The fan-fold geometry sequences the gore blanket assemblies during 

deployment. Each interconnected triangular shaped substrate/gore unfolds to form a tensioned ten-sided polygon 

structure when fully deployed. Radial spar elements attached to each substrate are elastically deflected to form a 

tension-preloaded concave paraboloid, providing high-deployed strength and stiffness. A layout of the deployed 

UltraFlex depicting its major components is shown in Figure 2. 

NASA/CP—2005-213431 60
Figure 2. UltraFlex Wing Components

When stowed, the UltraFlex occupies a triangular space that is approximately 20% the stowed volume of 

conventional planar arrays, which enables spacecraft to maximize launch volume for critical payloads. The 

honeycomb-composite static and pivot panels sandwich and preload the folded solar array blanket under 

pressure in the stowed configuration. In a standard crystalline photovoltaic-based design, the UltraFlex wing’s 

solar cells are protected from damage in the launch environment by foam interleaved in the folded blanket 

assembly. The stowed and compressed wing assembly is held to the space vehicle at three releasable launch tie 

points. Figure 3 shows the stowed UltraFlex configuration. 

Figure 3. UltraFlex Stowed Configuration

1.3 FLEXIBLE THIN FILM PV (FTFPV) 

The primary suppliers and technologies under consideration during this program included United Solar Systems’ 

(USS) amorphous silicon (α-Si) FTFPV, and ITN Energy Systems’ CIGS FTFPV technology. At program 

commencement, due to limited funding, only the USS α-Si technology was selected for implementation and test. 

Additional funding was obtained later in the program, and ITN Energy Systems CIGS will be included in similar 

coupon-level and engineering model level development tests to be performed in early 2004. 

The baseline α-Si FTFPV technology utilized for this program was USS “Q” solar cell. The “Q” cells are produced 

on a 25 micron (0.001-inch) thick stainless steel substrate, have a Ag/ZnO back reflector, a top side Indium-TinOxide (ITO) coating, and are tuned to the 1AM0 solar spectrum. The top surface conductive grid consists of a 

series of longitudinally oriented small diameter compression-bonded copper wires. The wires are 

terminated/connected to a silver coated copper bus that is used for cell-to-cell series interconnection. The cells 

measure approximately 4-cm X 17.8-cm and have an active area of 64.30 cm2

. The measured mass of each 

solar cell assembly is approximately 2.37 g (areal density of 0.37 kg/m2). A picture of the a-Si USS solar cell 

assembly is shown in Figure 4. 

Gore

Spar (typ.)

Mass Simulator Gore (typ.)

Primary Launch Tie

Hub Launch Tie

Base Mechanism Pivot Panel

Static Panel

Gore

Spar (typ.)

Primary Launch Tie

Hub Launch Tie

Base Mechanism Pivot Panel

Static Panel

NASA/CP—2005-213431 61
Figure 4. USS a-Si FTFPV (“Q”-cell) 

The baseline copper indium gallium diselenide (CIGS) FTFPV technology to be investigated during this program 

is the ITN Energy Systems Inc. (ITN) space-application solar cell mounted on 1 mil. stainless-steel substrates. 

ITN Energy Systems has developed a technology that adapts the cell material obtained from Global Solar Energy 

Inc. (GSE) into a product capable of surviving in a space environment. This is accomplished by adding AM0 

optimized space grids, thermal coatings to reduce on-orbit operating temperature and electrostatic discharge 

layers (ESD) for operation in a plasma environment to the basic cell material obtained from GSE. Additionally, 

the cells provided by ITN for this program will include ITN’s bonded interconnect technology that will allow 

integrated solar arrays for space applications to be implemented. As previously mentioned, validation of the ITN 

CIGS technology is currently underway with results anticipated for early 2004. 

1.4 THIN FILM ULTRAFLEX HARDWARE DEVELOPMENT & TESTING

1.4.1 Structural Optimization 

During this program many of the standard UltraFlex array’s structural design features were weight-optimized to 

take advantage of the FTFPV’s significantly lighter weight and robustness versus classical crystalline PV 

technologies. Because FTFPV technologies are projected to be almost four times lighter than the state-of-the art 

crystalline multi-junction cells on a weight per unit area basis, and are not subject to cracking or damage from 

vibratory motion, the method of tightly preloading the folded gore stack between layers of foam does not appear 

to be necessary with this technology. This allows the UltraFlex static and pivot panels (nominally of honeycombcomposite construction) to be further weight-optimized for use as stowed protective covers and deployed 

structural members, without the considerable stiffness and strength required to apply a uniformly-distributed 

preload over the stowed stack area, which is necessary for crystalline PV applications. ABLE has developed a 

concept for a lightweight iso-grid static and pivot panel construction that allows the panel stiffness to be tailored 

specifically for a low mass stowed FTFPV blanket stack. This panel construction saves significant weight (20-

40%), especially when scaling UltraFlex to the larger sizes (>4m diameter) required when implementing lower 

efficiency FTFPV. 

An alternate approach to reduce the weight of the static and pivot panel structure on large FTFPV UltraFlex wing 

configurations is to increase the number of launch tie locations so that the panels do not have to be as stiff and 

strong to apply preload to the stowed blanket stack. This concept requires the implementation of a weightoptimized launch tie assembly. The baseline UltraFlex launch tie is based on the heritage ABLE launch release 

that uses high-output paraffin actuators to release a tensioned cable. This design has extensive flight heritage 

and works well for applications requiring three or less tie-down release points, but would add unnecessary weight 

if applied at multiple (>4) launch tie points. For a larger-diameter FTFPV UltraFlex application requiring more 

launch tie locations, ABLE has initiated development of a next-generation launch tie release that utilizes small 

lightweight discrete bolt-separation actuators at each tiedown location. During this Phase 2 SBIR program, such 

a weight optimized launch tie release was considered in the FTFPV UltraFlex array sizing trades, with results 

indicating they have the potential to significantly reduce mass versus the existing launch tie technology, as well as 

reducing the necessity for ultra-stiff (and massive) structural panels. 

Additionally, the radial spar elements can be structurally optimized for FTFPV applications. With lightweight 

FTFPV technology the spar structures do not need to support as much weight when deploying or in the fully 

deployed configuration (under 1-G ground-test conditions). The standard UltraFlex design utilizes a flat 

rectangular composite spar construction with the lay-up optimized to support deployment loads under 1-G with 

minimal external support. A FTFPV array, with significantly larger wing area required for the same power output, 

NASA/CP—2005-213431 62
will require an external ground support structure to allow deployment under earth gravity. A rectangular spar 

applied to such a large wing could be susceptible to torsional instability and could buckle even when supported 

under 1-G unless it becomes very large and heavy. The design of a large TFTPV wing required re-assessment of 

the spar’s design to provide the most weight-optimal cross section that can still be packaged in the stowed 

configuration. ABLE has investigated concepts for deployable “lenticular” spar sections that stow in a flat 

package but expand out when deployed to form a very lightweight, torsionally stiff beam section that is less 

susceptible to lateral/torsional instability. 

1.4.2 Interconnect Development and Testing 

Previous solar array system studies suggest that a monolithically interconnected FTFPV blanket assembly is 

needed to truly realize the cost and mass benefits potential of this technology [1]. Because monolithic 

interconnection of FTFPV is still in development this technology was not available to implement on this program. 

Instead, the FTFPV interconnection technologies developed during this program focused on classical welding and 

soldering approaches. 

Because the UltraFlex blanket/gore assemblies are stowed face-to-face it is desirable to employ an interconnect 

that exhibits very low profile. To satisfy this need two interconnect approaches were ultimately pursued, a foldedflexible type that provided large stress relief and an in-plane type. A representative folded flexible interconnect 

prototype mock-up was made and tested for flexibility in two directions. The interconnect consisted of a folded 

copper strip that was welded to the top and bottom of adjacent cells. The fold in the copper strip allows the 

interconnect to flex during expansion and contraction of the circuit during thermal cycling. The folded interconnect 

testing successfully demonstrated the durability and flexibility of the design concept. A picture of the mechanical 

fold interconnect mock-up is shown in Figure 5. A prototype in-plane interconnect was also made and evaluated 

by ABLE and USS. The in-plane interconnect consisted of a straight copper strip that was welded to the top and 

bottom of adjacent cells using multiple welds. The in-plane interconnect testing successfully demonstrated the 

durability of this design. The results of this interconnect development effort indicated that either in-plane or 

stress-relief interconnects were suitable for UltraFlex and should be evaluated at the flight-like coupon level using 

the α-Si FTFPV. 

Figure 5. Folded-Flexible Interconnect Detail (Left) and In-Plane Interconnect Detail (Right)

The ITN CIGS cells will be interconnected using an adhesive bonding interconnect technology that combines the 

strength of a mechanical bond (to carry the CTE mismatch induced stress) with the electrical conductivity of a 

conductive bond. This interconnect design will be tested by ITN and ABLE to demonstrate the thermal-cycle 

durability of the design. 

1.4.3 Circuit/Laydown Development & Test 

To maintain blanket design commonality with previously qualified crystalline PV UltraFlex systems the FTFPV 

blanket/gore approach is comprised of a mesh substrate to which the interconnected cells are bonded. The 

amount of cell-to-gore mesh bond area can be significantly reduced, because the lightweight FTFPV will have 

less of a propensity to de-bond under its own accelerated mass. 

NASA/CP—2005-213431 63
An ultra-lightweight method for cell-to-gore mesh attachment was developed that utilized Kapton “dots” bonded 

directly to the cell backside and the UltraFlex gore/substrate, as shown in Figure 6. This feature resulted in an 

80% savings in cell-to-gore adhesive mass over the crystalline UltraFlex design. This low-cost method also 

provides a significant reduction in laydown assembly time compared to the standard UltraFlex cell attachment 

method. Engineering development tests were performed on a prototype incorporating this laydown approach to 

demonstrate concept feasibility. A localized proof load using a small vacuum suction cup was applied to each 

bonded cell to determine acceptable bond strength. Test results indicate that no bond degradation occurred 

under the normal load application. 

Figure 6. Lightweight Cell-to-Gore Attachment Using Kapton Dots

To further validate the cell/circuit interconnect approach, and circuit laydown, a flight-like engineering mock-up 

was fabricated. The engineering circuits incorporated the folded flexible interconnect design as well as the 

baseline in-plane welded interconnect and were mounted to a tensioned UltraFlex blanket/gore segment. The 

hardware employed the following three circuits: Circuit-1 consisted of three 25-micron-thick dummy cells with 

straight 100-micron-thick copper strip soldered interconnects, Circuit-2 consisted of two 125-micron-thick cells 

with a soldered flexure interconnect and improved grid wire terminations, and Circuit-3 consisted of two 

125 micron-thick cells with a soldered flexure interconnect and improved (segmented) grid wire termination. The 

original design for the grid wire terminations consisted of a layer of Kapton tape above and below the grid wires, 

and this double-layer termination method exhibited significant peeling following limited thermal cycles. An 

improved approach was developed that utilized a single layer of Kapton tape compression-bonded over the grid 

wires and to the cell edge surface. Additionally, this grid wire termination tape was segmented into three shorter 

sections on one of the samples in an effort to reduce thermal stresses and distortion. A photograph of the 

development circuit hardware is shown in Figure 7. Thermal cycle testing of the engineering gore segment was 

performed to determine any degradation and validate grid wire termination, cell/substrate attachment, and 

interconnect integrity/performance. The hardware was thermal cycled for a total of 266 cycles between –175C to 

+100 C at a rate of 2.5 cycles/hr. Visual inspections and interconnect resistance measurements were made at 

55, 104, and 266 cycles. Test results indicated no visible degradation of the straight and flexible interconnect 

configurations, no change in interconnect resistance, and the cell/substrate attachment was unaffected based on 

pull test measurements. The improved single layer grid wire termination (segmented and un-segmented) had no 

visible degradation following 266 thermal cycles. As a result of this testing the improved, segmented approach 

has been baselined. 

` 

NASA/CP—2005-213431 64
Figure 7. Development Circuits (Front and Back Views)

1.4.4 Thin-Film UltraFlex Engineering Circuit (E-Board) Testing 

To further validate FTFPV interconnection and laydown onto UltraFlex, two flight-like engineering circuits 

(E-Boards) were fabricated. Each circuit consisted of a 6-cell series interconnected a-Si FTFPV assembly 

bonded to the UltraFlex gore/substrate. Cell thickness was 25 microns (0.001 inch). Each circuit utilized a 

different interconnect method; the baseline in-plane welded tab design and the folded flexure design. The 

engineering circuits were inspected and tested with ABLE’s Large Area Pulsed Solar Simulator (LAPSS), and 

then mounted on two separate sections of tensioned UltraFlex gore/blanket using the previously verified Kapton 

“dot” attachment method. A pre-test electrical LAPSS performance evaluation was performed for the completed 

assembly. Pictures of the small flight-like E-Boards are shown in Figure 8. 

Figure 8. E-Board Circuit on UltraFlex Gore/Substrate

The completed E-boards were then mounted within a thermal chamber and were subjected to 500 thermal cycles 

at temperatures from -175°C to +100°C. Periodic inspections and continuity monitoring during the thermal cycle 

tests indicated no apparent degradation/failures of either circuit. Final post-test inspection and LAPSS verification 

of the circuits were performed, showing both circuits in good condition with only minimal localized physical 

degradation. 

Post-thermal LAPSS testing of the two circuits demonstrated no measurable degradation in cell performance 

when compared to the pre-thermal LAPSS data. The observed local separation of the Kapton tape and the rear 

contact on the flexible copper interconnect circuits did not affect the circuit performance. Pull tests on the cell-togore/substrate bonds demonstrated that the cell attachment was unaffected by thermal cycling. 

NASA/CP—2005-213431 65
The following conclusions were made from the fabrication and testing of the two E-Board circuits: 

1) The circuit electrical performance was unchanged by the thermal cycle testing. 

2) The straight welded interconnect is generally a good match with the UltraFlex gore/substrate and local 

attachment of the cell. No interconnect failure resulted from the thermal expansion and contraction of the 

cells relative to the UltraFlex gore/substrate. The flexible interconnect is probably not required when the cells 

are attached to the UltraFlex gore/substrate. 

3) The number of welds used to attach the rear copper contact to the stainless substrate may need to be 

increased to prevent the observed separation of the contact from the stainless substrate. 

4) The separation of the Kapton tape from the top ohmic points to a required change in surface preparation, tape 

adhesive, tape type, or material configuration when the cell is required to operate at the tested thermal 

environment. 

5) The cell attachment was generally unaffected by the thermal environment. 

1.4.5 Face-to-Face Survivability 

An investigation into the surface durability of the α-Si FTFPV modules was performed. This is an area of 

particular concern for the UltraFlex application because in the nominal stowed configuration, FTFPV cells are in 

direct face-to-face contact. A number of tests were performed initially to assess the relative durability of the 

(unprotected) α-Si FTFPV surface, as shown in Table 1. The tested FTFPV cells did not have any anti-reflective 

or oxide coatings on their front surface. Initial test results indicate that an unprotected front surface of the α-Si 

FTFPV is highly susceptible to damage from cell-to-cell contact and handling. 

Table 1. Surface Durability Test Results of Un-coated FTFPV 

The effectiveness in improving FTFPV durability by employing various thickness SiO2 coatings applied to the PV 

front surface was investigated. A number of 4.7” x 14” terrestrial α-Si cells were obtained from USS for use in 

determining optimal coatings to enhance cell front surface durability. The cells were coated with SiO2 of thickness 

varying from 0.1 μm to 3 μm, as shown in Table 2. Additionally, a 0.06 thick ITO layer was applied to the front 

surface of the cells. The reason for the ITO layer was to eliminate the potential for charge buildup on the SiO2

surface on orbit. Subsequent discharge from the cell surface may damage the circuit. 

Table 2. Matrix of ITO Coating Thickness Applied to Durability Test Cells 

Designation SiO2 Thickness (μm) ITO Thickness (μm) 

A 0.1 0.06 

B 1.0 0.06 

C 3.0 0.06 

Test Description Test Result

Two cells were placed face-to-face under 0.1 psi pressure No cell damage

Rounded edge of paperclip placed in contact with cell surface 

and lightly translated across cell

Cell completely 

shunted

Two cells were placed face-to-face. Light pressure was applied 

to top cell and top cell moved relative to bottom cell 

(displacement ~ .05 in., duration ~ 3 sec).

Cell completely 

shunted

Two cells were placed face-to-face. Foam was placed over top 

cell and weights applied to exert 0.1 psi pressure. The assembly 

was repeatedly moved on granite table (displacement ~ .05 in., 

duration ~ 3 sec.)

Cell completely 

shunted

Test Description Test Result

Two cells were placed face-to-face under 0.1 psi pressure No cell damage

Rounded edge of paperclip placed in contact with cell surface 

and lightly translated across cell

Cell completely 

shunted

Two cells were placed face-to-face. Light pressure was applied 

to top cell and top cell moved relative to bottom cell 

(displacement ~ .05 in., duration ~ 3 sec).

Cell completely 

shunted

Two cells were placed face-to-face. Foam was placed over top 

cell and weights applied to exert 0.1 psi pressure. The assembly 

was repeatedly moved on granite table (displacement ~ .05 in., 

duration ~ 3 sec.)

Cell completely 

shunted

Test Description Test Result

Two cells were placed face-to-face under 0.1 psi pressure No cell damage

Rounded edge of paperclip placed in contact with cell surface 

and lightly translated across cell

Cell completely 

shunted

Two cells were placed face-to-face. Light pressure was applied 

to top cell and top cell moved relative to bottom cell 

(displacement ~ .05 in., duration ~ 3 sec).

Cell completely 

shunted

Two cells were placed face-to-face. Foam was placed over top 

cell and weights applied to exert 0.1 psi pressure. The assembly 

was repeatedly moved on granite table (displacement ~ .05 in., 

duration ~ 3 sec.)

Cell completely 

shunted

NASA/CP—2005-213431 66
After coating, the electrical performance of the test cells was measured and the durability of the cells was 

evaluated. One cell was placed on a smooth granite surface with the PV side facing upward. The second cell 

was positioned so that the PV surface of the second cell contacted the PV surface of the first cell. A foamcovered plate was placed over the assembly and weights were applied to the aluminum plate in order to achieve 

a particular contact pressure. The top cell was then moved relative to the bottom cell with a displacement of 

approximately 0.05 inches along the direction of the gridlines for a period of 20 seconds. The process was 

repeated with the displacement direction perpendicular to the gridlines. The electrical performance of the test cell 

was then determined. The contact pressure was then increased and the durability test repeated. Three contact 

pressures (0.1, 0.2 and 0.3 psi) were studied. Two different contact surfaces (PV to PV and PV to ohmic) were 

evaluated. The durability test matrix is shown in Table 3. 

Table 3. Durability Test Matrix 

Sample No. Contacting 

Surfaces 

SiO2 

Thickness 

(μm) 

Pressure 

(psi) 

Duration 

(sec) 

1000A1 PV to PV 0.1 0.1, 0.2, 0.3 20/20,20/20,20/20* 

1000A2 PV to PV 0.1 0.1, 0.2, 0.3 20/20,20/20,20/20* 

1000A3 PV to Ohmic 0.1 0.1, 0.2, 0.3 20/20,20/20,20/20* 

1000A4 PV to Ohmic 0.1 0.1, 0.2, 0.3 20/20,20/20,20/20* 

10000B1 PV to PV 1.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

10000B2 PV to PV 1.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

10000B3 PV to Ohmic 1.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

10000B4 PV to Ohmic 1.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

30000C1 PV to PV 3.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

30000C2 PV to PV 3.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

30000C3 PV to Ohmic 3.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

30000C4 PV to Ohmic 3.0 0.1, 0.2, 0.3 20/20,20/20,20/20* 

* Duration: Test cell in contact with adjacent cell under specified pressure. Test cell 

moved relative to static adjacent cell (20 seconds along direction of gridlines, 

20 seconds across direction of gridlines). 

The post-test LAPSS data showed that the cells with the 1.0 μm and 3.0 μm SiO2 coatings had no measurable 

degradation in performance following the durability testing at the three contact pressure levels. The cells with the 

thinnest SiO2 coating (0.1 μm) showed a significant drop in peak power (10 -11 %). The results of these tests 

indicated that a 10000A-thick SiO2 coating to the front surface prevented the cell shunting failure mode caused by 

direct face-to-face contact as would be seen in a Thin Film UltraFlex application. Even with significant abrasion 

pressures (well above those expected in an UltraFlex application), damage was limited to minimal power 

degradation (< 2% reduction in PMAX) in the SiO2-coated cells. Further tests, including vibration, are planned as 

part of this study to more thoroughly validate the coatings. 

In addition to the protective SiO2 coating, several interleave options were evaluated in order to protect the cells 

within the stowed UltraFlex stack. The first interleave option studied was a layer of Tedlar film that rolls up via a 

lightweight constant force spring as the stowed array stack separates. The roll of interleave film would be located 

adjacent to the spar when the array was deployed. An engineering model was produced and used to verify the 

functionality of this interleave concept. This interleave method does add significant complexity to the array 

deployment. 

The second interleave option studied was a static layer of open-weave gore/substrate that covers the front side of 

the cells. Although this concept is relatively easy to implement, the gore/substrate produces a minimum power 

loss of approximately 5% over an uncovered cell. The thermal effects of the gore/substrate on array operating 

temperature need to be evaluated. The SiO2 coating is the best cell protective option investigated to date. 

1.4.6 Large Engineering Gore Circuit 

A flight-like engineering gore/blanket circuit of the α-Si FTFPV technology was fabricated that incorporated many 

of the lessons learned from the previous development/validation tests. A picture of this circuit is shown in 

NASA/CP—2005-213431 67
Figure 9. The circuit was laid-down to the UltraFlex substrate/gore and then will be integrated into ABLE’s 1st

generation IR&D qualification wing, as shown in Figure 10. Wing system level tests will then be performed to 

further validate the implementation of the α-Si FTFPV technology within the UltraFlex platform. 

Figure 9. Engineering Gore/Blanket Circuit 

Figure 10. EM Gore Circuit Integrated into Prototype UltraFlex Substrate Gore 

The 18-cell USS circuit mounted to the EM gore was coated with 1 μm SiO2 and 60 nm of ITO. LAPSS tests were 

performed on the circuit both before and after the coating operation. The preliminary results of the LAPSS tests 

performed on the circuit after coating indicated a 34% power reduction from the pre-coated circuit, as shown in 

Figure 11. 

Mass Equivalent 

Dummy FTFPV

A-Si 

Engineering 

Circuit

Gore/Blanket Assembly

Handling Frame

a -Si FTFPV Series 

Interconnected Circuit

EM Gore/Blanket Assembly in 

Thermal Chamber

Mass Equivalent 

Dummy FTFPV

A-Si 

Engineering 

Circuit

Gore/Blanket Assembly

Handling Frame

a -Si FTFPV Series 

Interconnected Circuit

EM Gore/Blanket Assembly in 

Thermal Chamber

NASA/CP—2005-213431 68
Figure 11. Pre and Post-Coating LAPSS Test Results for 18-cell EM Circuit 

Because of the observed degradation in power as a result of the coating, an investigation was initiated to evaluate 

the cause of the power reduction. The investigation indicated that the shorting between the conductive ITO 

coating and the cell front photovoltaic surface was occurring through pinholes in the SiO2. Currently, United Solar 

is researching suitable front-surface space-rated conductive coatings and application methods to eliminate this 

problem. Cells to be subsequently tested will be coated with 1 μm SiO2 only (no ITO). 

1.4.7 MISSE Flight Experiment Coupons 

Flight experiment coupons of the thin film UltraFlex designs were produced for the Materials of the International 

Space Station Experiment–5 program (MISSE-5). The purpose of the Materials International Space Station 

Experiment (MISSE) is to characterize the performance of new prospective aerospace materials when subjected 

to the effects of the space environment. Two FTFPV UltraFlex flight experiment coupons were produced, each 

with a different FTFPV technology. The first experiment employed USS α-Si FTFPV, while the second 

experiment used ITN Energy’s CIGS. Both experiments were laid down to a dedicated UltraFlex gore/blanket 

assembly. The USS α-Si circuit consisted of two AM0 cells with Al/ZnO back reflector on 1-mil stainless steel 

substrate with single layer side “B” Kapton tape termination. A redundantly welded copper interconnect was used 

to connect the two cells. Cell to cell interconnection, I-V performance, and operating temperature will be 

assessed during this experiment. Pictures of the MISSE-5 flight experiment coupons are shown in Figure 12. 

Figure 12. MISSE-5 FTFPV UltraFlex Flight Experiment Coupons 

(Uncoated Circuit) (Coated Circuit)

NASA/CP—2005-213431 69
1.5 ACKNOWLEDGMENT

This work was administered though the NASA Glenn Research Center and funded by the NASA SBIR Phase 2 

program under contract #NAS3-02078. 

1.6 THIN FILM ULTRAFLEX PROGRAM SUMMARY

The initial results of this program indicate that advanced FTFPV can be successfully integrated to a lightweight 

optimize UltraFlex solar array to produce a system that exhibits very high specific power. This program has 

provided an important foundation towards validating through hardware development cell-to-cell interconnection, 

electrical circuitry, circuit laydown and blanket/gore construction, survivability (thermal life cycling and face-to-face 

contact) and wing/array system level aspects. Future work will involve a series of wing system-level integration 

and validation activities that will further demonstrate concept feasibility, and more detailed hardware approach to 

assess face-to-face survivability of the FTFPV under a vibration environment. Additionally, an advanced UltraFlex 

incorporating the alternative ITN CIGS thin film technology will be developed and validated 

2. NASA SBIR PHASE 1 LOW COST/MASS ELECTROSTATICALLY CLEAN SOLAR ARRAY (ESCA) 

PROGRAM 

2.1 INTRODUCTION

Electrostatic interaction of solar arrays in a space plasma environment is of great concern as differential charging 

across panel portions can lead to high electric fields, charge loss to plasma, and arcing between critical 

components, which may disrupt scientific measurements of fields or particles [4]. The objective of this program is 

to develop a low cost and low mass Electrostatically Clean Solar Array (ESCA) which will not interfere with 

sensitive spacecraft instruments/measurements and ready this technology for commercialization and use on 

NASA and commercial programs such as the NASA MMS and GED missions. Previous developed ESCA 

designs are shown in references [5, 6, & 7]. 

Solar array electrostatic cleanliness is achieved when circuitry is sealed to prevent exchange of charge with the 

plasma, or unintentionally between cell/string components, and all exterior surfaces are sufficiently conductive to 

prevent charge differentials on the panels. In meeting this objective, two innovative ESCA solutions have 

successfully been developed and tested. These are shown in Figure 13 (Zone 1 and Zone 2). The first design 

uses conductive adhesive grout, which is grounded to the panel core, to electrically connect the ITO coated 

coverglass on the front of the cells and create an external ground plane. The second design uses conductive 

tape grounded to the panel core to seal circuitry, electrically connect the ITO coverglass, and create an external 

ground plane. In both designs, the front-side solar cell circuitry is fully encapsulated in non-conductive adhesive 

for electrical isolation prior to the application of the conductive exterior ESCA treatment. On the panel backside 

layers of non-conductive natural Kapton film and conductive polymer films are used for circuit sealing and 

conductivity. The ESCA designs use flight proven materials and processes to create ESCA systems that yields 

low cost, low mass, high reliability, high power density, and adaptability to any cell type and coverglass thickness. 

2.2 REQUIREMENTS

The NASA requirements and goals specified in Table 4 form the basis of the ESCA design solutions. All 

requirements and goals shown in Table 4 are met by ABLE’s ESCA designs. 

2.3 ANALYSIS

2.3.1 Mass Analysis 

Mass calculation summaries for the conductive grout and conductive tape ESCA treatments are presented in 

Table 5. Mass was calculated on a per Cell-Interconnected-Cover (CIC) basis so that the relative magnitude of 

the mass could be easily determined and so that the calculation would be independent of panel type. The results 

show that the mass added by either ESCA treatment compared to the CIC's is < 8% and compared to a typical 

solar panel assembly < 3%, on a per unit area basis. However, the mass of an ESCA solar array also increases 

by approximately another 10% on a kg/W basis because of the increased panel area required due to the ESCA 

treatment reducing the total illumination area. The mass of the conductive tape treatment was found to be 

NASA/CP—2005-213431 70
significantly lighter than the conductive grout treatment. Both treatments meet the NASA specification for mass 

increase (<20%). 

Conductive Adhesive A 

Conductive 

Adhesive B 

Ground 

Wire 

Conductive Tape B 

Conductive Tape A 

Ground 

Wire 

ITO Coated CMX (2kΩ/sq) 

72mm x 43mm x 200μm

16 Pieces total 

Coverglass is bonded with CV-2568 silicone to solar panel

(M55J/CE graphite facesheets with 1.8cm Al honeycomb core) 

Coverglass is bonded with CV-2568 silicone to solar panel

(M55J/CE graphite facesheets with 1.8cm Al honeycomb core) 

Conductive Adhesive A 

Conductive 

Adhesive B 

Ground 

Wire 

Conductive Tape B 

Conductive Tape A 

Ground 

Wire 

ITO Coated CMX (2kΩ/sq) 

72mm x 43mm x 200μm

16 Pieces total 

Coverglass is bonded with CV-2568 silicone to solar panel

(M55J/CE graphite facesheets with 1.8cm Al honeycomb core) 

Coverglass is bonded with CV-2568 silicone to solar panel

(M55J/CE graphite facesheets with 1.8cm Al honeycomb core) 

Figure 13. Phase I ESCA demonstration panel that was successfully thermal cycled 2,000 times between -

90°C to 100°C (LEO) and 1,350 times from –175°C to 75°C (GEO) 

Table 4. NASA ESCA System Requirements/Goals 

# NASA Requirement/Goal Description 

1 Regardless of size, there shall be no more than 100 millivolts between any two points on a given panel’s 

surface, even in the presence of a current of 0.1 microampere per square centimeter caused by the ambient 

plasma 

2 The ESCA system shall not induce an electric potential greater than 0.1 volts, at a distance of 0.1 cm or 

greater from the panel surface, compared with a perfectly conductive surface grounded to the spacecraft 

structure 

3 There shall be no exposed insulator on the panel’s front or rear surfaces to the charged particle environment 

4 The ESCA system shall be capable of having its quasi-equipotential surface connected to the spacecraft 

through a connector 

5 The proposed method to achieve electrostatic cleanliness shall be compatible with cover thicknesses from 

.004 inches to .060 inches 

6 The proposed method to achieve electrostatic cleanliness shall be compatible with any type of solar cell 

typically used on spacecraft solar arrays including single crystal silicon, GaAs/Ge, and multi-junction solar 

cells 

7 The ESCA design shall only use materials that exhibit no more than 1.0- percent total mass loss and no more 

than 0.1 percent-collected volatile condensable materials from outgassing in a vacuum 

8 The methods to achieve electrostatic cleanliness shall be compatible with typical spacecraft solar array 

materials and assembly processes 

9 The ESCA design shall use no magnetic parts 

10 The ESCA design shall lend itself to easy cell removal and replacement 

11 The method to achieve electrostatic cleanliness shall not increase array cost ($/W) by more than 20% 

12 The method to achieve electrostatic cleanliness shall not increase array mass (W/kg) by more than 20% 

13 The method to achieve electrostatic cleanliness shall not significantly decrease the reliability of the solar array 

14 The method to achieve electrostatic cleanliness shall be applicable to multi-AU missions 

Zone 1

Zone 2

Zone 3

NASA/CP—2005-213431 71
Table 5. Mass Summary for ESCA Treatments

2.3.2 Power Analysis 

ESCA treatments reduce power by blocking sunlight that would normally hit the edges of the cell and because 

ITO coated coverglass has lower transmittance compared to AR coated coverglass. The reduction in 

transmittance for ITO coated glass compared to standard MgF2 AR coated glass for the GaInP2 (top) junction of a 

triple-junction solar cell calculated from the manufacturer’s data is shown in Figure 14. The calculated power 

reduction for the conductive grout and conductive tape ESCA treatments are presented in Table 6. The total 

power reduction for both treatments is about 8%, of which 5% is due to light blockage and 3% due to reduced 

transmittance through the ITO. The power reduction can be reduced slightly by using oversize cover glass so that 

less or none of the cell is blocked by the conductive grout or tape. 

ITO coated glass vs. AR coated glass

0.92

0.93

0.94

0.95

0.96

0.97

0.98

0 5000 10000 15000 20000 25000

Sheet Resistance ohms/sq

Relative Transmittance

Figure 14. Relative Reduction in Solar Cell Current for ITO Coated Glass

Mass Calculation Summary for Conductive Grout ESCA Treatment

Cell Length (cm) 3.76

Cell Width (cm) 7.61

Cell-to-Cell Gap (cm) 0.0762

Grout Overlap on Cell (cm) 0.0508

Grout Thickness (cm) 0.0381

Grout Volume per CIC (cc) 0.077

Grout Density (g/cc) 3.4

Grout Mass per CIC (g) 0.262

Measured CIC Mass, 100μm cover (g) 3.34

Grout as a percentage of CIC mass 7.8%

Typical CIC mass fraction of solar panel 36.0%

Deduced grout mass fraction of solar panel 2.8%

Mass Calculation Summary for Aluminized Tape ESCA Treatment

Cell Length (cm) 3.76

Cell Width (cm) 7.61

Cell-to-Cell Gap (cm) 0.0762

Tape Overlap on Cell (cm) 0.0508

Tape Thickness (cm) 0.00635

Tape Volume per CIC (cc) 0.013

Tape Average Density (g/cc) 1.4

Tape Mass per CIC (g) 0.018

Conductive Adhesive Corner Dots per CIC (g) 0.055

Total ESCA mass per CIC 0.073

Measured CIC Mass, 100μm cover (g) 3.34

ESCA as a percentage of CIC mass 2.2%

Typical CIC mass fraction of solar panel 36.0%

Deduced ESCA mass fraction of solar panel 0.8%

NASA/CP—2005-213431 72
Table 6. Power Reduction for ESCA Treatments 

2.3.3 Cost Analysis 

The material cost in both the conductive grout and conductive tape ESCA treatments are relatively low. Most of 

the cost associated with both methods is due to labor. The estimated recurring cost for the ESCA treatments are 

presented in Table 7. The treatments are relatively inexpensive comprising only about 5% of the solar panel cost 

based on an area basis. However, more panel area (~8%) is required to obtain a given power level compared to 

an untreated panel. Thus, in terms of $/W the increase in cost is approximately (1.05)(1.08) = 1.134. Both ESCA 

treatments meet the NASA requirement for cost (<20% cost increase). These cost estimates do not include any 

non-recurring engineering, which would be required for a 1st of a kind array. 

Table 7. Estimated Recurring Cost for ESCA Treatments 

CV2-2646 $/Gram 5

Installed Grams/CIC 0.26

Waste Factor 1

Waste (g) 0.26

Total Grams/CIC 0.52

CV2-2646 Cost/CIC $2.62

Time Per Task (minutes)

Inspection 3

Masking 2

Grouting 3

Clean-up 3

Inspection 2

Rework 2

Misc 5

Total Labor (min) 20

Assumed Labor Rate $/min $1.00

Labor Cost per CIC $20.00

Material Cost per CIC $2.62

ESCA Treatment Total $22.62

Typical 3J CIC Cost $450.00

ESCA Cost/CIC Cost 5.0%

ESCA Cost Summary

Conductive Grout Treatment

Materials

Labor

Aluminized Kapton ($/cm2) 0.78

Installed Area (cm2

/CIC) 2.02

Waste Factor 1

Waste (cm2/CIC ) 2.02

Total cm2

/CIC 4.04

CV2-2646 Cost/CIC $3.15

Time Per Task (minutes)

Inspection 1

Taping 2

Adhsive Dots 2

Inspection 2

Rework 2

Misc 5

Total Labor (min) 14

Assumed Labor Rate $/min $1.00

Labor Cost per CIC $14.00

Material Cost per CIC $3.15

ESCA Treatment Total $17.15

Typical 3J CIC Cost $450.00

ESCA Cost/CIC Cost 3.8%

ESCA Cost Summary

Aluminized Tape Treatment

Materials

Labor

Cell Length (cm) 3.76

Cell Width (cm) 7.61

Cell Area (cm2

) 28.6

Overlap Along Edges (cm) 0.051

Covererd Area (cm2

) 1.14

Average Blocked Light % 4.0%

Cell-to-Cell Variation in Blocked Light % 1.0%

Worse Case Blocked Light % 5.0%

ITO Coverglass Transmittance Loss (%) 3.0%

Total 8.0%

ESCA Power Adjustment Factors Summary

Conductive Grout Treatment

Cell Length (cm) 3.76

Cell Width (cm) 7.61

Cell Area (cm2

) 28.6

Overlap Along Edges (cm) 0.051

Covererd Area Edges (cm2

) 1.14

Covered Area Dots (cm2) 0.14

Average Blocked Light % 4.5%

Cell-to-Cell Variation in Blocked Light % 1.0%

Worse Case Blocked Light % 5.5%

ITO Coverglass Transmittance Loss (%) 3.0%

Total 8.5%

ESCA Power Adjustment Factors Summary

Aluminized Kapton Treatment

NASA/CP—2005-213431 73
2.4 DEMONSTRATION HARDWARE AND TEST RESULTS

Based on the results of pathfinder coupon tests, the Phase 1 ESCA engineering model, shown in Figure 15, was 

fabricated using the most promising conductive adhesive and conductive tape treatments. T he engineering 

model contained sixteen pieces of ITO coated coverglass (72mm x 43mm x 200microns). The engineering 

model was thermal cycled 2,000 times in a LEO environment and 1,350 times in a GEO environment. Visual 

inspections and resistance measurements were made prior to and after thermal cycling. The visual inspection 

results and resistance measurements are shown in Figures 16. The conductive adhesive (zone 1) and 

conductive tape B (zone 2) passed both the mechanical and electrical test criteria. The conductive tape A 

treatment (zone 3) failed both mechanically and electrically. 

	







Tape A 

lifting and peeling

Adhesive B 

cracking 

Adhesive A 

crack along tape 

edge 

2 cracks at 

corner in 

adhesive A

2 pinholes in Adhesive A

Adhesive A 

cracks along 

tape edges 

1

6

5

4

3 7

8

9

2

13

10

11

12 16

15

14

Figure 15. Visual Inspection Results for Engineering Model 

After 2,000 LEO Cycles and 1,350 GEO Cylces Between –175°C and 75°C

Pre Test Post Test

1 Conductive Adhesive 1.0 1.1 Pass Mechanical and Electrical

2 Conductive Adhesive 1.5 1.6 Pass Mechanical and Electrical

3 Conductive Adhesive 1.5 1.2 Pass Mechanical and Electrical

4 Conductive Adhesive 1.5 1.3 Pass Mechanical and Electrical

5 Conductive Adhesive 1.0 1.0 Pass Mechanical and Electrical

6 Conductive Adhesive 2.2 1.5 Pass Mechanical and Electrical

7 Conductive Adhesive 3.0 1.2 Pass Mechanical and Electrical

8 Conductive Adhesive 1.4 1.1 Pass Mechanical and Electrical

9 Conductive Tape B 1.6 1.8 Pass Mechanical and Electrical

10 Conductive Tape B 1.6 2.7 Pass Mechanical and Electrical

13 Conductive Tape B 2.3 2.4 Pass Mechanical and Electrical

14 Conductive Tape B 6.3 8.5 Pass Mechanical and Electrical

11 Conductive Tape A 2.0 1.8 Fail, Tape Peeling, Adhesive Cracking

12 Conductive Tape A 2.0 3.3 Fail, Tape Peeling, Adhesive Cracking

15 Conductive Tape A 2.6 3.8 Fail, Tape Peeling, Adhesive Cracking

16 Conductive Tape A 6.0 OVLD Fail, Tape Peeling, Adhesive Cracking

*Less than 30kohm required to meet ESCA requirement of 0.1V maximum

2,000 LEO Thermal Cycle from -90°C to +100°C

1,350 GEO Thermal Cycles from -175°C to +75°C

Glass Treatment Comment

	
	
 

Figure 16. Coverglass Ground Resistance Results for Engineering Model

After 2,000 LEO Cycles and 1,350 GEO GEO Cylces. 

Zone 1

Zone 2

Zone 3

NASA/CP—2005-213431 74
2.5 ELECTRICALLY ACTIVE ESCA PANEL

An electrically active engineering model of the conductive tape ESCA treatment containing nine TECSTAR TEC1 

3J solar cells (Donated by EMCORE) wired in series was fabricated and is shown in Figure 17. The power of the 

electrically active engineering model was measured before and after application of the ESCA treatments to 

measure the effect on power. The results, which are shown in Figure 18 show a decrease in current (Isc) of 4.6% 

and a decrease in power (Pmax) of 6.3% which is in good agreement with the predicted light blockage of 5.5% 

calculated in Table 6. 

Figure 17. Electrically Active Prototype Before and After ESCA Treatment

ESCA Test Panel Before and After Treatment

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0 5 10 15 20 25

Voltage (V)

Current (A)

Pre ESCA Post ESCA Ratio (Post/Pre )

Isc (A) 0.371 0.354 0.954

Pmax (W) 6.352 5.954 0.937

ESCA Test Panel Before and After Treatment

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

0 5 10 15 20 25

Voltage (V)

Current (A)

Pre ESCA Post ESCA Ratio (Post/Pre )

Isc (A) 0.371 0.354 0.954

Pmax (W) 6.352 5.954 0.937

Figure 18. LAPSS Results for Active ESCA Test Panel

2.6 ACKNOWLEDGMENT

This work was administered though the NASA Goddard Space Flight Center and funded by the NASA SBIR 

Phase 1 program under contract #NAS5-03030. 

2.7 ESCA SUMMARY

Electrostatically Clean Solar Array (ESCA) technologies meeting NASA’s design requirements have been 

developed in preparation for commercialization and use on the NASA MMS and GED missions. The ESCA 

designs developed use flight proven materials and processes to create ESCA systems that yields low cost, low 

mass, high reliability, high power density, and is adaptable to any cell type and coverglass thickness. All Phase 1 

NASA/CP—2005-213431 75
objectives, which included developing specifications, creating ESCA concepts, concept analysis and trade 

studies, producing detailed designs of the most promising ESCA treatments, manufacturing ESCA demonstration 

panels and LEO (2,000 cycles) and GEO (1,350 cycles) thermal cycling testing of the down-selected designs 

were successfully achieved. It is hoped that these exceptional Phase 1 results and the high payoff of this 

technology to NASA GSFC’s planned MMS and GED missions result in a Phase 2 continuation of this program. 

3. NASA PHASE 1 SBIR HIGH POWER PLATFORM FOR THE STRETCHED LENS ARRAY 

3.1 INTRODUCTION 

The purpose of this program is to develop an extremely lightweight, high efficiency, high power, high voltage, and 

low stowed volume solar array suitable for very high power (multi-kW to MW) applications. This is achieved by 

integrating a Stretched Lens Array (SLA) blanket into the SquareRigger (SR) solar array structure. The resulting 

solar array is termed Stretched Lens Array SquareRigger (SLASR). 

SLASR uses SquareRigger, which is a well demonstrated structural approach for deploying extremely large 

photovoltaic blankets in terms of low mass, low stowed volume, and high stiffness and the SLA Fresnel 

concentrator, which is believed to be the best-demonstrated photovoltaic system in terms of the desirable 

combination of low cost, high efficiency (30%), ultra-low mass (>500 W/kg blanket), and high voltage (>1000V). 

The proposed SLA concentrator elements are highly optimized versions of the extremely successful Deep Space 

1 SCARLET solar array lenses [8]. An electrically active, flight-like, SLASR demonstration blanket was 

successfully fabricated and tested during Phase 1 and is shown in Figure 19. The excellent optical alignment 

achieved by the prototype SLASR blanket is shown in Figure 19 where the focus of natural sunlight on to the solar 

cells is demonstrated by the gossamer structure. 

The predicted near-term (<5 years) BOL performance for a 100 kW SLA SquareRigger array system is a specific 

power of 320 W/kg, a power per unit area of 310 W/m2

, and a stowed power per unit volume of 70 kW/m3 using 

the existing designs and demonstrated triple-junction solar cell technology. These characteristics of the SLA 

SquareRigger solar array, which are based on detailed models, make it an ideal candidate for meeting civil, 

military and commercial future power generation needs. 

Figure 19. SLASR Demonstration Blanket Built During Phase 1 

(Precise Optical Alignment of Prototype SLASR Demonstrated in Sunlight)

NASA/CP—2005-213431 76
3.2 BACKGROUND

The SquareRigger solar array structure (Figure 20) was developed by ABLE for the Air Force Research 

Laboratory (AFRL) to deploy and structurally support very large thin film solar array blankets for future high power 

Air Force missions. The key features of SquareRigger (using a hypothetical 0.2 kg/m2 thin film photovoltaic 

blanket made from 15% efficient CIGS) are very high power capability, low cost ($/W), low mass (>200 W/kg), 

and compact stowage for launch (>40 kW/m3

) [1]. These highly desirable characteristics may also enable 

numerous future NASA missions including Space Solar Power (SSP), high-power deep space SEP probes, Solar 

Electric Propulsion (SEP) tugs and large earth-orbiting science satellites. 

While the SquareRigger structure development has progressed rapidly, with all elements of 2nd generation 

hardware recently demonstrated at full scale as shown in Figure 21, the high efficiency (>15% AM0 & 0.2 kg/m2

) 

thin film photovoltaic blanket needed to complete this extremely lightweight, low cost, and high power solar array 

is currently not available. Additionally, there are many experts in the photovoltaic community who are concerned 

that the timeframe to develop a space-qualified, high efficiency, thin film blanket may be much longer than 

currently projected and that the attained performance will be lower than anticipated and the cost higher than 

expected. Thus, it is very desirable to develop alternative low cost and low mass photovoltaic blanket 

technologies for SquareRigger in case thin film photovoltaics are unavailable when needed. It is also highly 

desirable to develop a much more efficient photovoltaic system compared to thin films to greatly reduce the 

deployed area and the launch volume of the SquareRigger array. The Stretched Lens Array (SLA) blanket, under 

development for SquareRigger in this program, fits these requirements perfectly. 

... structure deploys synchronously under motor control...

...once structure is latched, motors then work to raise photovoltaic blankets.

Wing tiedowns release to initiate deployment...

Figure 20. SquareRigger Structure and Blanket Deployment

Stowed

Array

NASA/CP—2005-213431 77
Figure 21. Full-Scale SquareRigger Bay after Successful Ground Deployment

The Stretched Lens Array (SLA) [9] is a low mass/low cost high efficiency photovoltaic system that uses 

gossamer line-focus Fresnel lenses to focus sunlight at approximately 8.5X concentration on to spaced rows of 

high performance crystalline solar cells. The current embodiment of SLA on optimized rigid panels is well suited 

for lower power level spacecraft, offering a robust 180 W/kg, but not as well suited as SLA SquareRigger for high 

power applications (>15 kW wings). This is because SquareRigger was developed to specifically to overcome 

scale-up problems inherent in rigid panel platforms and provide lower mass and stowed volume. 

The SquareRigger platform can accommodate nearly any photovoltaic blanket technology. The AFRL is leading 

the development of thin films for space applications and the SquareRigger is a leading structural platform 

candidate. ABLE recently completed an AFRL funded program developing a thin film SquareRigger solar array 

system and also a crystalline (standard multi-junction cell) blanket as a backup technology to thin films, as this 

offers lower risk in the near term but much higher cost and mass. Alternatively, SLA technology adapted to 

SquareRigger simultaneously provides small deployed wing area and stowed volume (due to the use of high 

efficiency multi-junction cells), and low cost and low mass (enabled by a 70% to 80% reduction in cell area) 

approaching/surpassing the best thin film projections. 

Typical components of a SLA blanket, which can directly replace the thin film solar cell blanket in SquareRigger, 

are shown in Figure 22. The early version of the SLA photovoltaic blanket, also shown in Figure 22, was 

developed as part of the SSP program for use in a lower performing, earlier generation, “Space Station” style 

blanket array platform known as Aurora. Though built in 2000, this fully functional demonstration blanket was 

measured by NASA Glenn Research Center to have an output of greater than 375 W/m2

 at 20°C equating to a 

lens efficiency of 92%, a cell efficiency of 30.0%, and over 400 W/kg [4]. The SLA blanket developed for 

SquareRigger is an optimized version of this early concept. 

Low mass 

composite 

Node (corner) 

mechanism 

Spar joint 

and latch Ports for blanket 

deployment/tensioning 

cables 

This SquareRigger bay 

is the repeating 

element of the 

SquareRigger array 

NASA/CP—2005-213431 78
Figure 22. Stretched Lens SquareRigger (SLASR) Array

3.3 SLASR SYSTEM REQUIREMENTS AND GOALS

At the start of the Phase 1 program requirements/goals were developed for SLASR for both the near-term (2008) 

and mid-term (2013). These requirements, which are shown in Table 8, specify the performance metrics for the 

system (e.g. W/kg, W/m2, $/W etc.) and also define the launch and on-orbit environmental durability. These 

performance metrics were aggressively set to vastly improve the state of art in cost, mass, and stowed power 

density. The environmental durability requirements are representative of typical space solar array industry 

requirements and are based on ABLE’s broad experience in manufacturing solar arrays for commercial, NASA 

and military missions. These requirements were reviewed by NASA personnel and modified as required based 

on NASA’s input. All detailed design, analysis and testing performed during Phase 1 support the attainment (or 

very near) of all the requirements/goals as is shown in later sections. For the cases where projected 

performance, based on Phase 1 results, is now slightly less than the original goals it is shown in parenthesis. 

Technical specifications specific to the SLA blanket were then derived from the system requirements and the 

structural characteristics and capabilities of SquareRigger. These blanket specifications, which are shown in 

Table 9, were used as the design guideline for the SLA blanket during Phase 1.

Lens Ray Trace

Cell

Prototype SLA Panels Tensioned

Lenses

Lens Support Arch 

Cell String

Tensioned Lens Membrane 

140-Thick DC 93-500 Silicone

Tensioned 

Substrate 

5 m

2.5 m

3.75 kW SLA SquareRigger Bay – The 

Building Block of Future High 

Voltage/High power Solar Arrays 

750 kW Tug 

90 kW Array 

NASA/CP—2005-213431 79
Table 8. Original Requirements/Goals Developed for SLA SquareRigger System 

40%

120

390

500

125

Mid-Term Requirement (2013)

Ground Test Capability Pre-Launch LAPSS test at the bay level (5m x 2.5m section)

Micrometeoroid Degradation EOL Comparable to multi-junction planar array

Assumed Cell Efficiency (8AM0, 28°C) BOL 30%

Thermal Cycling Degradation EOL Comparable to multi-junction planar array

Voltage (V) EOL 100 to 1000

On-Orbit

On-Orbit

On-Orbit

EOL

On-Orbit

On-Orbit

On-Orbit

Launch

Launch

Launch

EOL

EOL

BOL

BOL

BOL

BOL

BOL

EOL

When?

Stowed Volume (kW/m 80 3)

Recurring Production Cost ($/ W) 250

Deployed Accelerations 0.001g’s 

Radiation Degradation (e Comparable to multi-junction planar array -, p+, UV)

Acoustic Vibration 146 dB OASPL for 120 seconds

Quasi Static Acceleration 22 g’s in each axis

Missions LEO, MEO, GEO & Interplanetary 

Pointing Accuracy at Base +/- 1 degrees

Deployment Time 1 hour total, 30 min structure, 30 min blanket

Deployment Torque Margin 200% (3 to 1)

System Reliability 0.99

Grounding, ESD No sustained arcs or degradation, minimize trigger arcs

Deployed Stiffness (fn Hz) > 0.1 Hz for 10kW , > 0.01 Hz for 100kW 

Stowed Stiffness (fn Hz) > 30 Hz

Contamination Degradation Comparable to multi-junction planar array

Deployed Power Density (W/m 300 2)

Specific Mass (W/kg) 330

Stowed Power (kW) 0

Deployed Power (kW) 10 to 100

Characteristic Near-Term Requirement (2008)

40%

120

390

500

125

Mid-Term Requirement (2013)

Ground Test Capability Pre-Launch LAPSS test at the bay level (5m x 2.5m section)

Micrometeoroid Degradation EOL Comparable to multi-junction planar array

Assumed Cell Efficiency (8AM0, 28°C) BOL 30%

Thermal Cycling Degradation EOL Comparable to multi-junction planar array

Voltage (V) EOL 100 to 1000

On-Orbit

On-Orbit

On-Orbit

EOL

On-Orbit

On-Orbit

On-Orbit

Launch

Launch

Launch

EOL

EOL

BOL

BOL

BOL

BOL

BOL

EOL

When?

Stowed Volume (kW/m 80 3)

Recurring Production Cost ($/ W) 250

Deployed Accelerations 0.001g’s 

Radiation Degradation (e Comparable to multi-junction planar array -, p+, UV)

Acoustic Vibration 146 dB OASPL for 120 seconds

Quasi Static Acceleration 22 g’s in each axis

Missions LEO, MEO, GEO & Interplanetary 

Pointing Accuracy at Base +/- 1 degrees

Deployment Time 1 hour total, 30 min structure, 30 min blanket

Deployment Torque Margin 200% (3 to 1)

System Reliability 0.99

Grounding, ESD No sustained arcs or degradation, minimize trigger arcs

Deployed Stiffness (fn Hz) > 0.1 Hz for 10kW , > 0.01 Hz for 100kW 

Stowed Stiffness (fn Hz) > 30 Hz

Contamination Degradation Comparable to multi-junction planar array

Deployed Power Density (W/m 300 2)

Specific Mass (W/kg) 330

Stowed Power (kW) 0

Deployed Power (kW) 10 to 100

Characteristic Near-Term Requirement (2008)

3.4 SLASR BLANKET DESIGN

The primary components of the dual-deck lens and cell panel blanket are shown in Figures 23, 24, 25, and 26. 

The ultra-thin Fresnel lenses used in the dual-deck SLA blanket, which are made of electron, proton, and UV 

radiation resistant DC 93-500 silicone, are 0.14-μm-thick and are spaced approximately 9 cm from the solar cells. 

The flexible lenses are tensioned over arches, with approximately one meter of span of lens between arches, to 

produce a curved shape that significantly increases shape-error tolerance and off-pointing performance compared 

to flat Fresnel lenses. The solar cells are mounted to thin (130micron or less) graphite-composite panels 

optimized for high stiffness and efficient waste heat removal via high thermal conductivity and emissivity. Both 

the lens film and composite panels are tensioned end-to-end between the struts of the SquareRigger structure 

when the array is deployed. The demonstrated tension to achieve precise lens shape and alignment is only 1.1 

Newton (0.25 lbf) per lens. 

Lateral lens-to-cell alignment is critical. Very accurate alignment is achieved by attaching the lens guides and 

panels to small diameter Kevlar cords tensioned straight between spreader bars, which are attached to the 

SquareRigger end struts. These primary support cords run continuously between the end interfaces under 

moderate tension and, therefore, form a near perfect datum for aligning the lens guides and panels. The support 

cord also doubles as the hinges between panels. 

(320) 

(70) 

(150) 

(110) 

NASA/CP—2005-213431 80
Table 9. Derived Technical Specification the for SLA SquareRigger Blanket

Required Pointing Accuracy (β) On-Orbit +/- 10 degrees

Storage Ground Meets all requirements after 1 year of stowage (20±5°C, <50%RH)

Operational Voltage EOL 100 V to 1000 V

8.25 kg maximum (0.64 kg/m2 Mass )

3.75 kW (290 W/m2 Power BOL )

Cell-to-cell centerline alignment (all cells) On-Orbit ± 0.025 cm (± 0.010 in) = 0.16 degrees error

Cell-to-lens centerline alignment (all cells) BOL ± 0.050 cm (± 0.020 in) = 0.31 degrees error

Deployment Force On-Orbit 66.8 N (15lb) Maximum

Grounding, ESD On-Orbit No sustained arcs or degradation, minimize trigger arcs

Micrometeoroid Degradation EOL Comparable to multi-junction planar array

Assumed Cell Efficiency (8AM0, 28°C) BOL 30%

Thermal Cycling Degradation EOL Comparable to multi-junction planar array

On-Orbit

On-Orbit

Launch

Launch

EOL

EOL

On-Orbit

On-Orbit

Launch

When?

Deployed Tension (Radiator + Lens) 44.5 N (10lb) Maximum

Radiation Degradation (e Comparable to multi-junction planar array -, p+, UV)

Acoustic Vibration 146 dB OASPL for 120 seconds

Quasi Static Acceleration 22 g’s in each axis

Deployed Dimensions 2.54 m wide x 5.08 m long (100 in x 200 in)

Required Pointing Accuracy (α) +/- 1.5 degrees

Deployment Time < 30 minutes

Contamination Degradation Comparable to multi-junction planar array

Deployed Lens- to-Cell Spacing 9.22 ± 0.05 cm (3.63 ± 0.02 in) from top of lens arch to face of cell

Stowed Dimensions 13.2 cm x 3.3 cm x 2.54 m (5.27 in x 1.35 in x 100 in)

Characteristic Blanket Specification

Required Pointing Accuracy (β) On-Orbit +/- 10 degrees

Storage Ground Meets all requirements after 1 year of stowage (20±5°C, <50%RH)

Operational Voltage EOL 100 V to 1000 V

8.25 kg maximum (0.64 kg/m2 Mass )

3.75 kW (290 W/m2 Power BOL )

Cell-to-cell centerline alignment (all cells) On-Orbit ± 0.025 cm (± 0.010 in) = 0.16 degrees error

Cell-to-lens centerline alignment (all cells) BOL ± 0.050 cm (± 0.020 in) = 0.31 degrees error

Deployment Force On-Orbit 66.8 N (15lb) Maximum

Grounding, ESD On-Orbit No sustained arcs or degradation, minimize trigger arcs

Micrometeoroid Degradation EOL Comparable to multi-junction planar array

Assumed Cell Efficiency (8AM0, 28°C) BOL 30%

Thermal Cycling Degradation EOL Comparable to multi-junction planar array

On-Orbit

On-Orbit

Launch

Launch

EOL

EOL

On-Orbit

On-Orbit

Launch

When?

Deployed Tension (Radiator + Lens) 44.5 N (10lb) Maximum

Radiation Degradation (e Comparable to multi-junction planar array -, p+, UV)

Acoustic Vibration 146 dB OASPL for 120 seconds

Quasi Static Acceleration 22 g’s in each axis

Deployed Dimensions 2.54 m wide x 5.08 m long (100 in x 200 in)

Required Pointing Accuracy (α) +/- 1.5 degrees

Deployment Time < 30 minutes

Contamination Degradation Comparable to multi-junction planar array

Deployed Lens- to-Cell Spacing 9.22 ± 0.05 cm (3.63 ± 0.02 in) from top of lens arch to face of cell

Stowed Dimensions 13.2 cm x 3.3 cm x 2.54 m (5.27 in x 1.35 in x 100 in)

Characteristic Blanket Specification

Lens Guide

Aluminum or Graphite

0.5 to 1.0 mm thick

DC 93-500 Silicone Lens

140 μm thick

Woven Graphite 

Composite Panel

125 μm thick

SquareRigger Frame

Thin Walled

M55J Graphite Load Spreader

Interface to Frame

Aluminum or Graphite

0.5 to 1.0 mm thick

Interconnected 

Triple-Junction

Solar Cells 

Bonded to Panels

Lens Guide

Aluminum or Graphite

0.5 to 1.0 mm thick

DC 93-500 Silicone Lens

140 μm thick

Woven Graphite 

Composite Panel

125 μm thick

SquareRigger Frame

Thin Walled

M55J Graphite Load Spreader

Interface to Frame

Aluminum or Graphite

0.5 to 1.0 mm thick

Interconnected 

Triple-Junction

Solar Cells 

Bonded to Panels

Figure 23. Lens Blanket Components Shown on the Phase 1 Demonstrator 

NASA/CP—2005-213431 81
Lens 9.1cm

Frame:

1. Primary solar 

array structure

Lens Guide:

1. Aligns lenses to panels (cells)

2. Keeps lens curvature consistent 

along length

3. Every 6 panels (0.8 m) End Interface:

1. Holds ends & tensions lenses

and primary support cords

Graphite panel: 

1. Supports cells

and wire

Curtain Line:

1. Keeps tension on 

blanket during 

deployment

Eyelet Guide:

1. Keeps blanket attached

to curtain line during

deployment

Primary Support Cord: 

1. Runs continuously between 

end interfaces (under tension)

2. Primary datum for blanket 

3. Supports and aligns panels and lenses

4. Serves as panel hinge

Notes:

1. Not to Scale

2. Shows primary components only

3. Full Size Blanket approximately 

36 panel sections long 

Notes:

1. Not to Scale

2. Shows primary components only

3. Full Size Blanket approximately 

36 panel sections long 

Figure 24. SLA SquareRigger Blanket Primary Components 

Figure 25. SLA SquareRigger Cell Panel Blanket

Blanket is comprised of 

108 graphite cell panels, 

3 across x 36 length

(only 6 shown)

Graphite Panel

with 30 series cells 

(0.45 A @ 60 Volts)

Cell Panel Support Cords, running on 

back of panels and tensioned between 

the end plates, carry and align cell 

panels and lens alignment arches

+

5.0 m

2.5 m

13.2cm

84.6cm

NASA/CP—2005-213431 82
CV-1142 Bond

Support cords 

align panels 

and lens 

guides with 

panels

.020”

Aluminum 

Lens Guide 

Aligns Lens 

to Cells

Approximately 30” 

(6 panels) to next 

Lens Guides

Tensioned 

support cords run 

whole length of 

blanket and are 

the primary 

alignment datum

Figure 26. Lens Guide Alignment

A typical deployment sequence for SLA SquareRigger is shown in Figure 27. During launch the dual-deck lens 

and cell blanket is stowed in a compressed stack in the cavity that exists between the SquareRigger struts and 

spars to protect them from vibration damage. As shown in Figure 28 and Table 10, the lenses are folded between 

the cell panels when stowed providing a compact blanket stack, 5.65cm (2.23 in.) tall, which is highly damped. 

The folded panels and lenses are deployed accordion style pulled by the same motor that unfolded the 

SquareRigger structural elements. Low spring rate tensioners are used to attach the lens/cell panel blanket to the 

SquareRigger frame to minimize the effect of thermally induced length change on lens and/or panel tension and 

to enhance structural stability. These designs for SLA blanket stowage and deployment are derived from the 

original thin film blanket stowage and deployment concepts. 

NASA/CP—2005-213431 83
Figure 27. SLASR Stowage and Deployment

NASA/CP—2005-213431 84
Substrate

Cell

Lens

Flexible Cord

Flex Hinge

End Interface

Lens Guide

Curtain Line

Eyelet Guide

5.275"

1.356"

To Motor

To High 

Tension

N’gator

To Low 

Tension

N’gator

Primary 

Support 

Cord

2.25”

Figure 28. SLASR Blanket Stowage 

Table 10. Stowed Blanket Thickness Calcuation 

Component Thickness Quantity Total Fraction

1/2 Cords on Back 0.005 36 0.180 8.9%

Substrate 0.005 36 0.180 8.9%

Flex Bond 0.000 36 0.000 0.0%

Flex Circuit 0.004 36 0.144 7.1%

Cell Bond 0.004 36 0.144 7.1%

Cell 0.006 36 0.216 10.7%

Glass Bond 0.002 36 0.072 3.6%

Glass 0.003 36 0.108 5.3%

Lens 0.008 36 0.288 14.2%

1/2 Lens Fold 0.012 36 0.432 21.3%

Sub Total 0.044 36 1.764

End Interface Plate 0.04 4 0.160 7.9%

Lens Guide 0.02 5 0.100 4.9%

Nominal Total 2.024 100%

Contingency 0.2024 10%

Total with Contingency 2.226

Stowed Blanket Thickness (Inches)

NASA/CP—2005-213431 85
3.5 SYSTEM ANALYSES

3.5.1 Sizing Model 

Development of an Excel spreadsheet to rapidly size and predict all key SLASR metrics was completed during the 

Phase 1 program. The model allows an analyst to quickly and accurately predict SLASR array size, power and 

mass for any reasonable configuration of bay size, quantity and arrangement. The model contains 22 detailed 

worksheets addressing the various details of the design such as mass properties, strength, stiffness, power, 

environmental degradation and packaging in a single Excel file. 

The model performs a detailed power analysis. On-Orbit EOL electrical performance is calculated from: 

• Measured BOL cell performance 

• Detailed radiation dose calculation 

• Operating temperature 

• Harness Losses 

• On-orbit degradation 

• Radiation, contamination, UV, micrometeorites, thermal cycling etc. 

The detailed mass analysis is based on prototype SquareRigger and SLA hardware as available. The model 

contains a complete parts list and there is a mass breakout and contingency for each component based on its 

maturity. Harness mass is determined by wire count and length. A mass contingency, which is currently running 

at about 5% at the system level, is included in all reported metrics and is calculated for each component based on 

the following criteria: 

• 2% for actual or PRO/E designed part 

• 10% for calculated 

• 15% for estimated 

The model also includes approximate structure strength and stiffness calculations to ensure that these very 

important features are consistent with the stated requirements. The key metrics of mass (W/kg) and stowed 

power (kw/m3

) calculated from the sizing model for already demonstrated 30% efficient concentrator cell efficiency 

is shown in Figure 28. A typical mass breakdown is also shown in Figure 29. A summary of a typical power 

analysis performed for a high power 15 year GEO mission is shown in Figure 30. 

Figure 29. Calculated SLASR Mass and Stowed Power Density and Typcical Mass Breakdown 

SLA SquareRigger Performance Versus Power

0

50

100

150

200

250

300

350

0 10 20 30 40 50 60 70 80

BOL Wing Power - On Orbit (kW)

Specific Power (W/kg)

0

10

20

30

40

50

60

70

80

90

100

Stowed Power Density (kW/m 3

)

kW/m3

W/kg

Structure

9%

Mechanism

Harnessing 7%

15% Substrates

3%

Yoke Assy

1%

Lenses & Cell 

Panels

61%

Root Assy

2%

Tiedowns

2%

NASA/CP—2005-213431 86
Isc Imp Vmp Voc Pmp Efficiency

Bare Cell Performance, BOL 28 °C 0.0177 0.0170 2.389 2.714 0.0406 30.00%

Temperature Coeffcients °C-1 1.00E-05 7.00E-06 -0.0065 -0.0059 -9.37E-05 -0.0693%

SLA Triple Junction Cell Data at 1 AM0 and 28 °C

 Amps/cm2 Volts

Cell Length (cm) 3.50

Cell Active Width (cm) 1.00

Lens Wdith (cm) 8.50

Geometric Concentration Ratio 8.50

Lens Gap or Blocking at Edges (cm) 0.076

Lens Optical Efficiency 0.92

Effective Optical Concentration Ratio 7.68

Net Optical Efficiency 0.904

Effective Cell Area (cm2) 26.88

Concentration Voltage Boost Factor 1.09

Input to Array 

IV curve calculator

15yr GEO Mission 15yr GEO Mission

System

Reference Value Loss Factor Isc Imp Vmp Voc Pmp Efficiency

0.476 0.457 2.604 2.958 1.19 29.54%

LAPSS Calibration to AM0 1.000 0.476 0.457

CIC Assembly Voltage Loss 0.995 2.591 2.943

CIC Glassing Loss 0.995 0.473 0.454

Sun Distance 1.0000 AU 1.000 0.473 0.454

Alpha Off-Pointing 1.000 0.473 0.454

Beta Pointing (Cosine) 3.00 deg 0.999 0.473 0.454

Beta Pointing (Shadowing) 1.000 0.473 0.454

UV degradation, Lens 0.980 0.463 0.445

UV degradation, Cell Cover 0.980 0.454 0.436

Micro-Meteoroid and Debris 0.990 0.449 0.432

Contamination on Lens and Cell 0.980 0.440 0.423

Isc Radiation Degradation 4.20E+14 1MeVe/sqcm 0.990 0.436

Imp Radiation Degradation 5.20E+14 1MeVe/sqcm 0.978 0.414

Vmp Radiation Degradation 5.20E+14 1MeVe/sqcm 0.899 2.330

Voc Radiation Degradation 5.93E+14 1MeVe/sqcm 0.897 2.639

CIC Operating Temperature 85.6°C

Isc Temperature Correction 10.0 uA/sqcm/C 0.0155 A 0.451

Imp Temperature Correction 7.0 uA/sqcm/C 0.0108 A 0.425

Vmp Temperature Correction -6.5 mV/°C -0.3744 V 1.956

Voc Temperature Correction -5.9 mV/°C -0.3398 V 2.299

Cell-to-Cell Interconnect Voltage Drop 0.01 ohms -0.0042 V 1.952

0.451 0.425 1.956 2.299 0.83 20.63%

Series Cells Per Module 4 4 0.451 0.425 7.824 9.196

Module-Module IC Voltage Drop 0.01 ohms -0.0042 V 7.819

0.451 0.425 7.819 9.196 3.32 20.62%

Thermal Cycle Current Loss 0.995 0.449 0.422

Thermal Cycle Voltage Loss 0.995 7.780 9.150

Cell/Module Stringing Mismatch 0.99 0.418

Lens Stringing Mismatch 0.99 0.414

Number of Series Modules per String 15 116.704 137.256

0.449 0.414 116.704 137.256 48.32 20.01%

Blocking Diode Voltage Drop 0.7000 V 116.004

Panel Harness Voltage Drop 0.28 ohms 0.1139 V 115.891

Strings per Power Circuit 6

Wing Harness Voltage Drop 0.95 ohms 2.3563 V 113.534

0.449 0.414 113.534 137.256 47.01 19.46%

EOL CIC Performance

Cell Module Performance

String Performance on Panel 

String Performance at SADA/Spacecraft Connector

Cell Under Lens @ 28°C

Peak Power

Power at 

Selected 

Load Points

Number of Strings = 1188 Volts Amps Watts

100.00 514.07 51406.7

Voc Isc Vmp Imp Pmp FF 110.00 503.48 55382.3

137.26 533.60 113.53 491.87 55843.7 0.76 113.53 491.87 55843.7

Performance at Load Voltage

EOL Equinox Performance

0

10000

20000

30000

40000

50000

60000

0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150

Volts

Power (Watts)

0

100

200

300

400

500

600

Current (A)

Figure 30. SLASR EOL Power Analysis Example for 56kW 15-year GEO Mission

NASA/CP—2005-213431 87
3.5.2 Detailed Structural Analysis 

ANSYS Finite Element models were developed for 14 kW, 42 kW and 56 kW wings to calculate natural frequency 

and deflection due to acceleration. It was determined that the baseline SquareRigger structure (tube crosssection and thickness), originally developed to support a thin film photovoltaic blanket, meets the SLASR stiffness 

requirements (frequency > 0.1 Hz for 10 kW Array - 2 Wings, frequency > 0.01 Hz for 100 kW Array - 2 Wings, 

and maximum a rotation < 1 Degree @ 0.01 g). Results of the finite element calculations are shown in Figure 31. 

3.6 MID-TERM (10 YEAR) TECHNOLOGY ROADMAP

All near-term analysis and performance projections are based on the existing designs and current technology 

capabilities, the most important of which is cell efficiency. Thus, a SLASR array constructed from existing 

components will achieve the near-term performance given in Table 11. Over the next ten years significant 

improvements are forecast for many key SLASR technologies. For example mass produced crystalline solar cells 

for space have achieved an average improvement of approximately 1% per year between 1995 (19% SJ GaAs) 

and 2003 (27.5% ATJ GaInP2/GaAs/Ge) based on the production of over 1.5 million cells and this trend is 

expected to continue. The key improvements in SLASR technology are listed in Table 11 along with the 

corresponding impact on the key metrics of efficiency (W/m2

), cost ($/W), mass (W/kg) and stowed power density 

(kW/m3

). 

Table 11. Mid-Term (10 year) Technology Roadmap for SLASR

Component 

Improvement

Current 

Performance

Projected 

Performance Basis W/m2 W/kg $/W kW/m3

Cell Efficiency 30% Efficiency 40% Efficiency EMCORE and SPL Technology 

Roadmap Trends 1.33 1.33 1.33 1.33

Solar Cell Module 

Mass

2.51 grams 

(4cells)

1.51 grams 

(4 cells)

Use 1/2 thickness cell being developed 

by EMCORE. Interconnect Cells 

without copper flex circuit. Co-cure 

kapton to panel.

1.00 1.09 1.00 1.05

Lens Mass & Cost 140 μm thick 115 μm thick Improved processing and materials 1.00 1.03 1.10 1.02

Panel Mass 100 μm thick 75 μm thick Higher conductivity graphite fibers 

including carbon nanotubes 1.00 1.05 1.00 1.02

Mechanism Mass 75% reduction 

from current Design and material optimization 1.00 1.02 1.05 1.00

Structural Size & 

Mass

75% reduction 

from current

Higher modulus graphite fibers 

including carbon nanotubes 1.00 1.02 1.10 1.10

Total Improvement Factor 1.33 1.64 1.69 1.60

Current Performance 300 320 250 70

10 Year Performance 400 523 148 112

Improvement

3.7 DEMONSTRATION HARDWARE AND TESTING

Based on the full scale SLASR design concepts developed, detailed drawings for the Phase 1 SLA SquareRigger 

prototype were prepared. As shown in Figures 32 through 35, the demonstration blanket contains four 12.7cm 

(5.0”) x 34.0cm (13.4”) woven graphite panels, four approximately 53.3cm (21”)-long lenses that are supported by 

a lightweight graphite SquareRigger frame. The lenses were manufactured and bonded to the aluminum lens 

guides by ENTECH. The lens and cell panel blankets are tensioned to approximately 4.5 Newtons (1lbf) each 

and this equates to approximately 4% stretch of the lenses. Left over (scrap) two-junction GainP2/GaAs/Ge 

TECSTAR cells from the DS1 SCARLET program were installed on the panels to enhance the quality of the 

demonstrator. Two strings, of 3 cells in series, were installed and interconnected on one of the panels for 

electrical performance testing. The optical alignment and flatness of the prototype was exceptional as proven by 

the ease at which all cells can be simultaneously brought into focus as shown in Figure 36. 

The blanket deployment concept was successfully verified by manually tensioning the blanket draw cords 

simulating tensioning of the cords by an electric motor as shown in Figure 37. The power production of the two 3-

cell strings was measured using ABLE’s Large Area Pulsed Solar Simulator (LAPSS) with and without the lenses 

installed to record system efficiency. The data from string 1 is shown in Figure 38. The increase in power and 

current from the lenses was as expected with a measured lens efficiency of 90%. The power and current 

increase for string 2 was similar to string 1, though it was noted that one of the cells in the strings was damaged, 

resulting in string 2 having low voltage compared to string 1. 

NASA/CP—2005-213431 88
Figure 31. SLASR FEA Stiffness Calculations for Several Configurations Showing 

Compliance to Requirements 

 

15 m 

14 kW Class

fn = 0.16 Hz 

β

α 

Maximum Rotation at 0.01 g 

α = 0.11 deg 

β = 0.79 deg

15 m 

42 kW Class

fn = 0.11 Hz 

Maximum Rotation at 0.01 g 

α = 0.57 deg 

β = 1.43 deg

25 m 

56 kW Class

fn = 0.04 Hz 

Maximum Rotation at 0.01 g 

α = 0.64 deg 

β = 5.69 deg

Requirements

• Frequency > 0.1 Hz 

(10 kW Array= 2 Wings) 

• Frequency > 0.01 Hz 

(100 kW Array = 2 Wings)

• Maximum a rotation < 1 

Degree @ 0.01 g 

NASA/CP—2005-213431 89
Figure 32. SLA SquareRigger Demonstration Blanket Components

Lightweight 

SquareRigger 

Frame 

17”

25” 

4X Woven 

Graphite Panel 

13.4” x 5.0”

DS1 SCARLET 

Solar Cells on All 

Panels 

21” Long Stretched 

Lenses 

Two electrically 

 Active Strings 

(3 DS1 Cells in Series)

13.4

5.0

DS1 Solar Cells

Woven Graphite Panel

0.005” Thick

Lens guides 

(Thin Aluminum Sheet)

NASA/CP—2005-213431 90
Figure 33. Phase 1 SLASR Prototype Top View

Lens Guide

Aluminum or Graphite

0.5 to 1.0 mm thick

DC 93-500 Silicone Lens

140 μm thick

Woven Graphite 

Composite Panel

125 μm thick

SquareRigger Frame

Thin Walled

M55J Graphite Load Spreader

Interface to Frame

Aluminum or Graphite

0.5 to 1.0 mm thick

Interconnected 

Triple-Junction

Solar Cells 

Bonded to Panels

Lens Guide

Aluminum or Graphite

0.5 to 1.0 mm thick

DC 93-500 Silicone Lens

140 μm thick

Woven Graphite 

Composite Panel

125 μm thick

SquareRigger Frame

Thin Walled

M55J Graphite Load Spreader

Interface to Frame

Aluminum or Graphite

0.5 to 1.0 mm thick

Interconnected 

Triple-Junction

Solar Cells 

Bonded to Panels

Figure 34. Phase 1 SLASR Prototype Oblique View

Figure 35. Phase 1 SLASR Prototype Side View

Gossamer 

Lens and 

Cell 

Blankets 

NASA/CP—2005-213431 91
Figure 36. Optical Verification of Phase 1 SLASR Demonstration Blanket

Lenses perfectly 

image cells verifying 

optical alignment 

and flatness

Lenses perfectly 

focus sunlight on 

cells verifying 

optical alignment 

and flatness 

NASA/CP—2005-213431 92
Figure 37. Phase 1 SLASR Demonstration Blanket Deployment

Prototype SLASR String 1

3 "Scrap" DS1 TECSTAR 2J Cells in Series

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

0.50

0123456789

Volts

Amps

With 

Lens

Without 

Lens

Ratio 

(With/Without)

Ratio/GCR 

GCR = 8.23

Isc 0.449 0.061 7.40 0.90

Imp 0.383 0.054 7.06

Vmp 6.340 5.460 1.16

Voc 7.810 7.210 1.08

Pmax 2.430 0.297 8.19 1.00

Prototype SLASR String 1

With Lens

Without Lens

Figure 38. Phase I SLASR Demonstration Blanket LAPSS Data

NASA/CP—2005-213431 93
3.8 ACKNOWLEDGMENT

This work was administered though the NASA Glenn Research Center and funded by the NASA SBIR Phase 1 

program under contract #NAS3-03017. 

3.9 SLASR SUMMARY

Excellent progress has been made in integrating a Stretched Lens Array photovoltaic blanket into the 

SquareRigger solar array structure. The result is an extremely lightweight, high efficiency, high power, high 

voltage, and low stowed volume solar array that may enable numerous future NASA missions such as Space 

Solar Power (SSP), high-power deep space SEP probes, Solar Electric Propulsion (SEP) tugs and large earthorbiting science satellites. 

REFERENCES: 

1) S. White, M. Douglas, B. Spence, P. Jones and M. Piszczor, “Development of an UltraFlex-Based Thin 

Film Solar Array for Space Applications”, 3rd WCPEC, May 2003 with, and of GRC. 

2) D. Murphy, M. Eskenazi, B. Spence & P. Jones, “Thin Film & Crystalline Solar Cell Array Performance 

Comparisons, IEEE Conference, 2002.

3) P. Jones, J. Harvey, S. White & B. Smith, “A High Specific Power Solar Array for Low to Mid Power 

Spacecraft”, SPRAT Conference, 1994. 

4) D. B. Snyder, "Dynamic Interactions Between Ionospheric Plasma and Spacecraft," 75th Anniversary 

U.R.S.I. Symposium, Brussels, Belgium, 26-27 April 1995. 

5) M. Kruer and J. Lyons, "The FAST Solar Array: Challenging Requirements, Novel Design," 1st World 

Conference on Photovoltaic Energy Conversion, December 5-9, 1994. 

6) S. Davis, R. Stillwell, W. Andiario, D. Snyder, and I. Katz, "EOS-AM Solar Array Arc Mitigation Design," 

1999 Intersociety Energy Conversion Engineering Conference, Vancouver, BC Canada, August 2-6, 

1999. 

7) T.G. Stern, et. al., “Development of an Electrostatically Clean Solar Array Panel”, IEEE Conference, 2002. 

8) B. Spence, D. Murphy, P. Jones, and M. Eskenazi, “The SCARLET Array for High Power GEO Satellites, 

26th IEEE Photovoltaic Specialists Conference, September 1997. 

9) O’Neill, M., et al, The Stretched Lens Array (SLA): A Low-Risk, Cost-Effective Concentrator Array Offering 

Wing-Level Performance of 180 W/kg and 300 W/m2

 at 300 VDC, Proceedings of the 37th Intersociety 

Energy Conversion Engineering Conference, July 29–31, 2002, Washington, DC IECEC 2002 Paper No. 

20115. 

NASA/CP—2005-213431 94
DEVELOPMENT OF A HIGH EFFICIENCY

MECHANICALLY STACKED MULTI-JUNCTION SOLAR CELL

Daniel Aiken, Paul Sharps, Mark Stan

Emcore Photovoltaics, Albuquerque, NM 87123

Harry Atwater, Anna Fontcuberta i Morral, James Zahler,

California Institute of Technology, Pasadena, CA 91125

Mark Wanlass

National Renewable Energy Laboratory, Golden, CO 80401

INTRODUCTION

Monolithic, 2-terminal, epitaxially grown multi-junctions represent the state-of-the-art in high efficiency photovoltaic

space power. Their in-situ monolithic integration results in an elegant device structure with high efficiency,

relatively high specific power, and a simple fabrication process. The monolithic, epitaxially grown nature of these

devices also imposes materials and design restrictions which impede the march to significantly higher

performance. These devices presently require lattice matched epitaxial templates (substrates) which limits the

materials choices for integration. The barriers to producing high photovoltaic quality epitaxial materials on lattice

mismatched templates have not been easy to surmount, despite extensive effort in the research community. Also,

series interconnection of subcells is by far the most applicable connectivity approach for these monolithic devices,

which further limits the connectivity design space.

In comparison, the mechanically stacked approach is advantageous because it relaxes or completely removes the

material and design restrictions inherent in monolithic approaches. Device designers can conceivably integrate

subcells of a chosen bandgap regardless of their lattice parameter, and in any connectivity arrangement. The

benefits of alternative connectivity arrangements have been demonstrated [1].

In contrast to the elegance of monolithically integrated solar cells, mechanically stacked solar cells can be

described as the “brute force” approach to higher performance. Mechanically stacked approaches have been

investigated for over 25 years, but have been deployed in space only on small-scale experimental testbeds.

Mechanically stacked approaches have historically been limited by several barriers. The cost associated with

integrating devices on multiple substrates has been economically prohibitive. The substrates used as templates

for III-V epitaxial solar cell structures comprise a very large fraction of the total cell cost. Techniques such as the

CLEFT process [2] or epitaxial lift-off [3] have been developed to allow substrate removal and re-use, but they are

difficult to implement cost-effectively in a production environment. Ex-situ integration is another barrier which

increases the processing complexity and cost, and also results in performance losses associated with parasitic

losses inherent to common interconnect schemes. Excessive weight also limits the practicality of these devices

for space applications because of the multiple substrates and/or fixturing hardware used to stack the subcells

together.

Here we present results of work performed on a new mechanically stacked solar cell approach for achieving very

high efficiency and high specific power for space concentrator applications. This work addresses some of the cost

and practicality limitations present in many historical mechanically stacked approaches, and is being funded by the

NASA Office of Space Science through the Advanced Cross-Enterprise Technology Development Program. The

goal of this project is to demonstrate the critical technology elements necessary for producing a 40% efficient,

lightweight solar cell for concentrator arrays. A specific power of 1kW/kg is projected at the cell level.

NASA/CP—2005-213431 95
DEVICE CONCEPT

The device structure being developed is shown in Figure 1. A high bandgap tandem based on the well-understood

InGaP/GaAs materials combination is grown on an inactive GaAs or germanium substrate. By similarity, this

tandem device structure capitalizes on the highly refined triple junction device structure already in high volume

production at Emcore. Subsequent glass encapsulation provides a superstrate for mechanical support of the

active solar cell, and a substrate selective removal process is used to decrease the weight of this high bandgap

tandem.

Figure 1 A schematic of the mechanically stacked 4-junction device integration scheme under development

Growth and processing of a low bandgap 1 eV InGaAsP / 0.74 eV InGaAs tandem is performed independently of

the top tandem on an engineered InP-on-silicon substrate. The engineered substrate in this application is

designed to provide an epitaxial template for active layers lattice matched to InP, while also benefiting from the

lower weight, lower cost, and superior mechanical stability of silicon as compared to InP. Silicon’s mass density is

less than half that of InP, and the superior mechanical strength of silicon permits the use of thinner substrates. A

thinner, lower density substrate allows a much higher power to weight ratio than would be possible with an InP

substrate. A thin InP film, shown in red in Figure 1, is transferred to a silicon host substrate from a re-useable InP

source wafer using a process that will be described in the next section.

A number of different schemes for mechanical and electrical interconnection of the two independently grown and

processed devices are being considered. As shown in Figure 2, this combination of bandgaps is nearly ideal for a

four junction series interconnected solar cell. This combination of bandgaps is therefore very compatible with two

terminal, series interconnected integration, provided that the tandem interconnect is sufficiently transparent so as

to transmit a near-theoretical spectral content to the bottom tandem. Achieving a highly transparent interconnect

while also maintaining both a low electrical specific resistivity and good mechanical stability at the interconnection

is critical to the success of this two terminal approach. Here we note that transmission loss through the tandem

interconnect can be partially compensated for by reducing the thickness of the top tandem, thereby decreasing the

fundamental, above-gap absorption in the top tandem.

ENGINEERED SUBSTRATES

The InP-on-silicon engineered substrate is being developed at Cal Tech using the Smart-cut process originally

developed by Bruel et al. for silicon-on-insulator applications [4]. The process is illustrated schematically in Figure

3. An InP wafer is implanted with hydrogen ions to a controllable projected range. The implanted side of the InP

GaAs or Ge substrate

interconnect formation

low-bandgap tandem growth

engineered substrate:

InP on silicon

high-bandgap tandem growth

(selective removal)

high efficiency, lightweight

2-terminal, 4-junction

solar cell

NASA/CP—2005-213431 96
wafer is then bonded to the silicon host substrate using modern direct wafer bonding methods [5]. The InP/Si

wafer pair is annealed in a controlled ambient at elevated pressure and temperature. The anneal simultaneously

strengthens the direct bond and perforates the InP wafer at the implanted region. The released InP substrate can

then be re-polished and repeatedly used in the layer transfer process. The engineered substrate can then in

principle be used to grow a high quality low bandgap tandem solar cell.

Figure 2 Bandgap versus lattice constant for the common compound semiconductor materials. Red lines

indicate the ideal bandgaps for a four junction series interconnected solar cell under the AM0 spectrum

at 25oC. Dotted lines indicate the bandgap combinations being investigated in this work.

Figure 3 A process for producing InP-on-silicon engineered substrates based on the Smart-cut technology.

Much of the work at Cal Tech has focused on perfecting the Smart-cut process for InP/Si and developing an

engineered substrate that is epi-ready. This requires a surface which is both free of dislocations that might

propagate into the epitaxial film, and has a characteristic surface roughness similar to that of epi-ready polished

InP.

InP-based

bottom tandem

GaAs/Ge-based

top tandem

O3 O3

O

O

CO2 CO2

• Hydrophobic surface

passivation:

• organic clean

• acetone

• methanol

• ultra-violet ozone

• Oxide removal  HF

• <5 Å rms roughness

H+

Donor Substrate

• H-implant of donor substrate to

desired peak range

• Room temperature bond initiation  ~1MPa

Handle Substrate

clean, hydro

-phobic surface

Device Substrate

Pressure, Temperature

• Wafer bonding and layer splitting:

• pressure  accommodate

thermomechanical stress

• temperature  blister; covalent bond

micro-crack

formation and

layer transfer

• Bonded Heterostructure  ~1 μm film

NASA/CP—2005-213431 97
Double heterostructures grown on as-transferred InP/Si engineered substrates have produced much lower

photoluminescence intensity as compared to control samples grown on thick InP substrates. One potential cause

of the presumed lower material quality on InP/Si is due to the presence of a region of highly defective InP at the

surface of the engineered substrate, presumably caused by the ion implantation and layer transfer process.

Figure 4 shows cross-sectional TEM images of the InP/Si pair, clearly showing an area of defective material at the

exposed surface. More importantly, InP near the bonded interface appears to be less defective than the surface

region. Polishing processes such as plasma etching, chemo-mechanical polishing, and wet chemical polishing

etches are being investigated as means for removing the surface damage, thereby creating a surface more

conducive to high quality epitaxial growth.

High resolution TEM images have revealed the presence of an oxide at the InP/Si interface. This interface layer is

formed during the anneal in a nitrogen ambient, presumably resulting from residual water at the clean interfaces

just prior to bonding. Despite the presence of this interfacial oxide layer, through-the-film current voltage

measurements indicate an ohmic, low resistance behavior. The specific resistivity of the interface is less than 0.1

Ω-cm

2

.

Figure 4 Cross-sectional TEM micrographs of an InP/Si engineered substrate. The images suggest a region of

damaged, highly defective material near the implanted region of the InP.

(a) (b)

Figure 5 (a) A high resolution TEM image indicating the presence of an interfacial oxide. (b) Through-the-film

current-voltage characteristics of an InP/Si pair, suggesting less than 0.1 ohm-cm

2

specific resistivity.

100 nm

defective region

bonded InP/Si interface

5 nm

Si

InP

Interfacial

oxide

   







    	 
  




      
     
 

 

      
    

	    	 
           

	    




 

C

urre

nt ( A.c

m-2 )

Voltage (V)

NASA/CP—2005-213431 98
TANDEM INTERCONNECTION

One of the tandem interconnect approaches being evaluated is a eutectic metal bond at the interface between two

perpendicularly oriented metallic grids, as depicted in Figure 6(a). Advantages of this approach include the

absence of a fixturing jig to hold the two cells optically in series, the absence of a peripherally formed interconnect

which otherwise complicates the fabrication procedure, and the elimination of the need for pattern alignment in

joining the two grids. Gold-germanium melts at a eutectic temperature of 356oC. Grid fingers coated with a goldgermanium alloy can therefore be bonded together at low temperatures using moderate pressure. The eutectic

metal bond approach is being experimentally evaluated using test structures such as that shown in Figure 6(b).

Gridded germanium wafers have been bonded together to test the mechanical stability and electrical conductivity

of the resulting interconnect. Figure 7(a) suggests intimate contact between the grid finger and germanium

substrate. The resulting test structures are mechanically robust and produce an interconnect with very low

specific resistivities of 0.024 Ω-cm

2

and 0.014 Ω-cm

2

for 1600 μm and 800 μm grid spacings, respectively. This

type of interconnect scheme results in a modest 2-4% transmission loss due to grid shading.

(a) (b)

Figure 6 (a) A conceptual schematic of the gridded interconnect scheme being evaluated. (b) A cross-sectional

optical micrograph (100x) showing a metallic grid successfully bonded to two germanium wafers.

(a) (b)

Figure 7 (a) A cross-sectional optical micrograph (1000x) showing metallic grid successfully bonded to two

germanium wafers. (b) Through-the-film current voltage characteristics of germanium wafers with a

gridded, bonded interface.

grid finger

cross sections

Ge

Ge

100x

1000x

-200

-150

-100

-50

0

50

100

150

200

-0.006 -0.004 -0.002 0 0.002 0.004 0.006

Volts

m

A/cm2

800 μm grid spacing

1600 μm grid spacing

NASA/CP—2005-213431 99
BOTTOM TANDEM DEVELOPMENT

Prototype lattice matched bottom tandems consisting of 1eV InGaAsP, a low bandgap tunnel diode, and 0.74 eV

InGaAs are being grown and processed on InP. The performance of these devices serves as a benchmark for

devices to be grown on the InP/Si engineered substrates. The bottom tandem is projected to contribute 9-10%

absolute to the overall device efficiency at ~10 suns AM0. Devices grown and processed at NREL have reached

6.1% efficiency without AR coating at 1 sun AM0, as shown in Figure 8. These 0.108 cm

2

devices were measured

under a simulated AM0 spectrum filtered by a low resistivity GaAs wafer coated on both sides with an AR coating.

Figure 8 Current density-voltage characteristics of a 1.0 eV InGaAsP, 0.74 eV InGaAs prototype tandem on InP,

measured under a fintered AM0 spectrum at 25oC.

ANTI-REFLECTION COATING DEVELOPMENT

The 4-junction, nominally current matched solar cell being developed here can achieve an unprecedented AM0

efficiency of 40%, which is near the theoretical limits of a 4-junction device. This is made possible by the highly

efficient utilization of photons over an extremely broad spectral range (300–1700 nm). Achieving a performance

near the theoretical limits requires near-unity quantum efficiency for all subcells in the solar cell structure. This

places requirements on AR coating performance that have previously not been necessary. The AR coating

transmission band of the top tandem must be much broader that is currently necessary for production triple

junction solar cells.

With the realization that the standard AR coating used for production triple junction solar cells would result in

substantial loss in a 4-junction cell, we have designed an improved AR coating using methods presented in earlier

work [6]. Anti-reflection coating design is accomplished here with optical modeling software which both optically

and electrically models device performance as a function of the AR coating structure. A new AR coating design as

presented in Figure 9 results in a modeled 1% absolute efficiency gain over the standard dual layer AR coating.

Table I summarizes the relevant modeled device parameters.

SUMMARY

An affinity for high performance solar cells in the space market, as well as the development of new engineered

substrate technology, has warranted a re-consideration of mechanically stacked solar cell technologies.

Mechanically stacked approaches result in a wider materials and interconnectivity design space compared to

monolithic approaches. We are investigating the Smart-cut technology to produce InP/Si engineered substrates.

Successful transfer of InP films to a silicon substrate has been demonstrated, and the resulting structure is

-15

-13

-11

-9

-7

-5

-3

-1

1

-0.1 0.1 0.3 0.5 0.7 0.9

Volts

m

A/cm2

Voc: 813 mV

Jsc: 13.4 mA/cm2

FF: 77%

Eff. 6.1%

NASA/CP—2005-213431 100
mechanically stable after MOCVD growth. The electrical resistivity of the bonded interface is low, despite the

presence of a very thin interfacial oxide. Work continues in regard to producing an epi-ready surface for the

subsequent growth of high quality MOCVD layers. Tandem cell interconnectivity approaches are being

investigated which can result in minimal parasitic loss and are also simple to process. Bottom tandem prototype

cells have been fabricated on InP, with encouraging results. New AR coatings have also been designed to

improve the optical coupling to these extremely wide-band solar cells.

Figure 9 Modeled reflectance of a top tandem solar cell. A new AR coating design is compared to Emcore’s

standard AR coating structure.

Table I Predicted performance metrics for a 4-junction cell, nominally current matched, using the reflectance

design modeled in Figure 9.

std. design new design

modeled efficiency (%) 34.23 35.67

maximum Jsc (mA/cm

2

) 17.86 17.86

achieved Jsc (mA/cm

2

) 15.54 16.65

weighted transmission (%) 87.03 93.21

REFERENCES

[1] L. Fraas, J. Avery, D. Scheiman, “AM0 calibration of 34% efficient mechanically stacked GaInP/GaAs-GaSb

Circuits”, Proceedings of the 29

th

IEEE Photovoltaics Specialists Conference, 2002, pp. 912-915.

[2] R.W. McClelland, C.O. Bozler, J.C.C. Fan, “A technique for producing epitaxial films on reuseable substrates”,

Appl. Phys. Lett., vol 37, no. 6, September 1980, pp. 560-562.

[3] E. Yablonovitch, T. Gmitter, J.P. Harbison, R. Bhat, “Extreme selectivity in the lift-off of epitaxial GaAs films”,

Appl. Phys. Lett., vol. 51, no. 26, December 1987, pp. 2222-2224.

[4] M. Bruel, B. Aspar, A.J. Auberton-Herve, “Smart-cut: A new silicon on insulator material technology based on

hydrogen implantation and wafer bonding”, Jpn. J. Appl. Phys., vol. 36, 1997, p. 1636.

[5] Q.-Y. Tong, U. Gösele, “Semiconductor Wafer Bonding”, John Wiley & Sons, 1999.

[6] D. J. Aiken, “High performance anti-reflection coatings for broadband multi-junction solar cells”, Solar Energy

Materials and Solar Cells, vol. 64, 2000, pp. 393-404

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

300 500 700 900 1100 1300 1500 1700

wavelength (nm)

reflectance

std design

improved design

NASA/CP—2005-213431 101
RECENT ADVANCES IN HIGH EFFICIENCY III-V MULTIJUNCTION SOLAR CELLS 

FOR SPACE APPLICATIONS: ULTRA TRIPLE JUNCTION QUALIFICATION 

Hojun Yoon, Jennifer E. Granata, Peter Hebert, Richard R. King, Christopher M. Fetzer, 

Peter C. Colter, Kenneth M. Edmondson, Daniel Law, Geoffrey S. Kinsey, Dmitri D. Krut, 

James H. Ermer, Mark S. Gillanders, and Nasser H. Karam 

Spectrolab, Inc., Sylmar, California 

INTRODUCTION 

The performance of GaInP/GaAs/Ge-based multijunction (MJ) space solar cells has continually 

advanced in the past several years, satisfying the ever-increasing power requirements of satellites. The 

first commercial satellite that employed this technology was launched in 1997 (HS 601HP), which used 

solar arrays populated with Spectrolab’s dual junction (DJ) GaInP/GaAs/Ge cells, with beginning of life 

(BOL) efficiency of 21-22% [1]. Building on this DJ benchmark, triple junction cells based on the same 

materials system have improved the performance to a new level. For example, Spectrolab’s Improved 

Triple Junction (ITJ) cells which have been in production for the past 3 years, have BOL efficiency of 

greater than 26.5%, providing ~20-25% more power than the DJ cells [2]. Today, there are more than 

225 kW of multijunction arrays on orbit that are performing to predict, a clear demonstration of the heritage and technology soundness of Spectrolab’s MJ cells. 

Spectrolab’s latest triple junction production space solar cell – Ultra Triple Junction (UTJ) – exhibits key performance advances over the ITJ cell, and pushes the performance of triple junction cells to yet 

a higher level [3]. The UTJ cell has a BOL average efficiency at maximum power of 28.0% (AM0, 28oC, 

135.3 mW/cm2

) and P/Po = 0.86 for 1 MeV electrons after 1E15 e/cm2 fluence. More importantly, the efficiency at an operating temperature of 60oC after 15 years in GEO environment is predicted to be 22.6%, 

which is higher by 5.5% relative than the ITJ cell. The purpose of this report is to describe some of the 

UTJ cell features as well as some of the key BOL and end of life (EOL) performance results stemming 

from the most comprehensive to date space-qualification program. The details of the qualification program, which includes electrical, mechanical, radiation, and environmental tests, have been described 

previously [3]. 

CELL DESIGN 

Device Structure

Figure 1 shows a schematic cross-section of a 

baseline triple junction cell structure, upon which the 

UTJ cell is based. It comprises of lattice matched (LM) 

GaInP and GaAs based subcells grown epitaxially on a 

p-type Ge substrate using metal organic vapor phase 

epitaxy (MOVPE) process. The precise lattice-matching 

of the epilayers to the Ge lattice constant promotes the 

growth of high crystallinity materials, and is one of the 

key factors in the resulting high performance device. 

Details of the cell development efforts, including optimization for EOL performance, that led to the final UTJ 

device structure have been described previously [4, 5].

Contact

p-Ge Substrate

n-Ge

n-GaAs buffer

n++-GaAs

p++-GaAs

p-GaInP

p-GaAs

n-GaInP

ICTJ

p-BSF

p-GaInP

n-GaInP

n-AlInP

AR

n-GaAs

n+-GaAs

Contact

Top

Middle

Bottom

Wide Bandgap

Top Cell Window

High Bandgap

Top Cell

Lattice-Matched

Structure

Contact

p-Ge Substrate

n-Ge

n-GaAs buffer

n++-GaAs

p++-GaAs

p-GaInP

p-GaAs

n-GaInP

ICTJ

p-BSF

p-GaInP

n-GaInP

n-AlInP

AR

n-GaAs

n+-GaAs

Contact

Top

Middle

Bottom

Wide Bandgap

Top Cell Window

High Bandgap

Top Cell

Lattice-Matched

Structure

Fig. 1. Cross section of a baseline triple junction GaInP/GaAs/Ge solar cell, highlighting 

some of the key features of the UTJ cell.

NASA/CP—2005-213431 102
While the GaAs-based middle cell of 

the UTJ cell is unchanged from that of the ITJ 

cell, the GaInP top cell has been modified to 

enhance its performance. First, the GaInP top 

cell is grown disordered, resulting in a higher 

bandgap material, and hence a higher cell 

voltage. The ordering refers to the atomic arrangement of the group-III sublattice sites by 

the Ga and In atoms. In a disordered GaInP 

material, the Ga and In atoms occupy the 

group-III sites in a random manner, whereas in 

an ordered GaInP material, they are arranged 

in a systematic manner. Second the AlInP top 

cell window material has been made more 

optically transmissive, providing a greater 

amount of incident light into the cell. The resulting UTJ structure has a higher cell load 

voltage by 80 mV, while maintaining the same 

current density compared to the ITJ cell. Consequently, the UTJ cell has an efficiency at load = 27.7%. A comparison of the light IV curves (AM0, 

28oC, 135.3 mW/cm2

) of a typical UTJ and ITJ cell is shown in Fig. 2. 

Minority carrier lifetime is an important parameter that strongly influences the solar cell device 

performance, and the disordered GaInP material has been measured to exhibit long lifetimes, in the tens 

of nanosecond range as measured by time resolved photoluminescence [4], nearly approaching the values measured on ordered GaInP material. These values translate to minority carrier diffusion lengths that 

are many times the cell base thickness, which is desired for high cell performance. 

Cell-Interconnect-Coverglass (CIC) Assembly

In addition to the improved device structure, the 

UTJ cell uses a 20% thinner (140 micron) Ge substrate 

compared to a standard ITJ cell, enhancing the powerto-weight ratio. All joints including the interconnects and 

the diode assembly onto the cell are welded for improved manufacturability. Two interconnects are welded 

on the cell front side and redundant “H-tab” diode assembly welded on the back side, with double-stick Kapton tape used for attachment and isolation of the diode 

assembly to the back side of the cell. The interconnects 

and the H-tabs are both made from silver-clad Kovar 

material. A photograph showing both the front and back 

sides of the UTJ CIC assembly is presented in Fig. 3. 

BOL PERFORMANCE 

 The efficiency at load (EffLoad) distribution of more than 1,000 large area (26.41 cm2

) UTJ bare 

cells from the qualification build is shown in Fig. 4. All measurements were performed using Spectrolab’s 

X-25 solar simulator (AM0, 28oC, 135.3 mW/cm2

), calibrated using balloon flight-traceable standards. 

The average EffLoad of this population is 28.1%, while the average efficiency at maximum power (Effmp) is 

28.3%. The highest Effmp measured to be 29.4%. Table 1 summarizes all the photovoltaic parameters 

for this population. As previously discussed, the improved performance stems primarily from the voltage 

increase – the load voltage for the UTJ cell has been increased by 80 mV (~3.5% increase) over the ITJ 

cell, while the current density remains approximately the same. The resulting efficiency makes the UTJ 

cell the highest efficiency space solar cell available today. 

0.000

0.002

0.004

0.006

0.008

0.010

0.012

0.014

0.016

0.018

0.0 0.5 1.0 1.5 2.0 2.5 3.0

Voltage (V)

Current (A/cm2

)

ITJ

UTJ

UTJ ITJ

I

sc (mA/cm2) 17.0 16.9

Imp (mA/cm2) 16.1 16.0

Voc (V) 2.660 2.565

Vmp (V) 2.345 2.270

VLoad (V) 2.310 2.230

FF (%) 83 84

EffLoad (%) 27.7 26.5

Effmp(%) 28.0 26.8

0.000

0.002

0.004

0.006

0.008

0.010

0.012

0.014

0.016

0.018

0.0 0.5 1.0 1.5 2.0 2.5 3.0

Voltage (V)

Current (A/cm2

)

ITJ

UTJ

UTJ ITJ

I

sc (mA/cm2) 17.0 16.9

Imp (mA/cm2) 16.1 16.0

Voc (V) 2.660 2.565

Vmp (V) 2.345 2.270

VLoad (V) 2.310 2.230

FF (%) 83 84

EffLoad (%) 27.7 26.5

Effmp(%) 28.0 26.8

Fig. 2. Comparison of light IV curves (AM0, 28oC, 135.3 

mW/cm2) of a typical UTJ and ITJ cell, highlighting the 

higher voltage and efficiency of the UTJ cell.

Front

Back

Diode Assembly

Fig. 3. Photograph of UTJ CIC assembly.

NASA/CP—2005-213431 103
 

The high efficiency performance of the UTJ cells builds upon the heritage of the ITJ cells which 

have been in production at Spectrolab since 2000. Figure 5 shows the EffLoad distribution of more than 

350,000 large area ITJ bare cells, with average EffLoad = 26.7%. Compared to this value, the UTJ qualification average is 5% higher, at EffLoad of 28.1%. 

Table 1. Average BOL photovoltaic parameters for large area UTJ bare cells.

Parameter 

Jsc

(mA/cm2

) 

Jmp

(mA/cm2

) 

Voc

(V) 

Vmp

(V) 

JLoad (2.31V)

(mA/cm2

) 

EffLoad (2.31V)

(%) 

Effmp

(%) 

FF 

(%) 

Qual Avg. 

(1045 cells) 

17.09 16.26 2.669 2.352 16.44 28.1 28.3 83.7 

Target/Spec 17.00 16.16 2.660 2.345 16.22 27.7 28.0 83.0 

EOL PERFORMANCE 

All space cell designs must take into account EOL performance in order to meet the spacecraft 

power needs throughout a typical mission life of 15 years. The key parameters that are necessary to 

predict on-orbit power performance at EOL are radiation resistance and temperature coefficients. Various 

electron and proton radiation energies present in space induce varying degrees of degradation to the solar cell. Taking these effects into account for a typical 15-year GEO mission, a 1 MeV electron equivalent 

fluence can be calculated in order to provide a single radiation degradation factor for EOL performance. 

This value has been determined for the UTJ cell to be approximately 6E14 e/cm2

, similar to Spectrolab’s 

other MJ solar cells. 

1 MeV Electron Radiation

Results of 1 MeV electron radiation effects on UTJ cell performance are summarized in Fig. 6, 

which plots the normalized degradation of the current (NIsc, NImp), voltage (NVoc, NVmp), and power (NPmp) 

as a function of fluence. Five fluences ranging from 3E13 to 3E15 e/cm2

 were used, with 15 cells (2 cm x 

2 cm) tested at each fluence. Similar to other MJ cells, the voltage generally degrades faster than the 

0

0.05

0.1

0.15

0.2

0.25

Fraction of Cells

<26.1 26.4 26.8 27.2 27.6 28 28.4 28.8

Efficiency @ 2.310V (%)

Fig. 4. Efficiency at load (2.310 V) distribution 

of 1,045 large area UTJ bare cells tested under 

AM0, 28oC, 135.3 mW/cm2

 condition. The average EffLoad = 28.1%.

Fig. 5. Efficiency at load (2.230 V) distribution of 

351,024 large area ITJ bare cells tested under 

AM0, 28oC, 135.3 mW/cm2

 condition. The average EffLoad = 26.7%.

NASA/CP—2005-213431 104
current, but the overall NPmp after 1E15 

e/cm2

 fluence has improved to 0.86, 

highest value achieved to date among 

all MJ cells. At 6E14 e/cm2

 (1 MeV 

electron equivalent fluence at EOL for a 

15-year GEO mission), the NPmp is 0.89. 

Temperature Coefficients

 It is well known that with increasing temperature, the efficiency of 

the solar cell decreases as a result of 

the competing effect between the voltage and current. The voltage decreases 

largely due to the increase in the saturation current and also due to a shift in the 

bandgap. The current increases due to 

a shift in the bandgap as well as a general increase in the carrier lifetime. The 

combined effects cause the cell efficiency to decrease with increasing temperature because the voltage drop is 

higher than the current gain. In MJ cells where the current produced by the entire stack is limited by the 

lowest current-generating subcell, the current balance among the subcells can also influence the measured current temperature coefficients. 

Temperature coefficients have been measured on UTJ cells after subjecting them to various 1 

MeV electron fluences. Figure 7 plots the cell efficiency at maximum power measured as a function of 

temperature in the range 15-75oC. The temperature coefficients (absolute % efficiency per oC) do not 

vary significantly as a function of fluence, ranging from -0.076%/oC for BOL condition to -0.065%/oC for 

3E15 e/cm2

 fluence. At 6E14 e/cm2 (1 MeV electron equivalent fluence at EOL for a 15-year GEO mission), the efficiency temperature coefficient is -0.070%/oC. 

EOL Power Estimate

 Using the 1 MeV electron radiation data for the UTJ cells, the remaining power after a 15-year 

GEO mission can be estimated. In addition, the temperature coefficients determined for various 1 MeV 

0.75

0.80

0.85

0.90

0.95

1.00

1.E+13 1.E+14 1.E+15 1.E+16

1 MeV Electron Fluence [e/cm2

]

Fitted Normalized Degradation

NIsc

NImp

NVoc

NVmp

NPmp

NPmp

NIsc

NImp

NVmp

NVoc

Fluence NPmp

5E14 0.89

1E15 0.86

6E14* 0.89

0.75

0.80

0.85

0.90

0.95

1.00

1.E+13 1.E+14 1.E+15 1.E+16

1 MeV Electron Fluence [e/cm2

]

Fitted Normalized Degradation

NIsc

NImp

NVoc

NVmp

NPmp

NPmp

NIsc

NImp

NVmp

NVoc

0.75

0.80

0.85

0.90

0.95

1.00

1.E+13 1.E+14 1.E+15 1.E+16

1 MeV Electron Fluence [e/cm2

]

Fitted Normalized Degradation

NIsc

NImp

NVoc

NVmp

NPmp

NPmp

NIsc

NImp

NVmp

NVoc

Fluence NPmp

5E14 0.89

1E15 0.86

6E14* 0.89

Fig. 6. Effects of 1 MeV electron radiation on UTJ cell performance. *6E14 e/cm2

 is the 1 MeV equivalent fluence after a 15-

year GEO mission.

16

18

20

22

24

26

28

30

10 20 30 40 50 60 70 80 90

Cell Temperature [o

C]

Cell Efficiency at Max Power [%]

Fluence %/°C

BOL -0.076

3E13 -0.073

1E14 -0.073

5E14 -0.071

1E15 -0.068

3E15 -0.065

Efficiency

temperature

coefficient

16

18

20

22

24

26

28

30

10 20 30 40 50 60 70 80 90

Cell Temperature [o

C]

Cell Efficiency at Max Power [%]

Fluence %/°C

BOL -0.076

3E13 -0.073

1E14 -0.073

5E14 -0.071

1E15 -0.068

3E15 -0.065

Efficiency

temperature

coefficient

Fig. 7. UTJ cell efficiency at maximum power measured as a function of temperature, for various 1 MeV electron 

fluences. The temperature coefficient values are noted in terms of absolute % efficiency per oC.

NASA/CP—2005-213431 105
electron fluences can be used to approximate the remaining power at an operating temperature. Table 2 

provides these estimates, in comparison to the ITJ, TJ, and DJ cells, using 60oC as an operating temperature. At EOL (60oC), the UTJ cell is projected to provide 5.5% more power than the ITJ and 25% 

more power than the DJ cell. The higher performance of the UTJ cell is attributed to an increase in the 

BOL efficiency as well as an improvement in the radiation resistance. 

Table 2. Power estimate for UTJ cells, compared to ITJ, TJ, and DJ cells. EOL condition assumes 15-year GEO 

mission and 1 MeV electron equivalence of 6E14 e/cm2

. Operating temperature of 60oC is assumed. 

UTJ 

2003 

ITJ 

2000 

TJ 

1999 

DJ 

1997 

Effmp at BOL, AM0, 28oC 28.0% 26.9% 25.1 21.7% 

P/Po (6E14 e/cm2, 1 MeV) 0.89 0.87 0.87 0.88 

Temp. Coeff. at EOL, AM0 -0.070%/oC -0.063%/oC -0.053%/oC -0.031%/oC 

Effmp at EOL, AM0, 28oC 24.9% 23.4% 21.8% 19.1% 

Effmp at EOL, AM0, 60oC 22.6% 21.4% 20.1% 18.1% 

SUMMARY 

Key performance results stemming from the Ultra Triple Junction cell qualification program have 

been described. The BOL average efficiency at maximum power for UTJ cells is 28.0% (AM0, 28oC, 

135.3 mW/cm2

), while the P/Po = 0.86 for 1 MeV electrons after 1E15 e/cm2 fluence. These results combined with temperature coefficient measurements, translate to an efficiency of 22.6% at an operating 

temperature of 60oC after a 15-year GEO mission. This represents a 5.5% relative higher efficiency at 

EOL compared to the previous generation ITJ cells, providing the highest power of any previously available space cell technology. Along with the improved bare cell performance, the UTJ CIC technology has 

been improved to provide enhanced manufacturability at the assembly level, adding to the many benefits 

of the latest production UTJ space solar cell technology. To date, the UTJ bare cell qualification tests 

have been completed, meeting all pass/fail criteria. The CIC assembly qualification is expected to be 

completed in October 2003. 

 Despite the impressive efficiency increases achieved from the GaInP/GaAs/Ge based triple junction cells in the past few years, there remain opportunities for further improvements. Spectrolab has identified these opportunities and is pursuing the next generation cell, the XTJ space solar cell, which is expected to provide 30% efficiency at BOL. As always, EOL performance will be emphasized as well as 

manufacturability in order to provide the highest value product at a lowest possible cost. 

REFERENCES 

[1] M.R. Brown, L.J. Goldhammer, G.S. Goodelle, C.U. Lortz, J.N. Perron, J.S. Powe, J.A. Schwartz, B.T. 

Cavicchi, M.S. Gillanders, and D.D. Krut, “Characterization Testing of Dual Junction GaInP2/GaAs/Ge 

Solar Cell Assemblies,” Proc. 26th IEEE Photovoltaic Specialists Conf. (IEEE, New York, 1997), p. 805. 

[2] J.E. Granata, J.H. Ermer, P. Hebert, M. Haddad, R.R. King, D.D. Krut, M.S. Gillanders, N.H. Karam, 

and B.T. Cavicchi, “Advancements in GaInP2/GaAs/Ge Solar Cells – Production Status, Qualification Results and Operational Benefits,” Proc. 29th IEEE Photovoltaic Specialists Conf. (IEEE, New York, 2002), p. 

824. 

[3] J.E. Granata, C. Fetzer, J.H. Ermer, R.R. King, K. Edmondson, A. Stavrides, P. Hebert, P. Colter, 

G.S. Kinsey, H. Yoon, M.S. Gillanders, N. Beze, K. Bui, J. Hanley, N.H. Karam, and B.T. Cavicchi, “Ultra 

Triple Junction GaInP2/GaAs/Ge Solar Cells – Cell Design and Qualification Status,” presented at the 3rd

World Conf. On Photovoltaic Solar Energy Conversion, Osaka, Japan, May 2003. 

NASA/CP—2005-213431 106
[4] R.R. King, C.M. Fetzer, P.C. Colter, K.M. Edmondson, J.H. Ermer, H.L. Cotal, H. Yoon, A.P. 

Stavrides, G. Kinsey, D.D. Krut, and N.H. Karam, “High-Efficiency Space and Terrestrial Multijunction Solar Cells Through Bandgap Control in Cell Structures,” Proc. 29th IEEE Photovoltaic Specialists Conf.

(IEEE, New York, 2002), p. 776. 

[5] G.S. Kinsey, R.R. King, K.M. Edmondson, A.P. Stavrides, H. Yoon, C.M. Fetzer, P.C. Colter, J.H. Ermer, M.S. Gillanders, P. Hebert, J.E. Granata, and N.H. Karam, “Ultra Triple-Junction High-Efficiency Solar Cells,” IEEE AESS Systems Magazine, March 2003, p. 8. 

NASA/CP—2005-213431 107
HIGH EFFICIENCY, MULTI-JUNCTION CELLS WITH MONOLITHIC BYPASS DIODES 

P. R. Sharps, M. A. Stan, D. J. Aiken, 

B. Clevenger, J. S. Hills, and N. S. Fatemi 

EMCORE Photovoltaics, Albuquerque, NM 

INTRODUCTION 

The manufacture of solar panels for both space and terrestrial applications involves stringing cells together such 

that they are connected electrically in series. If a solar cell in a series connected string is shaded, with the 

remaining cells still in the light, the cell is forced into reverse bias breakdown so that it can conduct the current of 

the string. The reverse biasing of the cell into breakdown, particularly for III-V multi-junction solar cells, will 

permanently damage the cell. To protect the cells from breakdown, discrete silicon bypass diodes are electrically 

connected in parallel to the cell, but with opposite polarity. If a cell is shaded, the diode is forward biased, and the 

string current safely passes through the diode. Several different approaches to attaching a discrete diode to the 

solar cell have been proposed and are in use, including front mounted, corner mounted (in-plane with the cell), 

and back mounted diodes [1]. 

The connection of the discrete silicon diode to the cell is done during the CIC (coverglass-interconnect) operation, 

and involves front and back interconnects to both the diode and the cell. Automation of the CICing process is 

complicated because of the number of parts involved, i.e., the cell, the diode, and the front and the back 

interconnects for the diode. 

 

A monolithic bypass diode, integral to the solar cell and not a discrete part, has the advantage of reducing the 

number of parts required during the CICing operation. However, there are several design criteria that the 

monolithic bypass diode must meet. First, it has to protect the cell, i.e., it has to be able to pass the string current 

without causing the cell to go into reverse bias breakdown. Secondly, it has to be reliable. The diode has to be 

stable in a space environment, as well as perform repeatedly without failure. Finally, it has to be simple. The 

inclusion of the diode must not add growth or processing complexity such that the reduction in CICing costs are 

outweighed by the cost of the monolithic diode. 

DIODE DESIGN 

The dark electrical schematic of the cell and diode is shown in Figure 1. As can be seen, the diode and the multijunction cell are electrically parallel. We have looked at two different approaches for the monolithic bypass diode, 

a Schottky junction approach and an epitaxially grown p-n junction approach. For performance and reliability 

reasons we have chosen the epitaxially grown p-n junction approach. 

For the Schottky diode approach, the diode can be located within the cell structure. What is required is a 

semiconductor layer that, when the appropriate metal is deposited on it, will create a Schottky junction with 

polarity opposite to that of the multi-junction cell. Figure 2 is a schematic of one Schottky junction approach, 

where the junction is created using one of the semiconductor buffer layers. In Figure 2 the Schottky diode is 

placed in the buffer layers, above the Ge junction, but beneath the epitaxially grown III-V layers. However, the 

Schottky diode could just as easily be placed on top of the III-V epitaxially grown layers. It is important to 

remember that the actual junction is at the metal semiconductor interface. During processing, the semiconductor 

layer has to be exposed (through appropriate etching), and the necessary metal deposited. A mesa etch also has 

to be done to electrically isolate the diode area from the rest of the cell. 

NASA/CP—2005-213431 108
Figure 1. Electrical schematic of the monolithic bypass diode, under dark conditions. 

Figure 2 also shows “metal shorts”, necessary to short out the junction underneath the Schottky diode. If these 

“metal shorts” are not included, the Schottky diode is in series with another junction underneath of opposite 

polarity. For the Schottky diode to protect the cell, the junction underneath of opposite polarity has to be forced 

into reverse breakdown to allow the current to be conducted. The “metal short” resolves this issue by providing 

an electrical short around the opposite polarity junction underneath the Schottky diode. The “metal shorts” also 

remove the necessity of including welded straps or “C-clamps” for completing the electrical circuit. The usual 

interconnect tab welded to the top of the Schottky diode and the cell bus pads completes the circuit. Emcore has 

both US and international patents pending on the “metal shorts” approach. 

MJ Solar Cell 

Metal 

Grid 

 

n +-

GaAs ARC 

Multi-junction 

Solar Cell 

Layers 

Monolithic Bypass Diode 

 

 Metal 

Short 

Schottky Metal 

Contact 

 Schottky 

Semiconductor 

Layer 

 Buffer Layers 

 Ge Junction 

Ge Substrate 

Metal Contact 

Figure 2. Schematic of the Schottky diode approach. Note that neither the horizontal nor the vertical distances 

are to scale. The monolithic bypass diode area covers a much smaller portion of the cell (e.g., 4 mm2

 for the 

diode versus 30 cm2

 for the cell) than depicted here. 

Monolithic

Bypass Diode

Bottom Subcell

Top Subcell

Middle Subcell

NASA/CP—2005-213431 109
Figure 3 shows a schematic of the epitaxially grown p-n junction approach. In this approach the bypass diode is 

grown on top of the multi-junction cell. The complete diode is epitaxially grown, and is of opposite polarity to the 

multi-junction device. The diode layers are removed from the areas above the multi-junction cell, and a mesa 

etch isolates the bypass diode area from the cell. As with the Schottky diode approach, a “metal shorts” approach 

is again used to short out the multi-junction cell layers underneath the bypass diode. The amount of processing 

required for either the Schottky or epitaxial diode approach are nearly identical. 

 

MJ Solar Cell Monolithic Bypass 

Diode 

 Metal Pad 

 p-GaAs 

Metal 

Grid 

 n-GaAs 

n +-GaAs 

ARC 

n+-GaAs 

Multi-junction 

Solar Cell 

Layers 

 Metal Short 

Multi-junction 

Solar Cell Layers 

 Ge Junction 

Ge Substrate 

Metal Contact 

Figure 3. Schematic of the epitaxial p/n diode approach. The same comments regarding scales and areas 

mentioned for Figure 2 apply for this Figure. 

In terms of electrical performance, the key diode parameters are the reverse bias leakage current (IRB) (measured 

at a value close to the cell load voltage, i.e., ~2.5 volts), the forward bias turn on voltage (VFB) (approximately how 

much voltage is required to pass the load current), and the diode series resistance (RS) (which is a measure of 

how much I2

R losses occur in the diode). Because the bypass diode is electrically parallel to the cell, any diode 

reverse bias leakage current reduces the total current through the cell/diode combination when the cell is at load. 

Hence, minimizing the bypass diode reverse leakage is critical to having a high efficiency cell/diode combination. 

DIODE PERFORMANCE 

 

We have found that the performance of the epitaxially grown p-n bypass diode is better than that of the Schottky 

diode. The reason for this is straight forward. Since the Schottky junction occurs at the metal/semiconductor 

interface, any variability in processing or contamination of the semiconductor surface will have an effect on the 

performance of the diode. Schottky diodes are very dependent on the surface conditions of the semiconductor 

prior to deposition of the metal. Since etching is required to expose the semiconductor surface, chemical residue 

can be left behind on that surface. Also, as the surface is exposed to air, an oxide layer will be created on the 

semiconductor surface, the thickness of which is time dependent. These varying conditions, inherent during 

processing of the diode, will alter the metal/semiconductor barrier height, and will in turn effect both the leakage 

current and forward bias turn on voltage. In addition, during the welding of the interconnect to the Schottky diode, 

the interconnect tab is actually welded to the metal part of the Schottky junction. The heat and pressure from the 

weld can affect the metal/semiconductor interface, further changing the properties of the junction. 

The epitaxial p/n junction approach has several advantages over the Schottky approach, one of which is that the 

junction interface can be precisely controlled during the MOCVD growth. The thickness and doping levels of the 

NASA/CP—2005-213431 110
p- and n-layers are precisely controlled, and the junction is also made in a tightly controlled environment (purified 

H2 in the MOCVD reactor), similar to the epitaxial growth of the solar cell. Hence, the manufacturing repeatability 

issues associated with the Schottky diode are avoided. In addition, welding to the epitaxial p-n diode is identical 

to welding to contact pads on the solar cell. A cap layer is included on top of the p-n junction, and the junction is 

removed from the metal to which the interconnect is welded. 

Testing of the diode in reverse bias indicates a considerable amount about a diode’s robustness. Because the 

bypass diode is in the reverse biased condition when the cell is at load (at approximately 2.3 volts), not only does 

the leakage current thorough the diode have to be minimal, it also has to be stable. Since these bypass diodes 

may operate for up to 15 to 20 years in space, there cannot be pre-mature failures. 

As part of the design phase of developing a monolithic bypass diode, we have taken both the Schottky and p/n 

diodes into far reverse bias to determine where breakdown occurs, what kind of breakdown mechanisms are 

involved, and how stable the breakdown is. We looked at two different types of p/n junctions, which we refer to as 

Design 1 and Design 2. Because of U.S. and International patents pending, we cannot discuss the details of 

Design 1 and 2. 

-50

-40

-30

-20

-10

0

-35 -30 -25 -20 -15 -10 -5 0

RB Voltage, Volts

RB Current, microamps

Series1

Series2

Series3

Design 2

Design 1

Schottky

Figure 4. Reverse bias I-V curves for Schottky, Design 1, and Design 2 monolithic bypass diodes. Note the 

qualitative difference between the Schottky and Design 1 diodes and the Design 2 diode. See text for further 

discussion. 

The results of the reverse bias tests are shown in Figure 4. The vertical axis on the graph is in micro-amps. 

Testing the cells in reverse bias at these currents is a sensitive measure of the type of breakdown that is 

occurring. Note the qualitative difference between the breakdown characteristics between p/n diode Design 2 and 

the other two diodes, the Schottky diode and p/n diode Design 1. Design 1 and the Schottky design show 

“microplasma” breakdown, associated with crystalline imperfections [2]. In the case of Design 1 and the Schottky 

design, we have been able to see the current pulses associated with the microplasmas on a curve tracer. The 

reverse bias I-V curves for the Schottky and Design 1 diodes will literally “flicker” as the microplasmas turn on and 

off. Because of the association of the microplasmas with crystalline defects and ensuing concerns about on orbit 

NASA/CP—2005-213431 111
device reliability, both the Schottky approach and the Design 1 approach are deemed undesirable. However, the 

reverse bias characteristic of Design 2 is clearly different, and indicates that it can support significant reverse bias 

voltage before breaking down. 

The breakdown seen in Design 2 in Figure 4 is consistent with the avalanche breakdown mechanism. Avalanche 

breakdown has a well defined, abrupt breakdown knee, while other types of breakdown (Zener and crystalline 

imperfection breakdowns), have a softer breakdown knee. In looking at the curve in Figure 4 for Design 2, the 

breakdown is abrupt, indicating that avalanche breakdown is the dominant mechanism. In addition, with 

avalanche breakdown, the breakdown voltage increases with increasing temperature. For Zener and crystalline 

imperfection breakdown, the breakdown voltage decreases with increasing temperature. To confirm that 

avalanche breakdown was controlling the Design 2 diode, the reverse bias characteristics of the diode were 

examined as a function of temperature. The results are shown in Figure 5. 

-50

-40

-30

-20

-10

0

-35 -30 -25 -20

Reverse Bias Voltage, volts

Current, microamps

20 C

30 C

40 C

50 C

60 C

70 C

80 C

90 C

100 C

Figure 5. Current-voltage curves as a function of reverse bias voltage, for various temperatures, for one of the 

epitaxially grown p-n junctions. 

Taking 30 microamps as a point to demarcate the onset of breakdown, it can be seen from Figure 5 that as the 

temperature increases from 20 oC to 70 oC, the voltage for the onset of breakdown increases in magnitude. After 

70 oC, the breakdown voltage decreases in magnitude with increasing temperature. In addition, the breakdown 

knee starts to soften around 60 oC. The conclusion is that up to 60 oC, the breakdown is through the avalanche 

mechanism, but as the temperature increases above this another mechanism controls the breakdown. Zener 

breakdown is associated with heavy doping on each side of the junction, and a thin depletion region. Since that is 

not the case in the Design 2 diode, it is thought that a localized crystalline imperfection starts to control the 

breakdown at higher temperatures. During the breakdown/temperature measurement it was noted that reverse 

biases of greater than 20 volts were still required to breakdown the Design 2 diode when the diode was at 150 oC. 

The Design 2 diode is robust and able to withstand temperatures of up to 150 oC without having any deleterious 

effect on solar cell operation. 

 

DIODE QUALIFICATION 

For use in space, the monolithic bypass diode has to pass several qualification steps. These include a welding 

test, a high temperature reverse bias soak, a high temperature forward bias soak, a high temperature cycling test, 

and performance comparisons before and after exposure to particle radiation. Since the standard diode 

NASA/CP—2005-213431 112
incorporated into the cells is 4 mm2

 in area, all of the diode qualification tests were done on Design 2 diodes of 

this size. Several tests were also done on the Schottky type diodes. 

The welding test was of particular concern for the Schottky diode approach. Since the junction in the Schottky 

diode actually occurs at the metal/semiconductor interface, any high temperature welding to the contact metal 

always had the possibility of altering that interface. We have found that the Schottky junction can be particularly 

sensitive to welds, with an increase in the reverse bias leakage current at 2.5 volts (IRB) due to the welding. The 

p/n junctions, both Designs 1 and 2, were much less sensitive to the welding process, since the actual junction is 

completely within the semiconductor. One result of the welding study was the institution of the criteria for a 

successful weld. The IRB before and after the weld process cannot change by more than 30 microamps. We 

found that it was much more difficult for the Schottky diodes to pass this bench mark. 

The reverse bias soak, the forward bias soak, and the cycling tests were all done under ambient atmosphere at a 

chamber temperature of 120 °C. The reverse bias soak involved holding the diode at 2.5 volts reverse bias for 

1,000 hours. During the forward bias soak, 650 mA was passed through the diode for 250 hours. For the cycling 

test, the diodes were switched between 2.5 volts reverse bias and 650 mA over 100,000 times. The dwells at 

each of the biases were 1 second, with an approximately 1 milli-second switching time. For all of the tests, the 

diode performance was measured before and after. The particle irradiation involved exposure to both 1 MeV 

electrons, and 10 MeV protons. For 1 MeV electrons, the diodes were exposed to a fluence of 1 x 1015 e/cm2

, 

while for 10 MeV protons, the diodes were exposed to a fluence of 1 x 1012 p/cm2

. 

 

The results of the diode qualification tests for the Design 2 diode are shown in Table 1. As can be seen, all of the 

tests were passed. 

Table 1. Summary of space qualification tests on the monolithic bypass diode Design 2, with 4 mm2

 area. IRB

was measured at 2.5 volts reverse bias, and ∆VFB refers to the change in the forward bias voltage needed to pass 

500 mA through the diode before and after the particular test. See text for further discussion. 

Test Item Requirement Results Status

Reverse 

Bias 

IRB < 30 μA 

∆VFB<10% 

IRB < 1 μA 

∆VFB=1.6% 

Pass 

Forward 

Bias 

IRB < 30 μA 

∆VFB<10% 

IRB < 1 μA 

∆VFB=0.4% 

Pass 

Diode 

Cycling 

IRB < 30 μA 

∆VFB<10% 

IRB < 1 μA 

∆VFB=0.2% 

Pass 

Irradiation: 

 Electrons 

 Protons 

IRB < 30 μA 

∆VFB<10% 

Electrons: 

 IRB < 1 μA 

∆VFB<1% 

Protons: 

 IRB < 1 μA 

∆VFB<1% 

Pass 

To further characterize the diode, the temperature of the diode was measured during forward biasing of the diode. 

The cell and diode were under ambient atmosphere in a chamber with a temperature of 120 oC. Vary small 

thermocouples were placed on the cell, 5 mm away from the diode, and also on the diode itself. The diode was 

forward biased to 650 mA, and the temperature of the cell and diode brought to a steady state. The cell 

temperature and the ambient chamber temperature were 120 oC, while the diode temperature rose to 133 oC. 

The implications of this are that the diodes tested above actually were at 133 oC, not at the ambient 120 oC. The 

I

2

R heating of the diode is generating some heat, which is being dissipated through the diode and into the cell. 

Further diode temperature measurements are underway in a simulated space environment, with cells mounted on 

substrates, and the measurements being done in a vacuum of less than 10-5 torr. 

NASA/CP—2005-213431 113
CELL PLUS DIODE PERFORMANCE 

 

We have made a number of cells with the monolithic bypass diode. Figure 6 shows the I-V curve for one of these 

cells, along with the device performance parameters. Figure 7 shows a histogram of the performance of 2,000 

large area (>30 cm2

) devices. The average efficiency for this group of cells is 27.5% (135.3 mW/cm2). In 

addition, we are working to increase the average lot efficiency to 28.5%. Figure 8 is a histogram of some recent 

data for some higher efficiency lots. 

Figure 6. I-V curve for a 30.2 cm2

 multi-junction solar cell with a Design 2 monolithic bypass diode. The cell Voc, 

Isc, ff, and efficiency are 2621 mV, 517 mA, 84.5%, and 28%, respectively, under 135.3 mW/cm2

. 

26.0 26.5 27.0 27.5 28.0 28.5 29.0

0

100

200

300

400

500

600

Y Axis Title

X Axis Title

Figure 7. Histogram for over 2,000 multi-junction cells grown with monolithic bypass diodes. All of these cells 

were 30 cm2

 in area or larger. The average efficiency is 27.5%, under 135.3 mW/cm2. 

Number of Cells 

-1.5

-1

-0.5

0

0.5

1

-2 -1 0 1 2 3

Voltage, volts

Current, amps

Series1

Series2 Dark

Light

Efficiency 

NASA/CP—2005-213431 114
Figure 8. Histogram for over 200 higher efficiency cells. The average cell efficiency is 28.5%, under 135.3 

mW/cm2

. 

CONCLUSIONS 

We have been able to demonstrate a monolithic bypass diode that meets the design criteria mentioned in the 

Introduction. The diode performs as intended as it prevents the cell from being reverse biased into breakdown. It 

is robust and reliable, as demonstrated by the qualifications tests. Finally, it is simple in design and does not add 

significant complexity during growth and processing. 

REFERENCES 

[1] See, for example, U.S. Patents 6,326,540 and 6,103,970 

[2] M. S. Tyagi, Introduction to Semiconductor Materials and Devices, (John Wiley & Sons, New York, 1991) pp. 

214-235. 

27.6

27.8

28

28.2

28.4

28.6

28.8

29

29.2

NASA/CP—2005-213431 115
DENDRITE GROWTH AND DEGRADATION IN MULTI-JUNCTION SOLAR 

CELLS 

Dr. Donna Speckman 

Dr. Dean Marvin 

John Nocerino 

The Aerospace Corporation, Los Angeles, CA 

Contact Author

Dr. Donna Speckman 

The Aerospace Corporation 

PO Box 92957 MS M2/244 

Los Angeles, CA 90009 

Donna.m.speckman@aero.org

310 336 9399 

ABSTRACT 

Silver dendrites have been shown to be responsible for shunting of p/n multijunction solar cell 

assemblies with silver plated interconnects. A variety of cell and interconnect designs has been 

studied in order to identify the key factors that result in silver dendrite formation and subsequent cell 

shunting. The experimental results suggest that in this case the shunting was due to the particular 

polarity of the cell and design of the interconnect that placed it in close proximity to the cell edge. As 

a result, mitigation strategies are readily available. 

INTRODUCTION 

Silver coated interconnects have historically been used for many years on single junction (SJ) solar 

cells for space applications, with no apparent difficulties. However, with the recent introduction of 

high efficiency, multi-junction solar cells with higher operation voltages for space-based solar array 

applications, some problems have arisen that appear to be related to the use of silver based 

interconnects. Solar panels constructed using p/n dual junction (DJ) cascade cells showed a severe 

degradation in performance upon long-term storage in room light and normal humidity conditions. 

Destructive analysis of solar cells removed from this panel revealed that the edges of the solar cells 

were highly decorated with silver dendrite growths, but the mechanism and conditions required for 

dendrite growth were not elucidated. Furthermore, identical solar cell assemblies that had not been 

mounted on a panel, but instead were stored in room air with occasional exposure to light, also 

were covered with dendrites along the cell edge. The performance degradation in the stored panels 

is believed to be due to the growth of dendrites that formed shunt paths across the cell junction. 

Although metal migration is well known in the printed circuit board industry1

, this is the first instance 

known to us of the phenomenon affecting solar cells. 

NASA/CP—2005-213431 116
The formation of silver dendrites in the junction region of a solar cell edge is a multi-step 

process. Since there is no silver inherently present in this region of the cell, silver ions must first be 

generated elsewhere. These ions must then be transported to the junction region, where they are 

reduced to silver metal. If enough silver metal is present to bridge the solar cell junction, a highly 

conductive shunt can be formed. There are several cell features that could contribute to the 

formation of dendrites. These include: a) cell voltage (to ionize silver, transport the ions to the cell 

edge, and reduce the ions at the cell junction), b) cell polarity (to transport silver ions towards the 

junction), c) amount of exposed silver in close proximity to the cell edge (interconnect design, 

presence of a crossbar, bus bar and solder pad locations), and d) presence of electrolytes on the 

cell surface to aid in ion transport (solder flux, contamination, moisture, solvents). In order to 

understand the mechanism of dendrite growth, and to identify the conditions that lead to this 

phenomenon, we have carried out a series of experiments to assess the influence of each of these 

features on the growth of dendrites and subsequent cell shunting. 

EXPERIMENTAL SECTION 

A variety of glassed solar cells with different polarities, cell voltages, interconnect designs, 

and interconnect attachment methods were obtained from vendors for comparison in our studies. 

Single junction p/n cells (Air Mass Zero open circuit voltage of as-received cells ~ 1.0v) were 

evaluated, as were several types of multi-junction cells (Air Mass Zero open circuit voltage of asreceived cells ~ 2.4v). Multi-junction cells with p/n polarity were compared with n/p multi-junction 

cells, and two interconnect designs were assessed: a "Z-bend" interconnect, which wraps closely 

around the solar cell edge and has the option of having a crossbar lying directly over the cell edge, 

and an "out-of-plane stress relief" interconnect, which does not contain a crossbar and does not 

conform closely to the cell edge. The influence of electrolytes from solder flux was evaluated by 

comparing interconnects attached to the cells either by soldering (using no-clean solder flux) or by 

welding. Interconnected solar cell assemblies obtained for testing included 1) p/n dual junction 

GaInP/GaAs/Ge (Tecstar) cells with soldered Z-bend interconnects with crossbars, 2) n/p dual 

junction GaInP/GaAs/Ge (Spectrolab) cells with soldered Z-bend interconnects with crossbars, 3) 

p/n dual junction GaInP/GaAs/Ge (Tecstar) cells with welded, out-of-plane stress relief 

interconnects with no crossbars, and 4) p/n single junction GaAs/Ge (Tecstar) cells with soldered Zbend interconnects with no crossbars. 

For our experiments, each interconnected cell assembly was suspended in a Petri dish 

above a fluid whose vapor might be in contact with a solar panel during storage, such as water, 

ketoxime (a cure product of the cell/substrate adhesive), or a mixture of both solvents. A control 

sample, in which no solvent was added to the Petri dish (room humidity exposure only), was also 

included for each set of samples studied. The assemblies were illuminated with quartz halogen 

lamps (approximately 0.2 suns) to achieve an approximate open circuit cell voltage while exposed to 

the solvent vapors (Figure 1). The open circuit voltages (Voc) of these samples were monitored 

continuously during the test. Solar cell assemblies that demonstrated a significant and permanent 

drop in Voc were removed from the test apparatus, and examined by scanning electron microscopy 

(SEM). I-V curves were also obtained for these degraded samples using an X-25 solar simulator as 

the illumination source. In all cases, I-V curves and SEM photos were obtained for the samples 

prior to testing, so that these could be used for comparison with the post-test results. All initial I-V 

curves showed good fill factors, and no dendrites were observed on any cell edges prior to testing. 

RESULTS AND DISCUSSION 

Open circuit voltage data as a function of time for the p/n dual junction Tecstar cascade cell 

assemblies with soldered Z-bend interconnects are shown in Figures 2a-c. These solar cell 

assemblies are the same design as those on the degraded solar panel array, with slight process 

differences. Five samples were initially placed under test (SN 112, 034, 053, 063, 008). All of these 

NASA/CP—2005-213431 117
assemblies were exposed to the vapor of a 50:50 mixture of water and ketoxime during illumination. 

One of these five samples (SN 034) was also treated with additional solder flux, applied to the 

underside of the crossbar of the Z-bend interconnect before testing. This solder flux-treated sample 

and one of the untreated solar cell assemblies (SN 112) exhibited a rapid and significant Voc

degradation within 24 hours of being placed under test (Figure 2a), after which, the Voc values 

remained relatively stable. After 12 days on test, these solar cell assemblies were removed for I-V 

measurements and SEM analysis. The final Voc value for the solder flux-treated sample was about 

1.3 volts and the non-treated sample had a final Voc of approximately 0.9 volts. The I-V curves for 

these two degraded samples are shown in Figure 3. Both solar cell assemblies exhibited a Voc

recovery upon illumination with the X-25 solar simulator. However, the fill factors for both of the 

samples were degraded compared to pre-test conditions. SEM analysis of the solder flux-treated 

sample (SN 034) shows extensive dendrite growth along the cell edge at what appears to be the 

cell junction region, as shown in Figure 4a. The untreated sample (SN 112) shows large silver 

crystals apparently growing upwards from the top edge of the cell towards the overhanging 

interconnect crossbar (Figure 4b). There are also silver crystals growing downwards from the 

crossbar. It is possible that these crystals once formed a continuous bridge between the 

interconnect crossbar and the cell edge and served as electrical shunts, but that these were broken 

prior to SEM analysis. Since the SEM pictures were obtained after the I-V measurements were 

made, they also reflect the solar cell assembly conditions following Voc recovery. A recovery in Voc

suggests that some shunts were destroyed during the I-V measurement, possibly by current 

overloading of the shunts produced by the greater light intensity from the X-25 simulator. 

Following removal of these two degraded samples from testing, two more interconnected 

solar cell assemblies were added to the illumination test in their place. One of these was another 

p/n dual junction cascade cell with soldered Z-bend interconnects (SN 069), which was placed into 

a Petri dish with no added water or ketoxime, to serve as a DJ cascade control. The second sample 

added was a single junction p/n GaAs/Ge cell (Tecstar) with soldered Z-bend interconnect with no 

crossbar (SN T-50), which was also placed in a Petri dish without added solvent. This sample was 

intended to serve as a SJ control. The DJ "control" sample began to show evidence of shunting 

within 24 hours on test, as indicated by a gradual and "noisy" loss in Voc (Figure 2b). The Voc of this 

cascade cell assembly appeared to stabilize at about 1.5 volts, and then partially recovered to a 

value of 2 volts after 22 days on test (Figure 2c). Three days later, the Voc of the cascade "control" 

assembly dropped again, this time to about 1 volt, and eventually stabilized at approximately 0.9v. 

These data demonstrate that shunt formation can occur in p/n cascade cell assemblies, even in the 

absence of high humidity or other solvent vapors, and that these shunts can occasionally be 

spontaneously reversed, resulting in partial cell recovery. However, degradation eventually results in 

cell failure. 

After three months of continuous illumination, two of the remaining three original p/n 

cascade solar cell assemblies had severely degraded (SN 063 and SN 008) and the last cascade 

assembly (SN 053) had begun to show slight signs of Voc degradation (which may be attributable to 

halogen lamp aging). In contrast, the single junction cell assembly remained stable throughout. The 

degraded Voc values for all of the p/n cascade solar cell assemblies were found to stabilize at one of 

two values: Voc ~ 1.3-1.5v or Voc ~0.9v. These Voc values are consistent with the loss of either one of 

the two junctions in these solar cells. The degraded p/n DJ cascade samples and the SJ control 

sample remaining on test were removed from the test apparatus and their I-V curves were 

measured under illumination with the X-25 simulator. In order to determine whether or not the high 

current generated by the X-25 was destroying some of the dendritic shunts in the degraded solar 

cell assemblies, Voc values were monitored as a function of time as the X-25 shutter was opened. 

The results of these measurements are shown in Figure 5. For each of the severely degraded 

cascade samples, the Voc value is seen to initially rise to its last test value, and then continue to rise 

as the shutter is opened. These observations are consistent with the hypothesis that high current 

destroys some of the dendritic shunts. In contrast, the Voc for the stable SJ cell assembly 

immediately rises to its final Voc test value and remains stable. The one cascade cell assembly that 

did not exhibit severe degradation (SN 053) also rose rapidly to its final test Voc value. The I-V 

NASA/CP—2005-213431 118
curves for the degraded cascade samples (SN 069, SN 063, SN 008) were found to be similar to 

those observed for SN 112 and SN 034. The fill factors for each of the degraded p/n cascade cell 

assemblies are significantly reduced compared to their pre-test values, and the depressed "knee" in 

the I-V curve is consistent with shunting in these cells. A representative I-V curve (SN 008) is 

shown in Figure 6. 

SEM pictures of the degraded p/n cascade cells SN 069, SN 063, and SN 008 are shown in 

Figure 7a-c. In all cases, silver dendrites and/or silver crystallites were found to have formed along 

the cell edges and between the cell and the overhanging crossbar of the Z-bend interconnect, which 

supports they hypothesis that cell degradation is due to shunting via these new growths. 

Three n/p dual junction solar cell assemblies with soldered Z-bend interconnects (same 

overall configuration as the p/n Tecstar cascade assemblies, but with a reverse cell polarity) were 

then placed in the Aerospace test apparatus under similar test conditions as the p/n cascade 

samples. Two solar cell assemblies were suspended in Petri dishes over 50:50 mixtures of water 

and ketoxime, and one assembly was suspended in room air only. The SJ control sample remained 

on test for comparison. After two months on test, none of the n/p DJ solar cell assemblies showed 

any signs of degradation. These data suggest that reversing the cell polarity mitigates the problems 

of dendrite growth and subsequent shunting, possibly by transporting any silver ions that are formed 

away from the solar cell junction region, instead of towards it. 

The last set of interconnected solar cell samples to be subjected to the Aerospace 

illumination testing consisted of three p/n dual junction cells with a welded out-of-plane stress relief 

interconnects. These assemblies were exposed to similar test conditions as the previous n/p DJ 

samples. One solar cell assembly was exposed to water vapor, one was exposed to ketoxime 

vapor, and one was exposed only to room air. Surprisingly, this last set of samples did not exhibit 

any Voc degradation after two months on test, despite the apparent undesirable polarity of the solar 

cells. This lack of degradation might be accounted for by the difference in interconnect design (no 

crossbar present, interconnect tabs are further away from the cell edge, and the silver weld pads 

are smaller than the Z-bend solder pads), or it might be explained by the lack of solder flux (which 

can serve as an electrolyte for silver ion migration). In an effort to differentiate between these two 

effects, solder flux was added to the backside of the welded interconnects on two of these solar cell 

assemblies (those exposed to excess solvent vapor), and all three samples were placed back on 

test. After more than 3 weeks of illumination testing following this change, no degradation has been 

observed. It may be that more time is required before an effect of the solder flux is seen, or it may 

be that the interconnect design prevents silver migration and subsequent degradation. These 

experiments are still ongoing. 

Finally, it is significant that the single junction p/n "control" sample did not show any signs of 

degradation, even after more than seven months on test, and SEM analysis shows no evidence of 

dendrite growth. This lack of degradation in the single junction solar cell assembly may be due to 

the low operating voltage of the SJ cell, but the use of a Z-bend interconnect without a crossbar may 

also contribute to the robustness of this assembly. 

SUMMARY 

The results of the Aerospace illumination experiments indicate that cell polarity and 

interconnect design play significant roles in controlling multi-junction cell dendrite growth and 

subsequent degradation in solar cell assemblies with silver-coated interconnects. The use of an 

out-of-plane stress relief interconnect appears to minimize dendrite formation, as does using an n/p 

cell polarity cell design. These results should aid in the design of robust multi-junction solar cell 

assemblies for future space applications. 

NASA/CP—2005-213431 119
Figure 1. Experimental apparatus for illumination testing 

2a) Five new p/n cascade assemblies placed on test (SN 112, SN 034, SN 053, SN 063, SN 008) 

NASA/CP—2005-213431 120
2b) One new cascade cell assembly (SN 069) and one single junction cell assembly (SN T-50) 

replace SN 112 and SN 034 on test. SN 069 begins to show immediate, noisy degradation; SN 063 

degrades abruptly after 25 days on test.

2c) SN 069 shows partial recovery followed by degradation; SN 008 gradually degrades

 

Figure 2. Open circuit voltage as a function of illumination time for p/n dual junction and single 

junction solar cell assemblies with soldered, Z-bend interconnects. 

NASA/CP—2005-213431 121
Figure 3. I-V curves for degraded p/n dual junction solar cell assemblies with soldered Z-bend 

interconnects. a) SN 034 (with added solder flux), and b) SN 112 

 

P/N Cascade Cell SN 034 Light IV 

-0.05

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.0 0.5 1.0 1.5 2.0 2.5

Voltage (Volts)

Pre-Experiment (FF = 0.770) 

After chemical exposure & SEM 

(FF = 0.460) 

a)

P/N Cascade Cell SN 112 Light IV 

-0.05

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.0 0.5 1.0 1.5 2.0 2.5

Voltage (Volts)

Pre-Experiment (FF = 0.760) 

10 Days of Exposure (FF = 0.670) 

After SEM 

b)

NASA/CP—2005-213431 122
Figure 4. SEM images of cell edges for degraded p/n dual junction solar cell assemblies, located 

under Z-bend crossbar. a) Cell edge of SN 034 (added solder flux), and b) Cell edge of SN 112 

Crossbar 

underside 

Ag dendrites 

GaAs/Ge 

interface 

a)

b)

Crossbar 

underside 

Ag crystallite 

GaAs/Ge 

interface

NASA/CP—2005-213431 123
Voc of Dendritic Cells As X25 Shutter Opens

0.00

0.50

1.00

1.50

2.00

2.50

3.00

0 20 40 60 80 100 120 140 160

Time ( x .04 seconds)

cell 008

cell GaAs

cell 063

cell 053

cell 069

Figure 5. Open circuit voltage of degraded and non-degraded solar cell assemblies as X25 solar 

simulator shutter opens

Figure 6. I-V curve for degraded p/n cascade cell SN 008 with soldered Z-bend interconnects, after 

several months on test. 

2J Cascade Cell 0563-008 Light IV

0.000

0.025

0.050

0.075

0.100

0.125

0.150

0.175

0.200

0.225

0.250

0.275

0.300

0.325

0.00 0.50 1.00 1.50 2.00 2.50

Voltage (volts)

Pre-experiment (FF = 0.800)

3 months of exposure (FF = 0.600)

NASA/CP—2005-213431 124
 

 

Figure 7. SEM images of cell edges for degraded p/n cascade solar cell assemblies, located under 

Z-bend crossbar a) SN 069 (DJ "control" sample, b) SN 063, and c) SN 008 

 

1

 Michalkiewicz R, Green J, and Opperhauser S. “Surface Insulation Resistance Testing of 

Soldering Pastes and Fluxes,” Proceedings of the Pan Pacific Microelectroinics Symposium. Kauai 

Hawaii. February 2001. 

Crossbar 

underside 

GaAs/Ge 

interface 

GaAs/Ge 

interface 

Crossbar 

underside 

GaAs/Ge interface

Ag 

crystallite

Ag dendrites

Ag dendrites

a) b)

c)

NASA/CP—2005-213431 125
INDIUM PHOSPHIDE WINDOW LAYERS FOR INDIUM GALLIUM ARSENIDE SOLAR CELLS 

Raj K. Jain* 

NASA Lewis (now Glenn) Research Center, Cleveland, OH 44135 

Window layers help in reducing the surface recombination at the emitter surface of the solar cells resulting in significant improvement in energy conversion efficiency. Indium gallium arsenide (InxGa1-xAs) 

and related materials based solar cells are quite promising for photovoltaic and thermophotovoltaic 

applications. The flexibility of the change in the bandgap energy and the growth of InGaAs on different 

substrates make this material very attractive for multi-bandgap energy, multi-junction solar cell approaches. The high efficiency and better radiation performance of the solar cell structures based on 

InGaAs make them suitable for space power applications. This work investigates the suitability of indium phosphide (InP) window layers for lattice-matched In0.53Ga0.47As (bandgap energy 0.74 eV) solar 

cells. We present the first data on the effects of the p-type InP window layer on p-on-n lattice-matched 

InGaAs solar cells. The modeled quantum efficiency results show a significant improvement in the 

blue region with the InP window. The bare InGaAs solar cell performance suffers due to high surface 

recombination velocity (107

 cm/s). The large band discontinuity at the InP/InGaAs heterojunction offers a great potential barrier to minority carriers. The calculated results demonstrate that the InP window layer effectively passivates the solar cell front surface, hence resulting in reduced surface recombination and therefore, significantly improving the performance of the InGaAs solar cell. 

INTRODUCTION

Window layers are quite important in improving the solar cell energy conversion efficiency. They help 

in effectively reducing the surface recombination at the emitter surface of the solar cell without absorbing the useful light required for the device. Unlike silicon, solar cells based on III-V compound semiconductors and related materials suffer from the lack of native passivating oxides and/or wide choice 

of suitable large bandgap energy window materials. Various window layer materials have been investigated for III-V compound semiconductor based solar cells [1]. Further optimization work on the existing choices of windows and more development work on new window materials are required. It is also 

important to understand the physics of the window-emitter hetero-interface as well as growth related 

issues. 

Indium gallium arsenide (InxGa1-xAs) material has been used widely in several state-of-the-art electronic and optoelectronic devices. InxGa1-xAs and related materials based solar cells are quite promising for photovoltaic (PV) [2] and thermophotovoltaic (TPV) applications [3]. InxGa1-xAs is a ternary 

semiconductor whose bandgap energy can be varied between 1.42 eV (of GaAs) and 0.36 eV (of 

InAs) by changing the In/Ga ratio (see Fig. 1). The flexibility of the change in the bandgap energy and 

the growth of InGaAs on different substrates (Si, Ge, GaAs, InP, InAs) make this material very attractive for various devices. InGaAs solar cells grown on gallium arsenide (GaAs) substrates were developed twenty years ago [4]. The 1.15 eV and 1.35 eV bandgap energy InGaAs on GaAs 

cells 

_______________________ 

*Presently employed at Space Systems/Loral, 3825 Fabian Way, M/S Z-51, Palo Alto, CA 95014 

NASA/CP—2005-213431 126
demonstrated efficiencies in excess of 24% at 400-sun concentration [5]. These developments were 

intended for the use of InGaAs cells in multi-bandgap energy, multi-junction solar cell approaches. 

Two-junction monolithic InP/InGaAs tandem cells grown on InP substrates have demonstrated AM0 1-

sun efficiencies over 22% [6]. 

Figure 1. Bandgap energy of indium gallium arsenide and related materials as a function of lattice constant. 

Record efficiency n-on-p and p-on-n lattice-matched InGaAs (0.74 eV bandgap energy) solar cells 

grown on InP were reported [7,8]. Lower bandgap energy (0.74 eV and below) InGaAs cells grown on 

InP substrates offer high efficiencies [9] for thermophotovoltaic applications and have been developed 

[8]. InGaAs based solar cells also find use in quantum well, operation under monochromatic light, and 

related applications. The demonstration of high PV and TPV efficiencies and better radiation performance [10] of the solar cell structures based on InGaAs solar cells make them attractive for space 

power applications. InGaAs and related materials based solar cells grown on GaAs substrates for applications in tandem cells are getting renewed interest and are under investigation. The InGaAs based 

cell technologies on GaAs substrates could easily be adopted later on germanium (Ge) substrates [2], 

similar to high efficiency GaInP/GaAs/Ge dual or triple-junction cells grown on Ge substrates. 

InGaAs solar cells have been fabricated with or without window layers but no quantitative theoretical or 

experimental data on the effects of window materials are available in the literature. This work investigates the suitability of indium phosphide (InP) window layers for lattice-matched In0.53Ga0.47As (bandgap 

energy 0.74 eV) solar cells. We present the data on the effects of the p-type InP window layer on pon-n lattice-matched InGaAs solar cells. p-on-n structure was selected for further investigation and 

development as it offered better optical and electrical properties than n-on-p structure. The calculated 

quantum efficiency results demonstrate the suitability of the InP window in effectively reducing the surface recombination at the front surface of the InGaAs cell resulting in significant improvements in the 

blue response. The large band discontinuities at the heterojunction provide a great potential barrier for 

the minority carriers. PC1D computer code [11] was used in performing the numerical calculations. 

NASA/CP—2005-213431 127
MODELING APPROACH 

InGaAs Material Parameters 

The InGaAs solar cell modeling requires an accurate knowledge about the various physical parameters. Little or no experimental information is available about the various physical parameters of latticematched or mismatched InGaAs material. The optical absorption coefficient and the intrinsic carrier 

concentration are very important parameters, which greatly influence the solar cell current and voltage 

respectively. Jain et al. [12,9] provided the first estimated and calculated data on these two important 

parameters based on available information. Figure 2 shows the optical absorption coefficient versus 

wavelength for the lattice-matched 0.74 eV bandgap energy InGaAs at 300 K. 

Figure 2. Optical absorption coefficient of lattice-matched InGaAs (0.74 eV bandgap energy) versus 

wavelength at 300 K. 

The intrinsic carrier concentration, ni

 for InGaAs was calculated by using the scaling relation described 

by Jain et al. [12], 

 ni

 (InGaAs) = ni (GaAs) exp{-•EG/(2kT)} (1) 

NASA/CP—2005-213431 128
where •EG is the bandgap difference, k is the Boltzman’s constant, and T is the temperature in K. Figure 3 shows the plot of the intrinsic carrier concentration vs. InGaAs bandgap energy at 300 K. An ni

value of 2.59x106

 cm-3 for GaAs at 300 K was used. As the bandgap energy of InGaAs decreases 

(with increase in In/Ga ratio), the ni value increases and correspondingly the open-circuit voltage of the 

InGaAs solar cell decreases. The absorption coefficient versus wavelength values (Fig. 2) and the ni

value at 0.74 eV (Fig. 3) were used in modeling the performance of the lattice-matched InGaAs solar 

cell. Similarly, the various physical parameters for the InP window material were used from the literature. 

Figure 3. Plot of the intrinsic carrier concentration versus InGaAs bandgap energy at 300 K. 

InP/InGaAs Interface Band Structure 

InP forms a heterojunction with InGaAs and the InP/InGaAs interface plays an important role in controlling the performance of the devices based on these materials. The bandgap energy difference between the two materials causes the mis-alignment of the conduction and valence bands resulting into 

band discontinuities. The band offsets in InP/InGaAs heterojunction structures have been studied experimentally and theoretically. Reference 13 describes the various values for the conduction band and 

the valence band discontinuities (•EC and •EV) measured and calculated by several research groups. 

NASA/CP—2005-213431 129
We have used average experimental values of 0.23 eV and 0.37 eV for •EC and •EV respectively in this 

work. Figure 4 shows our calculated band structure diagram for the p-InP/p-InGaAs heterojunction. 

These results have been discussed in detail in the next section. 

Figure 4. Band structure diagram of p-InP/p-In0.53Ga0.47As (window/emitter) heterojunction interface. 

RESULTS AND DISCUSSION 

For the past several years, we have been using a numerical code PC1D for silicon and III-V compound 

semiconductor based (GaAs, InP, InGaAs, GaInP, etc.) solar cell modeling and simulation of the 

measured cell performance to understand and provide feedback for further improvements to produce 

state-of-the-art devices. This work has resulted in several publications and a US patent [14]. NASA 

Lewis (now Glenn) Research Center has been continuing the pioneering work on the development of 

InGaAs based solar cells [15-17] and has reported record cell efficiencies [7,8]. 

In this work we have considered a p-on-n lattice-matched InGaAs solar cell grown on InP substrate. 

Figure 5 shows the cell structure considered. The emitter and base thickness and doping concentration were 0.3 μm, 2x1018 cm-3 and 4 μm, 1x1017 cm-3 respectively. The p InP window layer doping concentration was 2x1018 cm-3, same as in the emitter. No antireflection coatings were considered to 

separate the effects of the window layer on the performance of the InGaAs solar cell. Minority carrier 

diffusion lengths of 0.5, 1, and 4 μm have been assumed in the window, emitter and base layers. The 

front, interface and the back surface recombination velocities were assumed to be 107

, 107, and 104

cm/s. These values are typical of the current state-of-the art in p-on-n InGaAs solar cells. The heterojunction band structure parameters were used in modeling calculations. 

NASA/CP—2005-213431 130
Figure 5. Structure of an InP/In0.53Ga0.47As p-on-n solar cell. 

Figure 6 shows the calculated internal quantum efficiency response for the InGaAs solar cell with and 

without InP window layer. A 100 nm thick window layer was considered. Calculations 

were 

Figure 6. Plot of the internal quantum efficiency versus wavelength of p-on-n InGaAs solar cell with 

and without InP window layer. 

NASA/CP—2005-213431 131
performed by varying the window layer thickness and no significant difference in results was observed. 

This suggests that the wide-bandgap 1.35 eV InP window layer material does not absorb the light useful to InGaAs solar cell. There is a significant improvement in internal quantum efficiency, especially in 

the blue region, with window layer as shown in Figure 6. These results can be explained by the large 

conduction band discontinuity as shown in Figure 4. This discontinuity of 0.61 eV {= •EC (0.23 eV) + 

Band bending (0.38 eV)} acts as an effective potential barrier for the minority carriers (electrons) of the 

p-on-n InGaAs cell, thereby stopping them from recombining at the surface. We have also calculated 

the effect of InP as a back surface passivation layer for InGaAs solar cells and found that the red response increases significantly. Detailed results will be published elsewhere. 

CONCLUSIONS 

The modeling results demonstrate the suitability of InP window layer material for lattice-matched InGaAs p-on-n solar cells. The internal quantum efficiency response significantly improves with window 

layer. These results can be explained by the large conduction band offset which acts as a potential 

barrier to minority carriers. Hence, InP reduces the surface recombination and effectively passivates 

the InGaAs surface. InGaAs based solar cells have been developed in single and multijunction structures for space photovoltaic and thermophotovoltaic applications. The current work and future developments should benefit from the results reported in this paper. 

Acknowledgments 

This work was done when the author was working at the NASA Lewis (now Glenn) Research Center, 

Cleveland, Ohio and supported by National Research Council-NASA Research Associate Program. 

Helpful discussions with David M. Wilt, Dr. Geoffrey Landis, Dr. Dennis J. Flood, and several other 

NASA colleagues are greatly appreciated. 

REFERENCES 

[1] J. Lammasniemi, R. K. Jain, M. Pessa, “Status of Window Layers for III-V Semiconductor Cells”, 

Proceedings 14th European Photovoltaic Solar Energy Conference (1997) p. 1767. 

[2] R. R. King, C. M. Fetzer, P. C. Colter, K. M. Edmondson, J. H. Ermer, H. L. Cotal, H. Yoon, A. P. 

Stavrides, G. Kinsey, D. D. Krut and N. H. Karam, “High-Efficiency Space and Terrestrial Multijunction 

Solar Cells Through Bandgap Control in Cell Structures”, Proceedings of the 29th Photovoltaic Specialists Conference, (IEEE, New York, 2002) p. 776. 

[3] R. J. Wehrer, M. W. Wanlass, B. Wernsman, J. J. Carapella, S. P. Ahrenkiel, D. M. Wilt and C. S. 

Murray, “0.74/0.55-eV GaxIn1-xAs/InAsyP1-y Monolithic, Tandem, MIM TPV Converters: Design, 

Growth, Processing and Performance”, Proceedings of the 29th Photovoltaic Specialists Conference, 

(IEEE, New York, 2002) p. 884. 

[4] M. J. Ludowise, W. T. Dietze, R. Boettcher and N. Kaminar, “High-Efficiency (21.4%) 

Ga0.75In0.26As/GaAs (Eg=1.15 eV) Concentrator Solar Cells and the Influence of Lattice Mismatched 

on Performance”, Appl. Phys. Lett., Vol. 43, p. 468 (1983). 

[5] J. G. Werthen, B. A. Arau, C. W. Ford, N. R. Kaminar, M. S. Kuryla, M. L. Ristow, C. R. Lewis, H. 

F. MacMillan, G. F. Virshup and J. M. Jee, “Recent Advances in High Efficiency InGaAs Concentrator 

NASA/CP—2005-213431 132
Cells”, Proceedings of the 20th IEEE Photovoltaic Specialists Conference (IEEE, New York, 1988) p. 

640. 

[6] M. W. Wanlass, J. S. Ward, K. A. Emery, A. Duda, T. J. Coutts, Proceedings First World Conference on Photovoltaic Energy Conversion ( IEEE, New York, 1994) p. 1717. 

[7] D. M. Wilt, N. S. Fatemi, R. W. Hoffman, P. P. Jenkins, D. J. Brinker, D. Scheiman, R. Lowe, R. K. 

Jain, “High Efficiency Indium Gallium Arsenide Photovoltaic Devices for Thermophotovoltaic Power 

Systems”, Appl. Phys. Lett., Vol. 64, p. 2415 (1994). 

[8] D. M. Wilt, N. S. Fatemi, P. P. Jenkins, V. G. Weizer, R. W. Hoffman, R. K. Jain, C. S. Murray, D. 

R. Riley, “Electrical and Optical Performance Characteristics of 0.74 eV p/n InGaAs Monolithic Interconnected Modules”, Proceedings 3rd NREL Conference on Thermophotovoltaic Generation of Electricity (American Institute of Physics 1997) p. 237. 

[9] R. K. Jain, D. M. Wilt, R. Jain, G. A. Landis, D. J. Flood, “Lattice-Matched and Strained InGaAs 

Solar Cells for Thermophotovoltaic Use”, Proceedings 2nd NREL Conference on Thermophotovoltaic 

Generation of Electricity (American Institute of Physics 1995) p. 375. 

[10] I. Weinberg, R. K. Jain, C. Vargas, D. M. Wilt, D. A. Scheiman, Performance of Ga0.47In0.53As 

Solar Cells Over a Range of Proton Energies”, Proceedings First World Conference on Photovoltaic 

Energy Conversion ( IEEE, New York 1994) p. 2173. 

[11] P. A. Basore, “PC-1D Version 3: Improved Speed and Convergence”, Proceedings of the 22nd 

Photovoltaic Specialists Conference (IEEE, New York, 1991) p. 299. 

[12] R. K. Jain, D. M. Wilt, G. A. Landis, R. Jain, I. Weinberg, D. J. Flood, Modeling of Low Bandgap 

Solar Cells for Thermophotovoltaic Applications”, Proceedings 1st NREL Conference on Thermophotovoltaic Generation of Electricity (American Institute of Physics 1994) p. 202. 

[13] Properties of Lattice-Matched and Strained Indium Gallium Arsenide, INSPEC (1993) p. 86. 

[14] R. K. Jain and G. A. Landis, “InP Solar Cell with Window Layer”, US Patent Number 5,322,573, 

June 21, 1994 

[15] D. M. Wilt, et al., “Lattice-Mismatched InGaAs PV Devices for TPV Power Systems”, Proceedings 

First World Conference on Photovoltaic Energy Conversion ( IEEE, New York, 1994) p. 1738. 

[16] D. M. Wilt, et al., “Monolithically Interconnected InGaAs TPV Module Development”, Proceedings 

of the 25th Photovoltaic Specialists Conference, (IEEE, New York, 1996) p. 43. 

[17] D. M. Wilt, et al., “0.55 eV n/p/n MIM TPV Cell Development”, Proceedings of the 28th Photovoltaic 

Specialists Conference, (IEEE, New York, 2000) p. 1024. 

NASA/CP—2005-213431 133
DESIGN AND PERFORMANCE OF A TRIPLE SOURCE AIR MASS ZERO SOLAR SIMULATOR

Phillip Jenkins

Ohio Aerospace Institute, Brook Park, OH 44142

David Scheiman

Ohio Aerospace Institute, Brook Park, OH 44142

David Snyder

NASA Glenn Research Center, Cleveland, OH 44135

Simulating the sun in a laboratory for the purpose of measuring solar cells has long been a challenge for

engineers and scientists.1,2,3 Multi-junction cells demand higher fidelity of a solar simulator than do single junction

cells, due to a need for close spectral matching as well as AM0 intensity. A GaInP/GaAs/Ge solar cell for

example, requires spectral matching in three distinct spectral bands (figure 1). A commercial single source highpressure xenon arc solar simulator such as the Spectrolab X-25 at NASA Glenn Research Center, can match the

top two junctions of a GaInP/GaAs/Ge cell to within 1.3% mismatch, with the GaAs cell receiving slightly more

current than required. The Ge bottom cell however, is mismatched +8.8%. Multi source simulators are designed

to match the current for all junctions but typically have small illuminated areas, less uniformity and less beam

collimation compared to an X-25 simulator. It was our intent when designing a multi source simulator to preserve

as many aspects of the X-25 while adding multi-source capability.

The X-25 faithfully reproduces the AM0 spectrum from approximately 350-nm to 750-nm but beyond that region

the spectrum quickly diverges from AM0 (figure 2). This multi-source simulator takes advantage of this property

and uses the X-25 to provide the bulk of the UV and visible light. A non-absorbing dichroic mirror (designed and

fabricated by Displays and Optical Technologies, of Round Rock, Texas) is used as a beam splitter to reflect the

blue portion of the spectrum and transmit the infrared. The transmission and reflection curves of the filter (figure

3) show that there is a fairly good cutoff transition at 700-nm but that the reflection band pass is limited to

approximately one octave and begins to transmit excessively at wavelengths shorter than 450nm. This behavior

0

0.25

0.5

0.75

1

1.25

1.5

1.75

2

350 500 650 800 950 1100 1250 1400 1550 1700 1850 2000

Wavelength (nanometers)

mW/cm^2-nm

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

0.80

0.90

1.00

Solar Cell Quantum Efficiency

GaInP

GaAs

Germanium

AM0 Power 

Density

350nm-700nm

700nm

to

900nm

900nm-1700nm

Q.E.

0

0.25

0.5

0.75

1

1.25

1.5

1.75

2

2.25

2.5

2.75

3

3.25

3.5

350 550 750 950 1150 1350 1550 1750 1950

Wavelength (nanometers)

mW/cm^2-nm

AM0

X-25 Spectrum 

Figure 1) The quantum efficiency of a GaInP/

GaAs/Ge solar cell plotted with the AM0 spectrum. Figure 2) Spectrum of the Spectrolab X-25 solar

simulator5 compared to AM0.

NASA/CP—2005-213431 134
is typical for high efficiency, thin-film dielectric stack filters. Although not ideal, the filtered beam provides sufficient

UV and visible light for solar cells and excellent IR transmission for adding in the balance of the AM0 spectrum.

With the cell test plane and the X-25 simulator mounted 90° to each other, the dichroic mirror is mounted at a 45°

angle to reflect the blue portion of the X-25 beam onto the cell and allow for additional lights sources from above

the cell plane (figure 4).

The initial design strategy for the IR source was to pack as many lamps into as small an area as possible and

mount the lamps as far from the test plane as practical. This philosophy (in principle) will yield the best

collimation angle and beam uniformity. The lamps used are tungsten filament bulbs mounted in a gold reflector.

The lamps operate at 12 Volts with an output of 75 Watts and have an average life expectancy of 4,000 hours.

The vendor is Gilway Technical Lamp in Woburn, MA. The reflector has a nominal beam angle of 14°. The

attempt to pack as many lamps into as small a space as possible, was a rectangular array of 40 lamps. Although

a round lamp should pack more tightly in a hexagonal pattern, the lamp holders have a rectangular base and thus

the best packing is achieved using a rectangular pattern. This lamp arrangement had one major shortcoming; it

required that each lamp have two degrees of

pointing control in order to achieve acceptable

uniformity. Without this control, the beam had a

significant center hot spot. A second design,

which capitalizes on the natural radial symmetry of

the beam pattern, was designed. Here, the lamps

required only one degree of freedom of pointing to

achieve acceptable beam uniformity. In this

design, 36 lamps are oriented in a radial pattern

and each lamp can be pointed along a line that

sweeps from center to edge along a radius of the

beam. Pointing of each lamp is achieved by

mounting the lamp holder to a aluminum block

with four spring-loaded screws. The blocks are

fastened normal to the lamp mounting plate. The

four screws can be adjusted to tilt the lamp

housing approximately ±5°. This is enough

movement to sweep the beam from center to the

outside edge of the test plane. Figure 5 shows the

mounting arrangement for a single lamp. In

practice, all the lamps are splayed away from the

center to decrease the center beam hot spot.

X-25 Simulator

IR source

Unwanted red 

portion of X-25 

spectrum

Solar cell test 

plane

Dichroic mirror

Figure 3) Transmission and reflection of dichroic mirror. Figure 4) Optical path for combining the IR

source with the X-25 beam.

Figure 5) Each IR lamp is mounted on an aluminum block

with 4 spring-loaded adjustment screws.

NASA/CP—2005-213431 135
The spectrum of an unfiltered tungsten lamp does not adequately simulate the IR portion of AM0. Figure 6 shows

the spectrum of an unfiltered tungsten lamp versus AM0 in the infrared. Note that below 1000-nm the power

density of the tungsten lamp drops compared to AM0. This is a critical region of the spectrum for the GaAs

junction of a GaAInP/GaAs/Ge cell. When the intensity of the tungsten lamps is raised to match the GaAs middle

cell current, then the Ge cell has 40% excess current! Another filter was designed to shape the tungsten lamps

to match AM0. Figure 7 shows the (calculated) transmission of this filter. The filter is placed between the

tungsten lamps and the dichroic mirror. The resulting spectrum is close to AM0, but still needs an additional

degree of freedom to perfectly match all three junctions.

A third light source was created by filtering a subset of the 36 tungsten lamps. A non-absorbing filter was

designed to transmit only in the region of 700-900-nm. This filtration is optimized for adjusting the current of a

GaAs middle cell separate from the germanium or GaInP cell. The filters are placed in front of 6-12 lamps

depending on the amount of spectral mismatch of a particular cell (figure 8). The filtered lamps are put on a

separate power supply and their intensity is controlled independently of the 24 remaining lamps. For the 12 700-

900-nm filtered lamps, the lamps are connected as 4 parallel strings of 3 lamps each for a nominal operating

voltage of 36 Vdc at 25 Amps. The remaining 24 lamps are wired as 4 parallel strings of 6 lamps, and operate at

72 Vdc at 25 Amps. Using all three light sources, the short circuit current of each reference cell of a

GaInP/GaAs/Ge multijunction cell, can be simultaneously matched to their reference values to better than 0.5% in

less than 30 minutes set up time. The spectrum of the simulator is shown in figure 9.

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

300 500 700 900 1100 1300 1500 1700 1900

Wavelength (nm)

Transmission

Figure 6) Spectrum of the tungsten lamp compared to AM0.

Figure 7) Design of the filter to shape the IR lamp

spectrum to match AM0.

0

0.25

0.5

0.75

1

1.25

1.5

1.75

2

2.25

2.5

350 600 850 1100 1350 1600 1850

Wavelength (nm)

mW/cm^2-nm

Solar Simulator

AM0

Figure 9) Spectrum of the solar simulator when all three

junctions of a GaInP/GaAs/Ge solar cell are properly illuminate,

compared to AM0.

Figure 8) Photograph showing the 12 filtered

lamps of the 36 lamp IR source.

NASA/CP—2005-213431 136
The collimation half-angle is determined by the angle formed by the radius of the IR lamp cluster and the distance

the lamps are located above the test plane. This angle is 7.6 degrees. The uniformity of light intensity in the test

plane is most strongly affected by the number of individual lamps that can be articulated to reduce “hot or cold”

spots in the beam. All of the lamps contribute to the 700-900-nm portion of the spectrum and therefore the

number of lamps is effectively 36 in the 700-900-nm region. The intensity changes made with the 700-900-nm

filtered lamps are small (<5%) compared to the overall intensity in the 700-900-nm range and have little affect on

the uniformity. In the long IR region (900-1700-nm) the number of lamps is effectively 24. The intensity uniformity

of these lamps is ±1% over a large area of a 20-cm X 20-cm test plane as shown in figure 10. The uniformity of

the X-25 is similar to the IR lamps and is shown in figure 11. The intensity maps were made using an X-Y plotter

fitted with a photodiode. The method is similar to that used by Spectrolab to characterize the X-25 (X-25

documentation) and at other labs such as National Renewable Energy Laboratory (NREL). The stability of the

lamps has not been quantitatively characterized at this time but after an initial 30 minute warm-up period there is

no perceptible drift while making measurements over a 4 hour period. The IR lamps are also flicker free, due to

the nature of the IR lamp optics and the DC power used.

A schematic figure of the entire system is shown in figure 12. For many applications the single source X-25 is all

that is required. In such cases, the system can be quickly reconfigured by exchanging the dichroic mirror for a

simple aluminized first surface mirror. This also helped preserve the X-25 capability while the new system was

being built.

The cost of all the hardware (excluding the X-25 simulator) was 25,000 USD (in 2002). Operating the simulator

offers many challenges. There are a number ways to set up the intensity and spectrum of a multi-source

simulator. The easiest method is to have separate calibrated reference cells for each sub junction of the test cell.

If such cells are not available, then one must either apply spectral correction to existing reference cells that are

close in spectral response, or calculate the required spectral intensity from measurements of the spectral

response of the test cell and measurements of the simulator’s spectral intensity.

We have successfully converted a Spectrolab X-25 solar simulator into a multi-source simulator capable of

precisely matching the required currents in a triple junction GaInP/GaAs/Ge solar cell. The system is relatively

inexpensive to impliment and preserves the original functionality of the X-25 as well. It has excellent uniformity as

well as good stability and a collimation angle better than ±8°. Future plans for improving the performance, include

the addition of another xenon light filtered to provide a boost in the blue region of the spectrum to more closely

match AM0.

±1% Uniformity

±2%

±1% Uniformity

±2%

Figure 10) Uniformity of the long IR light source over a

20-cm x 20-cm area. The light blue contour represents

the additional area where the uniformity is at least ±2%.

Figure 11) Uniformity of the X-25 Simulator over a 20-

cm x 20-cm area.

NASA/CP—2005-213431 137
X-25 Solar

Simulator

(3kW Xe

lamp)

36 IR Lamps

6-12 lamps 

with 700nm900nm band 

pass filters

HEPA filter 

purge

IR Shaping Filter

Dichroic Mirror

Test 

Plane

2.1 meters

DC Power

DC Power

References

 1

 Solar Cell Standards Group, Energy Conversion Task Group, West Coast Solid State Devices, “Photovoltaic

Solar Simulator Specification, American Institute of Electrical Engineers, June 7, 1961, Revised September 5,

1961.

2

 D.B. Bickler, “The Hoffman Solar Simulator,” Proceedings of the Solar Working Group Conference (2nd PVSC),

Volume II, 6-1 to 6-21, Held, in Washington D.C., Feb. 27 and 28, 1962.

3

 Joseph H. Apfel and Alfred J. Thelen, “A Solar Simulator,” Proceedings of the Solar Working Group Conference

(2nd PVSC), Volume II, 7-1 to 7-21, PIC-SOL 209/2.1, Held, in Washington D.C., Feb. 27 and 28, 1962.

Figure 12) Schematic representation of the multi-source simulator.

NASA/CP—2005-213431 138
MEASUREMENT AND CHARACTERIZATION OF CONCENTRATOR SOLAR CELLS II

Dave Scheiman1, Bernard L. Sater2, Donald Chubb3, Phillip Jenkins1, Dave Snyder3

 1

OAI, 22800 CedarPoint Rd, Brookpark, OH 44142 2

PhotoVolt, Inc., 21282 Woodview Circle,Strongsville, OH 44149

 3

NASA GRC, 21000 Brookpark Rd, Cleveland, OH 44135

Concentrator solar cells are continuing to get more consideration for use in power systems. This interest is

because concentrator systems can have a net lower cost per watt in solar cell materials plus ongoing

improvements in sun-tracking technology. Quantitatively measuring the efficiency of solar cells under

concentration is difficult. Traditionally, the light concentration on solar cells has been determined by using a ratio

of the measured solar cell’s short circuit current to that at one sun, this assumes that current changes

proportionally with light intensity. This works well with low to moderate (<20 suns) concentration levels on “wellbehaved” linear cells but does not apply when cells respond superlinearly, current increases faster than intensity,

or sublinearly, current increases more slowly than intensity. This paper continues work on using view factors to

determine the concentration level and linearity of the solar cell with mathematical view factor analysis and

experimental results [1].

INTRODUCTION

Solar cells are designed to convert photon energy directly into electrical energy. The amount of natural

sunlight incident on a solar cell is considered the 1 sun power density. Higher than 1 sun power densities

achieved using lens or reflectors are concentrated sunlight. The increase in power density or ratio of power

density to the 1 sun power density is the concentration level. Efficiency measurement of concentrator solar cells

has always been a difficult task due to the lack of equipment available to measure the light power intensity. Most

laboratory measurements are done using a Large Area Pulse Solar Simulator (LAPSS) with data being taken in 1-

2 milliseconds. Measuring the power density/flux under these conditions requires a very fast thermopile or

pyroelectric device, if any exist. Consequently, determining the power density or light concentration ratio on a

solar cell is the ratio of the measured short circuit current (ISC) to the short circuit current at 1 sun. This process

can be inaccurate for solar cells that do not behave linearly with power density.

One way to obtain power density/flux is by using radiative energy transfer equations and calculating the view

factor between the source and the test sample. Energy radiating from a source incident on a surface at some

fixed distance away can be characterized by the view factor between the source and the surface. Knowing the

shape of the source and the incident surface and using numerical integration, the view factor can be calculated.

The view factor has always been used in determining the efficiency of thermophotovoltaic energy conversion

systems [2]. The view factor ratio between a location with a known power density and some other location can be

used to determine the concentration (change in power density) by ratio of the view factors. When dealing with a

point source, the view factor varies as 1/r2 where r is the distance between the source and the surface. The sun is

modeled as a point source because of its distance from the earth, closer objects can’t use this approximation.

THE VIEW FACTOR

The view factor is the fraction of the total energy emitted by one surface, the source, that is directly incident

on another surface. As the source and surface move farther apart or off-axis from each other, the view factor

becomes smaller. A point source with the incident surface forming a full hemisphere surrounding that source has

a view factor of 1 because the total energy from the source is incident on the surface. Additionally, the view factor

between a point source and the target surface decreases proportionally with the square of the distance between

them. Calculating the view factor between two finite shaped surfaces requires numerical integration of both

surfaces. The general form of the view factor calculation is defined by the following equation [3]:

† 

View Factor FAA' = 1

pA' Ú dA cosf' cosf

s

2 dA' Ú

NASA/CP—2005-213431 139
Where A and A’ are the two surfaces, s is

the distance between the two surfaces and ø

are the respective angles between a point on

surface A and a point on surface A’ and vice

versa. Generating and solving this equation for

specific surface geometries is beyond the

scope of this paper.

View Factor Solution for LAPSS

For the case described here, the source is

the Spectrolab LAPSS100, and the incident

surface is a planar solar cell. The source in

the LAPSS is a Xenon arc lamp that generates

a pair of 15 cm. arcs as shown in figure 1. To

simplify the view factor calculation, the light

source is modeled as a rectangle with a finite

length and width, assumed to be uniform and

isotropic. The LAPSS system contains two

identical arc lamps which fire simultaneously.

By placing the solar cell (incident surface)

symmetrically centered between the two lamps

only one view factor is calculated, and the

resulting view factor for both lamps is simply

twice that of the single lamp.

Using this model, a drawing of the test

geometry is made as shown in figure 2. There

are two identical sources, offset from the center

line of the test solar cell. Because of the

symmetry of the test setup, only one lamp

needs to be used to calculate the view factor

which is identical for the second lamp. These

two view factors can be added to get the actual

view factor. The lamp source is surface A1 in

the xy plane, the solar cell surface is A2 in the

zh plane. The distance between the two

planes is z and the distance between points the

two surfaces is s with angles Q1,2 between

them. Based on this model, discrete

coordinates of the rectangular endpoints can

be obtained for the view factor calculation.

The actual view factor calculation is shown in

appendix A, it is derived from two parallel and

perpendicular rectangular finite surfaces. 

 The view factor is calculated for a range of

distances z from the source using the numerical summation from appendix A [4]. By plotting the ratio of the view

factor at the farthest distance to shorter distances a graphical representation of this calculation is compared to a

1/r2 and shown in Figure 3. The concentration level of light on the solar cell is obtained by the following steps:

STEP 1: Measure the solar cell performance under a known or calibrated light source at 1 sun (this can be done

in a solar simulator other than the LAPSS with a standard cell). Light level of the simulator is adjusted

using a calibrated primary or secondary reference cell.

+

+

Trigger Wire (Outside of Lamp)

Trigger Wire 

(Outside of 

Lamp)

End Cap

Hose

 Clamp

Shield

 Wires

Center

 Wires

+

+ +

+

16.4 cm

12.0 cm

Figure 1 LAPSS lamp setup

Figure 2 LAPSS test geometry

NASA/CP—2005-213431 140
STEP 2: Set up a coordinate system for the cell and the LAPSS or concentrating light source, select a distance z

between the two surfaces farthest away as a 1 sun starting point. The coordinate system is based on

figure 2.

STEP 3: Using the coordinates in step 2 calculate a range of View Factors from z and shorter, ratio these view

factors to the view factor at z to get the concentration level. A computer program can be written for this

step.

STEP 4: Set up the solar cell at distance z from the light source(s), measure and adjust the light source to obtain

the 1 sun numbers from Step 1.

STEP 5: Using the data from step 3, move the solar cell closer to the light source to various distances as denoted

in Step 3, these measurements are at the calculated concentration levels.

The data used in figure 3 was for a 2 x 2 cm. solar cell, the light source is the Spectrolab LAPSS and the 1 sun

measurement was made at a distance of 635 cm. The view factor was calculated every 5 cm. going into the

source.

THE ARC LAMP

 The arc lamp is the most optimum single source solar simulator because of its close temperature match to the

solar spectrum. The disadvantages are that the arc lamp’s spectrum has many spikes due to the emission lines

of the gas in the lamp. The Spectrolab LAPSS system was designed to minimize the spikes and maintain good

uniformity at a 2 m x 2 m 1-sun plane within ±2%[5]. In order for the view factor calculation to be accurate, the

light source is assumed to be uniform and isotropic. The uniformity of the lamp can be generated from knowledge

of the plasma characteristics of the arc itself [6]. Arc lamps tend to be cooler near the electrodes but have

relatively constant temperature across the arc. To minimize this temperature effect, the lamps were baffled 1 cm.

above and below the electrodes. As can be seen in the photos, the arc completely fills the interior area of glass

between the electrodes, this is a result of electrode design, and insures ionization of all of the gas, resulting in

uniform light generation. The source is also isotropic, emitting light equally in all directions, however half of the

lamp diameter is not used. Any reflective optics or lenses used on this lamp would enhance the light intensity but

also complicate the view factor calculation.

 High speed photographs of the lamp during a flash are shown in figure 4. As can be seen from these

photographs, the arc fully saturates the glass housing and decays to a thin line. The solar cell performance is

only obtained during the full light portion of the arc with the best uniformity. This full light section of the arc also

has the flattest intensity and can last a few milliseconds. It is during these few milliseconds that the solar cell is

characterized, a monitor cell is also used to adjust for slight intensity variations within the flash and between

consecutive flashes. Typically, intensity between flashes of the arc lamp are controlled to be less than .25%

which is needed for measuring the cell at different positions or concentration levels.

0

250

500

750

1000

1250

1500

1750

2000

0 50 100 150 200 250 300

Distance From Source

Concentration (View Factor Ratio)

View Factor

1/r^2 % Difference Between View Factor and 1/r^2

0.0

20.0

40.0

60.0

80.0

100.0

120.0

140.0

160.0

180.0

0 20 40 60 80 100 120 140 160 180 200

Distance 

% Difference

Figure 3. Concentration by View Factor Ratio

NASA/CP—2005-213431 141
TEST

Several cells were measured at 1 Sun AM0 in a solar simulator using a calibrated reference cell. A

LAPSS100 was used for a light source, assumed to be uniform and isotropic. The LAPSS light intensity was

adjusted for 1 sun intensity at a fixed distance. Using the view factor ratio, the cells’ current-voltage

characteristics were measured at a variety of concentrations. Arc lamps exhibit uniformity in the center portion of

the arc away from the electrodes which tend to be cooler. Therefore the lamps were baffled 1 cm. from either

electrode. No additional optics were used with the lamps which would affect the view factor calculation. Based

on the mathematical analysis defined in section 2, a test was set up. Four different solar cells were selected to

cover a range of performance characteristics; Triple Junction InGaP/GaAs/Ge 1 cm. x 4 cm. from two different

vendors. These cells were designed as space concentrator cells to operate at <20 suns. Each of these cells

were characterized at 1 Sun AM0 using an appropriate reference standard.

All tests were performed at NASA Glenn Research Center. The Spectrolab LAPSS100 system performs IV

curves on solar cells/arrays in 1-2 milliseconds. It has a 12-bit A/D measuring cell voltage, current, and monitor

cell current simultaneously. IV curves consist of from 20 to 100 data points. The monitor cell was placed in a

fixed position within the beam path, its location was selected to monitor flash intensity and not interfere with the

test set up.

Test Setup

Before the test could be performed, several modifications had to be made to the existing LAPSS facility to

minimize any measurement errors. The LAPSS data acquisition system was calibrated with special consideration

for the scale range change accuracy. The lab was painted black with additional black cloth and drapes added to

eliminate any stray light. A rail system was added on the centerline between the two lamps. A metric ruler was

attached to the rail as an indicator of the distance from the lamps to the solar cell test plate. The solar cell test

plate was suspended from the rail so that its center was aligned to the center of the lamps. These changes are

shown in figures 5 and 6.

Test Procedure

The solar cell area was measured and its dimensions were used to calculate the view factor over a range of

distances from the lamp. The lamps were fired 20 times with one minute wait intervals between flashes to warm

up the electronics and lamps.

The solar cells were mounted to the test plate and set at a distance of 635 cm. Distances are estimated to

be within 1 mm. Lamp intensity (pulse network voltage) on the LAPSS was adjusted until the solar cell ISC

matched the 1 sun AM0 measured previously. The monitor cell nominal current was recorded and used for a

Figure 4 Arc Lamp Sequence (2000 frames/sec) frames 57,60, 61, & 62.

1 1 1 2

1 3 1 4

NASA/CP—2005-213431 142
correction to normalize the intensity variation. This test was repeated 5 times with the monitor cell correction less

than 0.25% for each of the IV curves to eliminate subtle variations due to lamp intensity and repeatability.

 After the 1 sun measurement, the test plate was moved to the ~2 sun position and the tests were repeated 5

times as before. No adjustments were made beyond the original settings. The test plate was then moved in to a

5 sun position and the IV curves were measured. Testing was continued at other concentration levels. With this

setup, a maximum ~62 suns is just outside the face of the lamp housing.

Test Results

The ISC of a solar cell is used for setting the light intensity and calibration. The 5 sets of data points at each

Figure 5. LAPSS test set up Figure 6. Test plate rail system

0.0

200.0

400.0

600.0

800.0

1000.0

0 2 4 6 8 10 12 14 16

Vendor 1

Triple Junction Cells (designed for ~ 10 Suns)

Isc (mA)

Concentration Based on View Factor

Linear Behavior

Linear/Sublinear Behavior

(measured Isc)

Figure 7 Vendor 1 Solar Cell Data

NASA/CP—2005-213431 143
concentration level were averaged and then the ISC is plotted as a function of the concentration level. The test

results are shown in figures 7-8. A black line is added to the plot to depict a “linear” behavior of the ISC as

commonly used in concentrator measurements. None of the cells exhibited a true ‘linear“ response.

The cells plots shown depict either slightly superlinear or sublinear behavior with concentration. Other

performance parameters such as Fill Factor, Maximum Power, and Open Circuit Voltage (VOC) all increased as

expected with increases in intensity. Details of the cell performance are not presented here as these test are

primarily for proof-of-concept.

CONCLUSION

 This paper is the second publication of this method. It is intended to provide a simple means of calculating the

view factor for a given source and sample and account for variations in light intensity at a given distance from the

light source(s). The mathematical calculations include the derivation and integration required to solve for the view

factor, its solution, and a way of applying that calculation to test systems. Knowledge of the light source is not

required. Once the view factor calculation is applied, the value was used to determine the intensity at a given

distance from the source. 4 solar cells were measured at a variety of intensities. The results show that this

method is viable for measuring and calibrating concentrator cells. Solar cells have been tested to concentrations

in excess of 500 suns.

These tests indicate that most solar cells do not exhibit a proportional 1:1 short circuit current behavior with

intensity. Solar cells and arrays measured under concentration determined by the ratio of ISC are suspect to the

accuracy of the concentration level. Multi-junction cells with inherent current limiting due to spectral separation

and current mismatch of junctions can exhibit even greater variation in linearity. This test eliminates the errors

0.0

200.0

400.0

600.0

800.0

1000.0

0 2 4 6 8 10 12 14 16

Vendor 2

Triple Junction Solar Cell (~10 Sun Design)

Isc (mA)

Concentration Based on View Factor

Linear Behavior

Linear Behavior

(measured Isc)

Figure 8 Vendor 2 Cell Data

NASA/CP—2005-213431 144
associated with determining the concentration by using mathematical analysis. Errors introduced with this

method include assumptions about the uniformity and geometry of the source, and positioning of the test cells.

A need for this measurement technique has arisen from the fact that Vertical Multi-Junction (VMJ) Cells

exhibit a superlinear behavior, having poor performance at 1 sun [7]. Other cells designed for high intensity also

experience this non-linearity. There are also no known reliable standards for calibration of solar cells at high

intensity, it is hard to assume that the current of a cell remains linear over many orders of magnitude in

concentration especially when series resistance, minority carrier lifetimes, and recombination centers change with

intensity.

Refinement of this method can be implemented for better accuracy at higher concentrations if so desired.

The model of the arc lamp as a rectangular source could be changed to reflect the actual shape of the lamp, all

depending on how far the modeler wants to go. Dimensional accuracy of the test plate could also be improved,

high precision optical rails with fine positioning can help to avoid errors at high-intensity. Positioning errors tend

to increase with higher concentration levels, at 60 suns a 1 cm. difference in position changes the intensity by

nearly two suns.

It is the intention of this paper to provide a demonstration of a method to calibrate concentrator cells. NASA

GRC facility is available for testing of cells/arrays with the limitations of the 50 Volts and 20 Amps.

REFERENCES

[1] D. Scheiman, B. L. Sater, D. Chubb, P. Jenkins, “Measurement and Characterization of Concentrator Solar

Cells”, 3P-C3-76, 3rd World Conference on Photovoltaic Energy Conversion, May 2003.

[2] B. Good and D. Chubb, “Effects of Geometry on the Efficiency of TPV Energy Conversion”,

Thermophotovoltaic Generation of Electricity, Third NREL Conference, 487 (AIP, 1997)

[3] R. Siegel and J. Howell, “Thermal Radiation Heat Transfer”, 2nd Edition, (McGraw-Hill, 1980)

[4] J.R. Ehlert and T.F. Smith, “View Factors for Perpendicular and Parallel Rectangular Plates”, Journal of

Thermophysics, Volume 7, No. 1: Technical Notes, p173-175, March 1992.

[5] R.L. Mueller, “The Large Area Pulsed Solar Simulator (LAPSS), JPL Publication 93-22, Rev A. , July, 1994

[6] S. C. Haydon, “An Introduction to Discharge and Plasma Physics”, 217-226, Department of University

Extension, University of New England, 1964 (Armidale, N.S.W., Australia).

[7] B. Sater and N. Sater, “High Voltage Silicon VMJ Solar Cells for up to 1000 Suns Intensities”, 29th IEEE

Photovoltaics Specialists Conference, 1019,

NASA/CP—2005-213431 145
APPENDIX A : View Factor (Rectangle to Rectangle)

A view factor is defined as the fraction of area in the total field of view (180° hemisphere) of one surface as seen

from another surface, its value is a maximum of 1. The general equation for the view factor between two finite

areas is;

†

F1-2 = 1

A1

cosq1 cosq 2

ps

2

A1

Ú

A2

Ú dA1dA2 1

where F1-2 is the view factor from finite surface A1 to finite surface A2. Angles q1 and q2 are the angles between the

two surfaces from a point dA1 to dA2 and vice versa. Applying the general equation 1 to two surfaces which are

parallel rectangles of different size and not coaxial (see figure 1), the following substitutions can be made from

this graphical layout.

† 

s

2 = z2 + (x -x)2 + (y -h)2 cosq1 = cosq 2 = z

s

2

by substituting the above, equation 1 can be re-written to the following integral based on the four corner

coordinate terms (x, y) for the source rectangle A1, and (h, x) for the rectangle A2. The integral is in the following

form:

† 

G(x, y,h,x) = z2

p

x

Ú

h

Ú

y

Ú xÚ xx

z2 + (x -x)

2 + (y -h)2 [ ]

2 dxdydhdx 3

and solving equation 3 using symbolic mathematics, the final solution becomes a series of summations based on

the corner coordinates (x1,y1; x1,y2; x2,y1; x2,y2) for A1 and (h1, x1; h1, x2; h2, x1; h2, x2) for A2 of the parallel

rectangles. The location of the axis for either coordinate system is random, as long as both rectangles are

parallel to each other and their relative positions do not change. This solution is shown below;

NASA/CP—2005-213431 146
† 

F1-2 = 1

(x2 - x1)(y2 - y1) (-1)

(i+ j +k+l)

G(xi, y j ,hk ,x l)

l=1

2

Â

k=1

2

Â

j=1

2

Â

i=1

2

Â 4

where

G(x,y,h,x) is : 

† 

G = 1

2p

(y j -hk ) (xi -xl)

2 + z2 [ ]

1/ 2

tan-1 y j -hk

(xi -x l)

2 + z2 [ ]

1/ 2

Ï 

Ì Ô 

Ó Ô 

 ̧ 

 ̋ Ô 

 ̨ Ô 

+ (xi -xl) (y j -hk )

2 + z2 [ ]

1/ 2

tan-1 xi -x l

(y j -hk )

2 + z2 [ ]

1/ 2

Ï 

Ì Ô 

Ó Ô 

 ̧ 

 ̋ Ô 

 ̨ Ô 

- z

2

2

ln (xi -x l)

2 + (y j -hk )2 + z2 [ ]

Ï 

Ì 

Ô 

Ô 

Ô 

Ô 

Ô 

Ó 

Ô 

Ô 

Ô 

Ô 

Ô 

 ̧ 

 ̋ 

Ô 

Ô 

Ô 

Ô 

Ô 

 ̨ 

Ô 

Ô 

Ô 

Ô 

Ô 

5

A computer program can be written to calculate the view factor. Please note that double precision numbers

should be used to avoid errors, large numbers are added and subtracted with small differences in the lower digits,

single precision does not have the resolution to retain these small differences.

Example:

Rectangle A1 has corners located at coordinates (-2,1), (-2,10), (4,1), (4,10)

Rectangle A2 has corners located at coordinates (-3,2), (-3,11), (3,2), (3,11)

They are the same 6 x 9 size except offset by 1 unit in both axes.

Calculate the view factor if they are separated by 2 and 10 units, z = 2, 10

Plugging the coordinates into the view factor calculation equations 4 and 5 and performing the summation results

in the view factors: F1-2 (z=2) = .54926 and F1-2 (z=10) = .12407

NASA/CP—2005-213431 147
HIGH-ALTITUDE AIR MASS ZERO CALIBRATION OF SOLAR CELLS1

James R. Woodyard 

Wayne State University, Detroit, Michigan 

David B. Snyder, NASA Glenn Research Center 

Cleveland, Ohio 

ABSTRACT 

Air mass zero calibration of solar cells has been carried out for several years by NASA Glenn Research Center

using a Lear-25 aircraft and Langley plots. The calibration flights are carried out during early fall and late winter 

when the tropopause is at the lowest altitude. Measurements are made starting at about 50,000 feet and

continue down to the tropopause. A joint NASA/Wayne State University program called Suntracker is underway

to explore the use of weather balloon and communication technologies to characterize solar cells at elevations 

up to about 100 kft. The balloon flights are low-cost and can be carried out any time of the year. AM0 solar cell

characterization employing the mountaintop, aircraft and balloon methods are reviewed. Results of cell 

characterization with the Suntracker are reported and compared with the NASA Glenn Research Center aircraft 

method. 

INTRODUCTION

It is important in characterizing solar cells for use in space-power applications that the spectral irradiance of the 

calibration-light source is within a percent of the spectral irradiance of air mass zero conditions (AM0). Spectral 

irradiance differences greater than a few percent can result in calibration errors; the magnitude of the errors 

depends on the structure of the solar cell. In the case of single-junction cells, the current-voltage characteristics 

are not very sensitive to small differences in the spectral irradiances of calibration-light sources because the 

spectral response is not sensitive to spectral irradiance. The current in a single-junction solar cell under AM0 

normal incidence operating at a voltage V is given by 

∫ =

2

1

( ) ( ) ( , )

λ

λ

Io V So λ R λ V dλ (1) 

where So(λ) is the absolute AM0 spectral irradiance of sunlight and R(λ,V ) is the spectral response of the cell 

at wavelength λ and voltage V . In the ideal case, the spectral response is independent of the irradiance of the 

light source. The spectral response depends on the opto-electronic properties of the materials used in the 

fabrication of the cell that include, but are not limited to, the wavelength dependence of the optical absorption

coefficient; optical band gap, material thickness, doping, temperature and quality; and carrier mobility and 

lifetime. λ1 and λ 2 are the lower and upper cut-off wavelength values where the spectral response no longer 

contributes to cell current. 

The spectral irradiance of laboratory-based solar simulators is different than the AM0 spectral irradiance. The 

simulator is set to “AM0” intensity by adjusting the intensity to produce the short-circuit current in a standard cell, 

 1

 This work was supported in part under NASA Grant NAG3-2801. 

NASA/CP—2005-213431 148
i.e., a cell calibrated under AM0 conditions. This approach may be used because the spectral response of 

single-junction solar cells is somewhat insensitive to spectral irradiance. Adjusting the intensity of the simulator 

will compensate for spectral irradiance differences when compared to AM0 over the range of the spectral 

response of the cell. The adjustment produces a spectral irradiance that is larger than AM0 in some regions of 

the spectrum and smaller than AM0 in other regions of the spectrum. Following adjustment of the simulator 

intensity, cells may be characterized under “AM0” conditions. This method may be used as long as two 

conditions are met. First, it is necessary that the simulator is stable, meaning that spectral irradiance remains 

constant during the measurements on the standard cell and the cells to be characterized. Second, the voltage 

dependence of the spectral responses of the standard cell and cells to be characterized must be the same and 

not influenced by differences in the spectral irradiances of the solar simulator and AM0. The method requires 

stable standard cells for each of the types of single-junction cells to be characterized. Laboratory-based 

“AM0”characterization of single-junction solar cells has been carried out for many years with good results using 

this method. 

The evolution of solar-cell technology for space applications has resulted in “state-of-the-art” cells with four and 

five junctions in series. Each junction is designed with a spectral response matched to one region of the 

spectral irradiance of AM0 in order to optimize the efficiency of solar cells. The current in a four-junction solar 

cell operating at a given voltage is given by: 

∫ =

8

1

( ) ( ) ( , )

λ

λ

Io V So λ R λ V dλ and (2) 

 

∫ ∫ ∫ ∫ = = = =

8

7

4 4

6

5

2 2 3 3

2

1

4

3

( ) ( ) 1( , 1) ( ) ( , ) ( ) ( , ) ( ) ( , )

λ

λ

λ

λ

λ

λ

λ

λ

Io V So λ R λ V dλ So λ R λ V dλ So λ R λ V dλ So λ R λ V dλ (3) 

 

where the variables in Equation 2 are the same as in Equation 1 except λ1 and λ8 are the lower and upper cutoff wavelength values, above and below which the spectral response is negligible and no longer contributes to 

cell current. The spectral response in Equation 2 characterizes the overall operation of the four junctions in 

optical absorption and carrier transport. However, the spectral responses and voltages in Equation 3, 

Ri(λ ,Vi) and V respectively, are subscripted to show that they are different for each of the four junctions. The 

voltage across the cell is equal to the sum of the voltages across each of the four junctions, namely, 

. The wavelength ranges on each of the integrals, in the most general case, will overlap since it is 

not possible to fabricate materials with sharp wavelength cut-offs. Equation 3 shows the series nature of the 

current in multi-junction solar cells, namely, the current is the same in each of the junctions. 

i

∑ =

= = 41

i

i V Vi

The sensitivity of a four-junction solar cell to spectral irradiance can be illustrated with an example. Consider a 

cell that has been optimally designed for AM0 is to be characterized with a solar simulator. Assume the solar 

simulator has a spectral irradiance that is less than AM0 in the λ1 and λ 2 wavelength range and the same as 

AM0 in the other three wavelength ranges shown in Equation 3. The lower spectral irradiance will result in less 

current in the junction optimized for the λ1 and λ 2 wavelength range which in turn will limit the current in the 

cell due to the series nature of the four junctions. Equation 3 shows that the current reduction in the four 

junctions must be accomplished through changes in spectral responses of the other three junctions; this is the 

case because the spectral irradiances in the other three wavelength ranges are assumed to be the same as 

AM0. The collective interaction of the four junctions will result in redistribution of the cell voltage across the four 

junctions, which in turn changes the spectral responses of the four junctions and the cell current. 

The role of the interaction of four junctions in the operation of a multi-junction solar cell, as compared to a 

single-junction cell, can be illustrated with an example. Assume the average spectral irradiance and the 

average spectral response are the same in the four wavelength regions in Equation 3. A one percent decrease 

in the spectral irradiance relative to AM0 over the λ1 and λ 2 wavelength range will result in about a one percent 

decrease in the cell short-circuit current. A single-junction junction solar cell that responds in a similar fashion 

NASA/CP—2005-213431 149
1.5

400 500 600 700 800 900

2.0

Q

u

a

ntu

m

Efficie

n

c

y

Wavelength (nm)

lt

)

Triple-Junction QE Dependence on Voltage Bias

Under AM0 Light Bias

Figure 1. Dependence of quantum efficiency of an a-Si:H alloybased triple-junction thin-film solar cell on forward bias [1] 

0.0

0.1

0.2

0.3

0.4

0.5

Vo

0.5

1.0

age (V

over the λ1 to λ8 wavelength range will experience only a 0.25 % decrease in short-circuit current. The reason 

is a one percent decrease in the integrated spectral irradiance over the λ1 and λ 2 wavelength range in the 

multi-junction cell corresponds to a 0.25 % decrease in the integrated spectral irradiance over the λ1 to λ8

wavelength range in the single-junction cell 

 

 

A calibration procedure for multi-junction solar cells that uses a standard cell to set a solar simulator to “AM0” 

intensity may result in data that are not useful in optimizing the design of a test cell for space power generation. 

Assuming the voltage dependence of the spectral responses of each of the junctions in the standard and test 

cells are the same under the simulator “AM0” conditions, the junctions may be operating under conditions that 

are vastly different than AM0 conditions. It is possible that the test cell current-voltage characteristics measured 

under “AM0” conditions may not be useful in optimizing the cell design to improve efficiencies at the one percent 

level. Moreover, the complex nature of the interaction of the junctions does not lend itself to the use of an 

optical technique to compensate for the deficiencies in the “AM0” spectral irradiance. 

The differences in the “AM0” and 

AM0 spectral irradiances are more 

problematic at the maximum power 

point than short-circuit conditions. 

The reason is the electrostatic 

potential barriers in each of the 

junctions are relatively small at the 

maximum power point as compared 

to short-circuit current conditions. 

Redistribution of voltages across the 

junctions can produce relatively large 

changes in the electrostatic potential 

barriers and produce major changes 

in the spectral responses of the 

junctions. Figure 1 shows the effect 

of forward bias on the quantum 

efficiency of a solar cell [1]. The 

solar cell is a triple-junction a-Si:H 

alloy-based thin-film solar cell that 

was illuminated with a solar simulator 

with an AM0 spectral irradiance. The 

spectral irradiance was within one 

percent of AM0 in the wavelength 

range where the spectral response 

contributed to cell current. The figure shows the maximum quantum efficiency is at a wavelength of about 450 

nm, serving as evidence that the top junction in this particular cell was limiting the current of the cell under shortcircuit conditions. The maximum in the quantum efficiency shifted from 450 to 600 nm as the forward bias 

approached the maximum-power point showing that the middle and bottom junctions limited the cell current. 

The quantum efficiency of the cell changed markedly when the spectral irradiance of the simulator was altered 

[1]. A history of particle irradiation can also have a large effect on the dependence of the quantum efficiency of 

multi-junction cells on forward bias thereby further complicating the optimization of design of cells for space 

power generation in radiation environments. 

It is clear that the voltage dependence of the spectral responses of multi-junction solar cells complicates 

optimization of cell design. While there are characterization methods that make it possible to use solar 

simulators in advancing the multi-junction solar cell technology, the series nature of the cells places more 

demands on the need for standard cells characterized under AM0 conditions. AM0 conditions are available only 

in space; near AM0 conditions can be achieved at altitudes in excess of 100,000 ft. The demand for greater 

access to AM0, and the costs associated with AM0 calibration, has generated interest in exploring lost-cost 

methods for AM0 solar cell calibration. The NASA supported Suntracker program is an attempt to meet this

challenge. 

NASA/CP—2005-213431 150
AM0 SOLAR CELL CALIBRATION MEHTODS 

Efforts to develop new methods for AM0 calibration of solar cells should be founded in an awareness of current 

calibration methods, a knowledge of fundamental principles, and possible shortcomings of existing methods. 

Reviewing analyses of data collected by various methods is also an instructive way to gain a better 

understanding of the methods. Mountaintop, aircraft and balloon-based methods for AM0 calibration of solar 

cells are reported in the literature. While there have been a number of satellite-based measurements, no space 

calibration method has emerged that is available to the photovoltaic community for producing solar cell 

standards. A photovoltaic engineering test bed facility for use on the International Space Station has been 

designed but not implemented [2]. This section will review the mountaintop, aircraft and balloon-based methods 

used in AM0 calibration of solar cells. 

Mountaintop method 

Laboratory-based solar simulators have been used since solar cells became attractive for space-power 

applications. However, it was recognized by Zoutendyk that sunlight should be used “ to diminish uncertainty in 

the design of space solar cell power systems” [3]. He was one of the first investigators to attempt to correct for 

the effects of the atmosphere on the spectral irradiance of sunlight. A review of his work with silicon singlejunction solar cells serves as a basis for understanding some of the challenges associated with AM0 calibration 

methods. 

Zoutendyk assumed the spectral irradiance at a given air mass Sm(λ)is given by: 

m Sm So e ( ) ( ) ( ) α λ λ λ − = (4) 

where α(λ) is the monochromatic atmospheric absorption coefficient per unit air mass and is the geometric 

air mass. He defined the geometric air mass as the ratio of the path length of the sunlight through the 

atmosphere at a zenith angle

m

θ to the path length for the sunlight when the sun is overhead and the zenith 

angle is zero. The geometric air mass was taken as the secant of the zenith angle, namely, m = sec(θ ). The 

sea-level irradiance at a given air mass is: 

∫

∞

°

− = λ λ α λ S S e dm m o ( ) ( ) . (5) 

The monochromatic short-circuit current at a given air mass was assumed to be given by: 

m Iscm Isc e ( ) ( ) 0( ) α λ λ λ − = where Isco(λ) = R(λ)So(λ) (6) 

where Isco(λ) is the monochromatic short-circuit current under AM0 conditions. The short-circuit current of a 

single-junction cell over λ1 to λ 2 wavelength range where the spectral response contributes to the current is: 

λ λ λ λ λ

λ

λ

α λ α λ

λ

λ

I R S e d I e dm sc m scm o ∫ ∫ − − = =

2

2

( ) 0 ( )

2

1

( ) ( ) ( ) . (7) 

Equation 7 serves as the basis for the use of Langley plots to characterize solar cells. The exponential term 

may be factored out of the integral if α(λ) is assumed to be constant over the λ1and λ2 wavelength range. 

The short-circuit current for a given air mass can then be written as: 

NASA/CP—2005-213431 151
sco m sco m Iscm e I d e I α

λ

λ

α λ λ − − = = ∫

2

1

( ) (8) 

where Isco is the AM0 cell short-circuit current. Taking the Log10 of both sides of Equation 8 gives: 

Log Iscm m Log Isco 10 4343 10 = −0. α + (9) 

which is the theoretical equation used to determine the short-circuit current of solar cells under AM0 conditions. 

The logarithm of the short-circuit current is plotted on the ordinate of a semi-log graph and the air mass on the 

abscissa. The graph is referred to as a Langley plot. The data are fitted to a straight line using a least-squared 

method and the line extrapolated to m . The intercept of the straight line with the ordinate is taken as the 

short-circuit current under AM0 conditions. The slope of the graph is

= 0

− 0.4343α and may be used to determine 

the atmospheric optical absorption coefficient. It is important to emphasize the constancy of the atmospheric 

optical absorption coefficient and the use of the “air mass” concept implies the following: 

1. The optical absorption coefficient must be constant with respect to wavelength over the range of 

wavelengths where the solar cell spectral response contributes to cell current. If it is not constant, 

using Equation 9 to analyze data will produce errors in the extrapolated AM0 short-circuit current. 

2. The concentration of optically absorbing atomic and molecular species in the atmosphere and their 

altitude dependence must not change for the duration of the short-circuit current as a function of air 

mass measurements. If the concentrations are changing during the measurements as a result of 

weather fronts, turbulence in the atmospheric, solar heating of the atmosphere etc., Equation 9 may not 

be linear and linear extrapolation of the short-circuit current to zero air mass may be in error. 

3. The optical absorption coefficient must not be large enough to totally absorb the AM0 spectral 

irradiance at any air mass over the range of wavelengths where the solar cell spectral response 

contributes to cell current. If there are regions of the spectral irradiance where the sunlight is totally 

absorbed as it travels through the air mass, then the use of the Langley method to determine the AM0 

short circuit will produce erroneous results. 

4. Only normally incident sunlight must contribute to the short-circuit current. Scattered sunlight, referred 

to as “sky radiation” by Zoutendyk, must not contribute to the short-circuit current. Additionally, the 

presence of reflected light, or light produced by any other mechanisms, may introduce errors in the 

determination of the AM0 short-circuit current. 

Zoutendyk set up a tracking system with silicon single-junction solar cells at an elevation of 7.4 kft on a 

mountaintop and carried out diurnal measurements of cell short-circuit current and temperature as a function of 

the zenith angle as the sun moved across the sky. The cell short-circuit current was defined at the current

through a 1.000 precision resistor in series with the cell. The sea level geometric air mass was calculated 

using 

Ω

m = sec(θ ). The data were analyzed using a Langley plot to arrive at cell AM0 short-circuit currents. 

The short-circuit current was corrected for cell temperature, precision resistor temperature and the earth-sun

distance. The cells were then flown on the Ranger III spacecraft and cell data downlinked. The agreement 

between the AM0 short-circuit current measurements on the mountaintop and space was reported to be 

± 2% [3]. 

It is noteworthy to evaluate the constancy of the atmospheric optical absorption coefficient in mountaintop work 

to understand the utility of Langley plots. Equations 8 and 9 show that the exponential term is assumed to be 

constant over the λ1 and λ 2 wavelength range in order to permit factoring it out of the integral. The cut off 

wavelength of Zoutendyk’s solar cell at low wavelengths was λ1 = 450 nm because of the optical properties of 

the cover glass on the cells. The high wavelength cut off was about λ 2 = 1200 nm due to the band gap of the 

silicon material used in the solar cells. Analyses were carried out using values of the atmospheric optical 

NASA/CP—2005-213431 152
absorption coefficients in the wavelength range that were reported by Moon [4]. The coefficients 

ranged between 0.05 and 0.96 per air mass. The spectral response of the solar cells peaked at about 850

where the atmospheric optical absorption was about 0.1 per air mass. The geometrical air masses used by 

Zoutendyk must be multiplied by 0.7 to correct for the 7.4 kft altitude [5]. For 

450 − 1100 nm

nm

α = 0.1 per air mass and m = 0.7 , 

the exponential term in Equation 8, has a value of about 0.93. At the largest and smallest values of the optical 

absorption coefficient, 0.96 and 0.05 per air mass, the values of the exponential term will be 0.51 and 0.96, 

respectively. Clearly the exponential term varies with wavelength when Moon’s atmospheric optical absorption 

coefficients are used in Equation 8. However, as shown in Equation 7, the exponential term is convoluted with 

the cell spectral response. The spectral response is always less than one; it decreases from a maximum value 

at 850 to approximately zero at the cut off wavelengths. The effect of convolution of the spectral response 

with the exponential term in Equation 7 is to decrease the weighting of the exponential term in the integral. A 

non-constant exponential term in Equation 7 will produce a concave up feature in Langley plots [4]. There is no 

evidence of a concave up feature in the Langley plots in Zoutendyk’s work. This suggests variations in the 

atmospheric optical absorption coefficients were small enough so as to not invalidate the use of Langley plots to 

determine solar cell AM0 short-circuit currents. 

nm

It is surprising the extrapolated AM0 short-circuit currents agree with the space measurements to within 2 %. It 

may be the case that the optical absorption coefficients used by Zoutendyk are not appropriate for the 

conditions under which the mountaintop measurements were carried out. There are three reasons for this

conjecture. 

1. The ratios of Zoutendyk’s measured and calculated short-circuit currents as a function of air mass differ 

considerably. He used Equation 7 to calculate short-circuit currents along with a standard AM0 spectral 

irradiance [5], the spectral response of the cells and atmospheric optical absorption coefficients [6]. In 

every case, the calculated short-circuit currents are smaller than the ones measured, suggesting the 

atmospheric optical absorption coefficients used are larger than the effective optical absorption 

coefficients at 7,400 ft. 

2. The irradiances measured as a function of air mass are also considerably larger than the irradiances 

calculated using Equation 5. Zoutendyk plotted measured irradiance as a function of air mass on 

semi-log plots. The curves are clearly concave up providing convincing evidence of the effect of nonconstant atmospheric optical absorption coefficients. In the case of the irradiance curves, Equation 5 

shows the integral extends over a larger wavelength range and is not convoluted with the cell spectral 

response. The larger wavelength range and absence of the convolution both lead to the full effect of 

the atmospheric optical absorption coefficients on the transmitted sunlight and a concave up feature in 

irradiance plots. 

3. An analysis of Zoutendyk’s data in six Langley plots yields atmospheric optical absorption coefficients 

ranging between 0.079 and 0.101 per unit air mass; the average is 0.087 per unit air mass. The 

average value of Moon’s optical absorption coefficients is about 0.15 per unit air mass in 

the 700 to range where the solar cell spectral response is the largest. The fact that the average 

slope is about 60 % of Moon’s optical absorption coefficients suggests either Moon’s coefficients are 

too large to be used in predicting AM0 short- circuit currents or the atmospheric conditions that

prevailed during Zoutendyk’s measurements are different than the conditions under which Moon’s 

coefficients were determined. Additionally, the variation in the slopes of the Langley plots measured

from day-to-day suggests changing atmospheric conditions may have played a role in the mountaintop 

measurements. 

900 nm

Ritchie recognized the problems associated with using Moon’s atmospheric optical absorption coefficients to 

correct solar cell short-circuit currents. He employed measurements on a mountaintop to produce secondary 

standards [7] that did not employ Langley plots. The secondary standards were based on the use of primary 

standards calibrated with the balloon method and the following equation: 

NASA/CP—2005-213431 153
s

p scm

scm

p

s sco

sco I

I

I I = × (10)

where and are the calculated secondary and measured primary standard AM0 short-circuit currents, 

respectively; and are the secondary and primary standard short-circuit currents, respectively,

measured at the same time on a mountaintop. The balloon method was used to measure . Following the 

mountaintop measurements, the secondary standards were flown on a balloon flight and the AM0 short-circuit 

currents measured; the currents agreed to within 0.5% with the currents predicted using the mountaintop 

measurements based on Equation 10. It is important to note that the spectral responses of the primary and 

secondary standards must be the same when using a primary balloon standard, Equation 10 and mountaintop 

measurements to produce secondary standards. 

s

sco I psco I

scm

s I p

scm I

p

sco I

Aircraft method

The use of an aircraft to carry out high-altitude solar cell measurements at altitudes between 47 kft and 6 kft and 

air masses in the 0.180 – 0.862 range was first reported by Brandhorst [8]. It was suggested that the aircraft

method is attractive when compared to the mountaintop method for three reasons. First, measurements are 

made at lower values of air mass than the mountaintop method resulting in shorter extrapolations of the shortcircuit current on Langley plots. It is expected that the more accurate values of the AM0 short-circuit currents 

will be obtained if the extrapolation is over a smaller range of air masses. Second, the atmosphere should be 

less prone to compositional changes during the relatively short time of the aircraft measurements as compared 

to diurnal mountaintop measurements, i.e., minutes versus hours. Third, the measurements are made at 

altitudes that are above ground haze and low-altitude atmospheric disturbances. 

The aircraft method employed a 4.5” diameter windowless collimator with a collimation ratio of 4:1 that was 

mounted inside the aircraft and extended through a hole in the side of the tail section [9]. The collimator was 

designed to over-fill the cell holder so that the cells were uniformly illuminated even when the orientation of the 

aircraft resulted in a 2-degree error in the pointing of the collimator. The collimator angle was set before each 

flight to the zenith angle of the sun during the measurements. The tail section was not pressurized and the cells 

were exposed to the low pressure and temperature environment that is characteristic of the altitudes at which 

the measurements were carried out. Single-junction silicon solar cells were mounted on a heated stage and the 

cell temperature maintained between 15 and 30 oC with a variation of less than 4 oC. The cell short-circuit

current was taken as the current through a 1.000 

±

Ω precision resistor that was placed in series with the cell, as 

was done by Zoutendyk. The aircraft altimeter was used to measure pressure to an accuracy of 75 ft. The pilot 

used a sight tube mounted next to the controls in the cockpit to orient the aircraft and control the pitch, roll and 

yaw so as to point the collimator at the sun with a pointing accuracy of better than ± 2 degrees. Altitude, cell 

short-circuit current and cell holder temperature were measured at altitude intervals of 5 kft during descent from 

47 to 6 kft. 

A standard atmospheric model was used to convert the altitude measurements to pressure [5]. The air mass 

was calculated using: 

m = (P / Po)secθ (11)

where P is the pressure at which the cell short-circuit current was measured and is the sea-level pressure. 

Langley plots were produced and extrapolations carried out to determine the AM0 short-circuit current of the 

single-junction silicon solar cells. The AM0 short-circuit currents were corrected for cell temperature, precision 

resistor temperature, ozone absorption and the earth-sun distance. The extrapolated AM0 short-circuit current 

was corrected for ozone absorption using the cell spectral response; ozone absorption coefficients in 

the wavelength range [10]; ozone altitude profile [11]; and the percent of the total column ozone 

above the aircraft during measurements. The effect of ozone absorption on the short-circuit current of singlejunction Si and GaAr solar cells was estimated to be 1.04 and 1.23 %, respectively. Brandhorst reported that 

Po

400 − 700 nm

NASA/CP—2005-213431 154
all the Langley plots were straight lines [8,9]. However, there were differences in the slopes of the Langley plots 

from flight-to-flight suggesting atmospheric conditions, while perhaps constant during a flight, changed from 

flight-to-flight. The atmospheric optical absorption coefficients, as determined from the slopes of the Langley 

plots in the publications, ranged between 0.09 and 0.30 per air mass. The change in the slopes suggests there 

were variations in the concentration of optically absorbing atomic and molecular species in the atmosphere from 

flight-to-flight. The agreement in the AM0 short-circuit currents, measured by the aircraft method and the 

mountaintop method that used Equation 10, was ± 0.9 %. The AM0 short-circuit currents measured during three 

separate flights were reproducible within 1 % even though the slopes of the Langley plots, and therefore the 

atmospheric conditions, were different. 

±

Hadley analyzed data from three single-junction silicon cells. Two were balloon calibrated primary standards 

and one was a secondary standard calibrated using mountaintop measurements with Equation 10. All three of 

the cells were characterized with the aircraft method. He found that the cell AM0 short-circuit currents 

measured by the aircraft method were consistently about 1.6 % lower [12]. Hadley pointed out that the Langley 

plots for different days had different slopes and some of the plots appeared to be concave up . While there was 

a need to make spectral corrections because the spectral responses of the cells were slightly different, it was 

not possible since the different slopes for different days “indicate that there cannot be a unique value of the 

spectral correction factor.” The effect of Hadley’s work was to suggest that while it was agreed that the use of 

the Langley plots in the mountaintop method was faulted, there was evidence that there were also problems in 

using Langley plots to analyze data from the aircraft method. 

Subsequently, the aircraft method was used to measure silicon single-junction solar cell short-circuit currents 

versus altitude. The measurements resulted in Langley plots that exhibited an anomalous behavior. [13]. The 

plots had a curve with two linear regions, each of which had a different slope. There was a “break” in the curve 

where the linear segments met. It was determined that the “break” occurred at an air mass that corresponded 

to the altitude of the tropopause. The short-circuit currents measured at altitudes above the tropopause 

produced a linear plot with a larger slope than the slope of the currents measured at altitudes below the 

tropopause. Extrapolation of each of the linear segments produced different AM0 short-circuit currents. The 

Langley plots for data collected at altitudes above the tropopause extrapolated to larger AM0 short-circuit 

currents than the plots for data collected at altitudes below the tropopause. The slopes of the Langley plots for 

data collected above the tropopause were the same on a month-to-month basis while the slopes for data 

collected below the tropopause were different on a month-to-month basis. The observation that the slope above 

the tropopause was constant on a month-to-month basis suggests the concentrations of the optically absorbing 

atomic and molecular species are the same on month-to-month basis for species that absorb in the region of the 

solar spectrum that contributes to cell current [13]. The atmospheric optical absorption coefficient determined 

from one of the Langley plots for the data collected above the tropopause was about 0.20 per air mass. On the 

other hand, the fact that the slopes below the tropopause changed on a month-to-month basis suggests 

concentrations of the absorbing atomic and molecular species are different below the tropopause on a month-tomonth basis. The atmospheric optical absorption coefficient determined from one of the Langley plots for the 

data collected below the tropopause was about 0.09 per air mass. 

A comparison was made of AM0 short-circuit currents of five silicon single-junction solar cells measured with the 

balloon and aircraft methods [13]. The accuracy of the balloon measurements was estimated at 0.9 %. Only 

aircraft data collected above the tropopause were used. Differences in the AM0 short-circuit currents measured

by the two methods ranged from 0.3 to 1.2 % with an average of 0.7 %. The AM0 short-circuit currents 

measured with the aircraft method were always larger than those measured with the balloon method. It was 

concluded that the measurements “show excellent agreement”, and the aircraft method must be used above the 

tropopause in order to produce good results. 

±

Brandhorst carried out a series of experiments with the aircraft method to determine the wavelength region of 

the anomalous effect [14]. Optical filters were placed over solar cells and data collected both above and below 

the tropopause. The Langley plots had straight lines for data collected with filters that transmitted red and green 

light; an anomalous plot with a “break” was produced with data collected with a filter that transmitted blue light. 

The measurements suggested the anomalous effect was due to an atomic or molecular species that absorbed 

in the blue wavelength region of the solar spectrum, and that the species existed primarily above the 

tropopause. 

NASA/CP—2005-213431 155
The anomalous Langley plot can be understood by modifying Equations 7, 8 and 9 to include the effect of two 

atmospheric optical absorption coefficients. The atmospheric optical absorption coefficient for an optically 

absorbing species that exists primarily above the tropopause is taken as α1 for a range of wavelengths from 

λ1 to λ2 , the wavelength range where the spectral response contributes to the cell current. The optical 

absorption coefficient in the same wavelength range is α2 for a different species that exists primarily below the 

tropopause in the same range of wavelengths. Equations 7, 8 and 9 above the tropopause become: 

 and sco

m

o

m

scm I e R S d e I 1

2

1

1 ( ) ( ) α

λ

λ

α λ λ λ − − = = ∫

(12)

scm sco Log I m Log I 10 4343 1 10 = −0. α + . 

The slope of the Langley plot above the tropopause is − 0.4343α1m and the extrapolated AM0 short-circuit 

current is . Below the tropopause: sco I

 and cot

2

2

1

2 ( ) ( ) s

m

ot

m

scm I e R S d e I α

λ

λ

α λ λ λ − − = = ∫

(13)

 10 4343 2 10 cot 0. scm s Log I = − α m + Log I

where (λ) Sot

)

is the solar spectral irradiance just below the tropopause; it is corrected for absorption due to the 

optically absorbing species above the tropopause. is the short-circuit current due to the convolution of s cot I

Sot(λ with R(λ). The slope of the Langley plot below the tropopause is − 0.4343α2m and the extrapolated 

AM0 short-circuit current is . The “break” in the straight line segments on the Langley plot defines two 

regions of air mass where the data have linear characteristics, namely, a straight line segment with a larger

slope at lower air masses and a straight line segment with a smaller slope at larger air masses. It follows that 

s cot I

α1 in the region of lower air is greater thanα2 in the region of larger air masses. Hence, the anomalous plots 

may be explained by assuming the atmospheric optical absorption coefficients are different above and below the 

tropopause. 

The aircraft method has been developed over the years by investigators at NASA Glenn Research Center

[15,16]. A large body of calibration data has been collected and AM0 standards provided to the PV community. 

The aircraft has been replaced twice and the method improved. The current aircraft is a Lear 25 that houses the 

instrumentation and collimator in a pressurized and temperature controlled compartment. Photographs of the 

Lear 25 aircraft, collimation tube and test cell may be viewed on the NASA Glenn Research Center Web site 

[17]. The ratio of the collimation has been increased to 1:4.5. The method was upgraded to carry out 

measurements every nine seconds during a 6E-4 air mass per second rate of descent from 50 kft down to the 

tropopause. Sources of random error were estimated to be about 0.04 % and agrees with measurements. The 

difference in the average of measurements on a single-junction cell carried out over a twenty-year period and a 

recent measurement was at the 0.05 % level. Systematic errors were estimated to be at the one percent level. 

Space shuttle AM0 short-circuit current measurements on two cells were compared with the aircraft method. 

The aircraft measurements were less than the shuttle measurements by 1.0 and 0.8 % for the two cells; the 

errors were consistent with the estimated systematic errors. 

The role of ozone on the solar cell AM0 short-circuit current measured with the aircraft method has been 

investigated by Snyder and collaborators [18]. The extrapolated AM0 short-circuit current was corrected using 

an ozone correction factor. The correction factor was determined using a calculated AM0 short-circuit current 

NASA/CP—2005-213431 156
and a calculated short-circuit current that included the effects of ozone absorption. The WMO solar spectral 

irradiance [19] and solar cell spectral response were convoluted using Equation 1 to calculate the AMO shortcircuit current where x corresponds to Si, GaAs and InGaP single-junction solar cells. The short-circuit 

current was calculated using Equation 7 and the WMO solar spectral irradiance, solar cell spectral 

response, ozone absorption coefficients [20] and total column ozone valueT . The ozone correction factor 

 for each of the types of solar cells was found using: 

x

scoc I

x

scdu I

co

x Fo















 − =

oc

x

scdu

x

scdu

x

x scoc

o I T

I I F (14) 

where the units ofT are Dobson units, d.u. The correction factors were found to be insensitive, at the 5 % 

level, to the value of the total column ozone, different sets of spectral response data, and the resolution of the 

spectral response data used in the calculations. The ozone correction factor was reported to increase with an 

increase in the optical band gap of the materials used in the solar cells. 

co

The ozone correction method consisted of employing a Langley plot to first determine the extrapolated AM0 

short-circuit current . Then the current was corrected for optical absorption in the ozone column on the day 

of the flight. The sun zenith angle, total column ozone on the day of the flight and the ozone correction factor 

were used in the following equation to determine the AM0 short-circuit current for the various types of cells: 

x

scol I

0.83 secθ . (15) oc

x

o

x

scol

x Isco = I + F T

The 0.83 factor results from assuming that 83 % of the total column ozone was above the aircraft during the 

measurements. The correction method was used to analyze Si solar cell data collected during 20 flights. The 

correction resulted in an increase of 0.52 % in the Si AM0 short-circuit current as compared to the earlier one 

percent correction method introduced by Brandhorst [9]. Applying Equation 15 to the earlier aircraft 

measurements also reduced the differences noted by Hadley in currents determined with the aircraft and 

balloon methods [12]. The percentage standard deviation of the ozone corrected currents decreased from 0.49 

% with the earlier correction method to 0.28 %, thereby suggesting the importance of using Equation 15 to 

correct for the effect of ozone. 

Applying the correction method to a GaAs solar cell, which has a larger band gap than Si, resulted in AM0 shortcircuits that were different from flight-to-flight. The corrected currents increased with increases in total column 

ozone suggesting that the method was sensitive to the band gap of the cell. The correction method was revised 

in an effort to eliminate the dependence of 

currents on total column ozone [21]. The revised 

method called for first correcting the short-circuit 

current measured at each altitude for ozone 

absorption, then plotting the data and 

extrapolating to air mass zero. The approach 

requires the column ozone at each of the

altitudes at which the cell current is measured. 

The TOMS standard ozone profiles were used to 

calculate column ozone as function of altitude 

[22]. Figure 2 shows the TOM ozone profiles 

used in calculating the column ozone. The 

average ozone density in the figure is plotted at 

the midpoint of each Umkehr layer [23], starting 

at layer zero corresponding to an altitude of 9 kft, 

through layer 9 at 143 kft. The density shown in 

0

20

40

60

80

100

120

0 20 40 60 80 100 120 140 160

Altitude (kft)

Ozone Density (matm-cm/layer)

175M

225M

275M

325M

375M

425M

475M

Figure 2. TOMS ozone profiles as function of altitude and 

total column ozone at mid latitudes [22] . 

NASA/CP—2005-213431 157
the figure at 156 kft includes the ozone above 156 kft. The profiles in the figure are for mid latitudes and total 

column ozone values ranging from 175 to 475 d.u. Mid latitudes are around 45 degrees north, the latitude of the 

NASA Glenn Research Center flights. The maximum altitude for the flights is about 50 kft. Figure 2 shows the 

fraction of the total column ozone above 55 kft decreases with increasing total column ozone. The fraction at 55 

kft is about 0.87 of the total column ozone when the total column ozone is 175 d.u.; the fraction decreases to

about 0.77 at 475 d.u. The fraction is about 0.92 and fairly independent of total column ozone at 40 kft. 

The revised AM0 short-circuit current correction method included converting the aircraft altitudes to atmospheric 

pressure, p . The total column ozone on the day of the flight was obtained and used to select the appropriate

TOMS ozone profile in Figure 2. The fraction of the total ozone column above the aircraft during each of 

the measurements was calculated. The ozone corrected short-circuit current was calculated for each 

measurement using: 

f ( p)

I ( p) scm

I ( p) sc

( ) ( ) ( ) secθ (16) oc

x Iscm p = Isc p + f p Fo T

The ozone corrected short-circuit currents were plotted as a function of pressure on a Langley plot, instead of a 

function of air mass. The linear plot was extrapolated to zero pressure to determine the AM0 short-circuit 

current. Results with the revised method were compared to the one percent method using data for a Si singlejunction solar cell. Data collected over two time intervals, namely a short term and long term, were compared. 

The short-term data were collected on one cell during twenty flights in one year. The long-term data were 

collected on the same cell during thirteen flights over a period of eight years. Table 1 shows differences in the 

average AM0 short-circuit currents using the one percent, Av. Isco 1%, and revised ozone correction methods, 

Av. Isco Revised. The percentage difference in the averages of the currents are 0.52 and 0.54 % for the short 

and long terms, respectively, showing the percentage differences are 

essentially the same for the two periods. The percent standard 

deviation decreases from 0.49 to 0.26 % and 0.72 and 0.48 % for the 

short and long terms, respectively, showing the importance of the

revised ozone correction method in reducing the systematic error in

the aircraft method. The percentage differences between the highest 

and lowest AM0 short-circuit currents decreases from 3.2 to 1.2 % and 

2.5 to 1.0 % for the short and long terms, respectively, again showing 

the improvement of the results with the revised ozone correction 

method. The relatively larger percentage differences in the highest 

and lowest AM0 short-circuit currents is due to the fact that the zenith 

angle ranged between about 48 and 68 degrees during the various 

flights. Both methods show a decrease in the ozone corrected AM0 

short-circuit currents as the zenith angle increases. However, the 

revised ozone correction method is considerably more effective in 

correcting for the effect of the zenith angle. Not shown in Table 1 is

the percentage difference in the average AM0 short-circuit currents

using the revised method for the short and long terms; the difference is 

0.26 % which is within the standard deviation for the measurements. 

Table 1. Comparison of ozone 

correction methods for Si, GaAs 

and InGaP single-junction solar 

cells [21].

Si Short Term 

Av. Isco Revised - Av. Isco 1%=0.52%

1% Revised

(%) (%)

Std. Dev. 0.49 0.26

High-Low 3.2 1.2

Si Long Term 

Av. Isco Revised - Av. Isco 1%=0.54%

1% Revised

(%) (%)

Std. Dev. 0.72 0.48

High-Low 2.5 1.0

GaAs Long Term 

Av. Isco Revised - Av. Isco 1%=1.0%

1% Revised

(%) (%)

Std. Dev. 0.75 0.47

High-Low 2.6 2

InGaP 

Av. Isco Revised - Av. Isco 1%=2.3%

1% Revised

(%) (%)

Std. Dev. 0.45 0.14

Table 1 also shows the differences in the average AM0 short-circuit 

currents for GaAs and InGaP single-junction solar cells using the 1%, 

and revised ozone correction methods. The GaAs data were 

measured in 31 flights over a nine-year period. The InGaP data were 

collected during seven flights over a few years. The band gap of GaAs 

is 1.43 eV. The band gap of the InGaP cell is not known because the 

alloy concentrations were not reported for this particular cell; however, 

it is expected that it is greater than the band gap of GaAs. The 

percentage difference in the averages of the currents corrected by the 

1% and revised methods are 1.0 and 2.3 % for the GaAs and InGaP 

NASA/CP—2005-213431 158
cells, respectively, showing the revised method gives larger currents as the band gap increases. The percent 

standard deviations are reduced for both the cells as was observed for the Si cell, again confirming the 

importance of the ozone correction. The differences in the high and low current values are reduced for both 

cells when the revised method is used. A larger zenith angle results in a larger correction in the AM0 shortcircuit currents for both cells. The revised method is more effective in correcting for larger zenith angles as is 

illustrated by the smaller high-low percentage differences in the AM0 short-circuit currents. However, there is a 

trend for corrected currents to exhibit a decreasing trend with increasing zenith angle. 

Balloon Method 

The first balloon method measurements to determine the AM0 short-circuit current of single-junction silicon solar 

cells was reported by Zoutendyk [24]. A sun tracker weighing about 70 lbs was flown mounted on the top of a 

balloon, as opposed to in a gondola suspended below the balloon. Mounting the sun tracker on top of the 

balloon also reduced the effect of sunlight reflections from the balloon on the cell measurement. The top 

mounting also eliminated anticipated complicating effects of pendulum and torsional motion of the gondola on 

tracking of the sun. The downlinked data included three solar cell temperatures, fourteen cell short-circuit 

currents and seven precision standard voltages. The cell short-circuit current was defined at the current through 

a 1.000 precision resistor that was placed in series with the cell. The weight of the balloon, parachute, sun 

tracker, batteries, transmitter, data acquisition system and other incidental items was 425 lb; the batteries and 

transmitter were mounted in a gondola attached to the parachute which is turned was attached to the bottom of 

the balloon. The maximum volume of the balloon was 175,000 cubic feet and required 7,150 cubic feet of

helium to provide an ascent rate of about 1000 feet per minute. It took about two hours for the balloon to reach 

the 80,000 ft float altitude. The float time for flights was about four hours with two hours of floating time before 

solar noon and two hours after solar noon. Two vehicles were used in support of flights. One was a bus with

communications equipment that maintained a distance of less than 200 miles from the balloon and received the 

downlinked data. The other vehicle was a chase truck that retrieved the equipment and solar cells. An aircraft 

was also available to assist if necessary with tracking the balloon. 

Ω

The solar irradiance at a zero zenith angle and the 80 kft float altitude was estimated to be 95 % of AM0, 

corresponding to an air mass of 0.05. A reduction of 0.3 % was estimated in the air mass 0.05 cell short-circuit 

current, as compared to AM0, considering the spectral response of the solar cells. The sun tracker was able to 

track the sun to within 4 degrees. An error analysis of the measurement system showed a maximum error of 

about 0.6 %. Three flights were carried out over a period of three months and the results of the calibration of 

two silicon solar cells were reported. The cell temperatures were 19

±

±

± 1 oC during the first flight and 32 ± 1 oC 

during the last two flights. The data were corrected for temperature. AM0 short-circuit currents were reported 

for the mean earth-sun distance and 28 o

C cell temperature. No corrections were made for ozone absorption. 

The repeatability in AM0 short-circuit currents for three flights was better than ± 0.5 % and the accuracy 

estimated at ± 0.6 %. 

The balloon method has been further developed over the years by Anspaugh and collaborators [25]. It has 

been used by the NASA Jet Propulsion Laboratory to provide primary standards to the space photovoltaic 

community. The standards are calibrated at about 120,000 ft, the float altitude of the balloon. Balloon flights 

are carried out at least once a year during the summer months. The combined weight of the balloon and top 

and bottom payloads is about 1,400 lb. The main balloon has a volume of about 3.6 E6 cubic feet and requires 

about 24,000 cubic feet of helium. Launching requires the use of a tow balloon, spool vehicle, launch vehicle 

and helium tanker. The top payload weighs about 140 lb and consists of the sun tracker, solar panel, 

transmitters, command receivers, data acquisition electronics, video camera, batteries and miscellaneous items. 

The bottom payload is tethered in a gondola and includes of batteries, ballast module, flight terminate 

equipment, pressure transducer, transmitters, GPS receivers, and a variety of supporting electronic systems. 

The top and bottom payloads have separate parachute and release mechanisms. In excess of 75 cells and 

modules can be calibrated during a flight. An on-board microprocessor-based system is used for data 

acquisition and storage during flights. Current-voltage measurements are made over a voltage range from 

about 100 mV to the cell open-circuit voltage. Cell temperatures are measured both during ascent and at float 

altitude. Cell temperatures are between 30 and 40 oC during ascent and about 75 oC at float altitude. The 

balloon ascent rate was about 900 ft/min and the time from launch to a 120 kft float altitude was about two 

NASA/CP—2005-213431 159
hours; float times ranged between 1.5 to 3.0 hours. Data are downlinked to a base station at the launch site. 

An aircraft with a two-person crew direct both the termination of the flight and recovery activities. The aircraft 

crew in retrieving the two payloads directs a chase truck equipped with communications equipment. 

Laboratory characterization of solar cells and modules is carried out both before and after balloon flights to 

insure the cells are not damaged. Cells with a variety of structures have been flown including single and multijunction cells fabricated from crystalline and amorphous materials. Both thick crystalline and thin-film cells have 

been calibrated. Only data collected with the tracker pointed to within ± 2 % of the sun are used in 

characterizing cells. “Wild” data points have been observed in recent years with modifications in the equipment, 

e.g., adding a video camera and placing some of the transmitters in the top package [26]. The data acquisition 

system has been programmed to exclude “wild” data points from the data files that were used in determining the 

cell AM0 characteristics. The cause of the “wild” points was not know but it was suggested that they may be 

due to water condensation or radio interference. 

A module was flown for 41 flights to evaluate both measurement repeatability and the role of position on the

solar panel. It was felt that changing the position of the module on solar panel could be used to evaluate the 

geometric quality of the solar irradiance “with regard to uniformity, shadowing, or reflections” [27]. The standard 

deviation for 41 flights was 0.46 % and it was concluded that there are no geometric problems. The cell used in 

the 41 flights was damaged and replaced. It was replaced with a set of nine standard cells and modules to 

continue checking the repeatability of the balloon method. The nine standard cells and modules were not flown 

on every flight. Cells and modules were flown between 6 and 26 times. The standard deviations of the AM0 

short-circuit currents for the flights range between 0.23 and 1.0 %. The percent differences between the 

maximum and minimum AM0 short-circuit currents range between 1.3 and 4.4 % for all the standard cells [28]. 

Two of the silicon single-junction standard cells were flown on the Discovery Shuttle and the AM0 short-circuit 

current measured in a manner similar to the balloon method. The differences between one measurement on 

the shuttle flight for each of the two cells, and the average of the AM0 short-circuit currents measured on 

multiple balloon flights, were 0.21 and 0.11 %, respectively. The agreement is seen as “verifying the accuracy 

of the calibration procedures used on the balloon flights” [29]. 

The agreement between shuttle and balloon calibrations for two Si solar cells was 0.21 and 0.11 %, yet the set 

of standard cells have standard deviations for multiple balloon flights that range between 0.23 and 1.0 %. 

Moreover, the percentage differences between the maximum and minimum AM0 short-circuit currents range 

between 1.3 and 4.4 % for the set of nine standard cells [28]. If the agreement between the shuttle and balloon 

methods is taken as representative of the accuracy of the balloon method for silicon single-junction cells, then 

the differences in the balloon measurements may be due to either the spectral responses or instabilities in the 

standard cells. The cell structure and materials are not given for the nine standard cells [28], and for this 

reason, it is not possible to correlate the differences with cell spectral response. The standard cell with a 1.0 % 

standard deviation and a 4.4 % difference between the maximum and minimum AM0 short-circuit current was 

flown on 13 balloon flights. The data may be used to evaluate the role of cell instability on measurements. The 

first current measurement was 166.83 mA; the current measured on the thirteenth flight was 611.11 mA. 

Inspection of the measurements for the 13 balloon flights shows the trend is for the cell current to decrease with 

each successive flight. The data show that the reason for the relatively large cell statistics is a gradual decrease 

in the cell’s current. The behavior suggests the existence of a cell degradation mechanism. It in noteworthy 

that the cell operates at 75 o

C at the float altitude. The cell will operated at 75 oC in excess of twenty hours 

during the 13 flights. It is probable that thermal degradation resulted in the relatively large change in cell 

current. However, opto-electronic degradation and other sources of instability should also be considered. 

SUNTRACKER PROJECT 

A joint NASA Glenn Research Center/Wayne State University program called Suntracker is underway to explore 

the use of weather-balloon and communication technologies to characterize solar cells at elevations up to 

120,000 feet [30-32]. The balloon flights are low-cost and can be carried out any time of the year. 

The Suntracker scientific package includes a collimator with a cell to be calibrated, two GPS receivers, two 

transmitters, two separate battery packs, control electronics and a video camera. Figure 3 shows the scientific 

package. The main body of the package is cylindrical and fabricated from urethane foam; it measures 11” in 

NASA/CP—2005-213431 160
diameter and 11” high. The 

package is about one inch thick 

and provides thermal insulation for 

the electronics. Two battery 

packs are supported 22” from the 

main body by light carbon 

composite tubes that are arranged 

in a “z” configuration. The battery 

packs are mounted in this fashion 

to increase the rotational inertia of 

the package; the 22” length is 

limited by the interior width of the 

mobile unit. The two battery 

packs are enclosed in separate 

carbon tubes and wrapped in 

foam insulation to provide thermal 

insulation. The dimensions of the 

collimator are 1.25"x1.25"x4.00"; 

the front aperture is 1.00"X1.00" 

and the cell area 0.79"x0.79" [x]. 

The dimensions of the collimator 

were selected to provide a 1:4 

collimation ratio and a one degree 

pointing accuracy. The collimator 

prevents light scattered from the 

balloon, earth, moon or clouds 

from contributing to solar cell 

current . Two motors and 

supporting electronics control the 

altitude and bearing angles of the 

collimator during a flight. A video 

camera is mounted on top of the 

package and used to observe the 

operation of the Suntracker as it 

ascends though low temperatures to the stratosphere. The weight of the payload is about six pounds. 

Photographs showing the system components inside the scientific package may be viewed on the Suntracker 

Web site [33]. 

Figure 3. Suntracker scientific package developed for characterizing 

multi-junction solar cells in the stratosphere.

Table 2. Summary of Suntracker Flights 

ST Date Launch Site Landing Site Alt. (ft) Range (mi) 

 I 8/29/99 Findlay, OH Marion, OH 87,000 40

II 9/9/00 Findlay, OH Aborted due to hardware problems

III 9/14/02 Findlay, OH Greenwich, OH 94,256 55 

IV 10/12/02 Findlay, OH Brunswick, OH 87,127 94

V 11/9/02 Portland, IN Aborted due to surface winds

VI 12/7/02 Portland, IN Canton, OH 95,956 186 

VII 8/16/03 Findlay, OH Amlin, OH 99,500 68

The scientific package is attached to a parachute that is affixed to a latex balloon. The balloon weighs about 

three pounds and is inflated with about 250 cubic feet of helium and launched by a three-person team. The 

Suntracker scientific package is tethered below the balloon and the electronics is programmed to point the 

collimator at the sun during ascent. The balloon ascends at a rate of about 800 ft/min to a burst altitude ranging 

between about 87 and 100 kft in about two hours and then parachutes to a landing site. Data are downlinked 

continuously during the flight on 2 m and 70 cm bands to a mobile unit that is equipped with receivers, 

computers and tracking software. The three-person team in the mobile unit chases the balloon, records the

data and retrieves the scientific package. The data includes cell voltage, cell temperature, electronics module 

temperature, video, reference voltage, and atmospheric pressure. The short-circuit current is determined from 

the cell voltage across a 1.000 Ω precision resistor that is in series with the cell. 

RESULTS 

Seven flights have been attempted with five successful launches. Table 2 shows the launch dates and 

locations, burst altitudes, landing sites and balloon trajectory ranges. The scientific package was retrieved on 

the same day for the Suntracker I, III and IV flights. Hardware problems developed during the Suntracker VI 

and VII flights that resulted in the loss of GPS signals; the package was found within a few days of the launch by 

individuals and subsequently retrieved. A single-junction silicon solar cell was mounted in the collimator during 

NASA/CP—2005-213431 161
the flights. The cell voltage data downlinked during the Suntracker IV and VI flights have been analyzed using 

Langley plots to determine the AM0 

short-circuit current. The Suntracker IV 

uncorrected short-circuit current versus 

altitude is shown in Figure 4. Only the 

maximum currents were selected for use

in the Langley plot. The cell current data 

illustrate the tracking characteristics 

during the ascent. For the most part the

Suntracker was not locked on the sun 

during the flight. The video data showed

the motors slowed down during the 

ascent as a result of the low atmospheric 

temperatures. Motor assemblies using 

lubricant with lower temperature 

specifications will be evaluated in future 

flights. Additionally, the stability of the 

scientific package and the collimator 

control algorithm will be investigated in 

order to improve the performance of the

Suntracker system. 

ST VI Cell Current vs. Altitude

0

40

80

120

160

0 20 40 60 80 10

Altitude (1000 ft)

C

urre

nt

0

(mA)

Figure 4. Suntracker VI short-circuit current versus altitude for a 

single-junction silicon solar cell. The cell temperature versus altitude 

during the Suntracker VI flight is shown 

in Figure 5. Also shown are the 

radiosonde data reported by the 

National Weather Service (NWS) on the 

day of the flight. The effect of solar 

heating on the cell current is apparent . 

While the solar cell temperature 

increased from –10 to about 0 o

C as the 

balloon ascended from 80 to 96 kft, the 

atmospheric temperature remained at 

about –45 oC. The dependence of the 

cell and NWS temperatures in this 

altitude range suggests that the cell 

temperature may be higher at higher 

altitudes. If this is the case, it will be 

possible to operate cells closer to 25 o

C 

at higher altitudes, and to determine the

temperature coefficient of the shortcircuit current as the package ascends. 

Figure 6 is a Langley plot of the data for 

a single-junction silicon solar cell from 

the Suntracker IV and VI flights. The 

optical air masses were calculated using

Equation 11. The data have been corrected for the earth-sun distance and cell temperature, and fit with straight 

lines. The extrapolated AM0 short-circuit currents are 144.32 and 144.38 mA for the Suntracker IV and VI 

measurements, respectively. The average AM0 short-circuit current is 144.35 mA ± 0.02 %. The resolution of

the eight-bit ADC in the Suntracker data acquisition system is ± 0.2 %, showing that the agreement between the 

two flights is better than the uncertainty in the measurements and probably reflects the statistics of the curve

fitting etc. The AM0 short-circuit current of the single-junction silicon solar cell flown on the Suntracker flights 

was determined using the aircraft method at NASA Glenn Research Center [21]. The AM0 short-circuit current 

was 144.88 mA and within ± 0.36 % of the Suntracker average value. The results agree to within the statistics 

Temperature vs. Altitude

-70

-50

-30

-10

10

30

0 20 40 60 80 100

Altitude (1000 feet)

Temp. (Celsius)

NWS Cell

Figure 5. Suntracker VI downlinked solar cell and atmospheric 

NWS temperatures versus altitude. 

NASA/CP—2005-213431 162
of the two methods, namely about ± 0.2 % for the Suntracker measurements and ± 0.6 % for the aircraft 

method. 

It is instructive to determine the 

atmospheric optical absorption 

coefficients for the two methods using 

Equation 9. The slopes of the two 

straight lines in Figure 6 were analyzed

to determine the absorption

coefficients; the coefficients are 0.265 

and 0.293 per air mass for the 

Suntracker IV and VI data, respectively. 

The average value of the atmospheric

optical absorption coefficient is 0.280 

per air mass 5 %. An analysis of the 

Langley plot produced with aircraft data

gives absorption coefficients of 0.125 

per air mass. The Suntracker value is 

somewhat larger than the 0.20 per air 

mass determined from the earlier 

aircraft measurements [13] while the 

absorption coefficients determined with 

the current aircraft data is considerably 

less. The reasons for these differences

are not understood and will be the subject of future investigations. 

±

Figure 6. Langley plot of corrected Suntracker IV and VI data for a 

single-junction silicon solar cell. 

CONCLUSIONS 

The voltage dependence of the spectral responses of multi-junction solar cells complicates optimization of cell 

design. The series nature of multi-junction solar cells places more demands on the need for standard cells 

characterized under AM0 conditions. Cells selected for use as standards should have a history of thermal 

cycling and light soaking that provides evidence of the level of cell stability. The AM0 short-circuit current of a 

single-junction silicon solar cell was determined using data collected during two Suntracker flights. The 

agreement in the two measurements was 0.02 %. The agreement in the AM0 short-circuit current of the cell

measured with the Suntracker balloon method and NASA Glenn Research Center aircraft method was 

±

± 0.36

%, which is within the uncertainty of the two methods. There is a need to understand the role of ozone and 

atmospheric optical absorption on the calibration of solar cells in the stratosphere. 

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NASA/CP—2005-213431 163
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19. http://rredc.nrel.gov/solar/spectra/am0/NewAM0.xls 

20. B. Leckner, “The Spectral Distribution of Solar Radiation at the Earth’s Surface – Elements of a Model,” 

Solar Energy, volume 20, page 143, 1978. 

21. David B. Snyder, Philip P. Jenkins and David A. Scheiman, “Historical Precision of an Ozone Correction 

Procedure for AM0 Solar Cell Calibration,” Proceedings of the Space Photovoltaic Research and 

Technology Conference, 2003, In Press. 

22. Richard D. McPeters, P. K. Bahartia, Arlin J. Krueger, Jay R. Herman, Charles G. Wellemeyer, Colin J. 

Seftor, Glen Jaross, William Byerly, Steven L. Taylor, Tom Swissler and Richard P. Cebula, “Earth 

Probe Total Ozone Mapping Spectrometer (TOMS) Data Products User’s Guide,” NASA Technical 

Publication 1998-206895, page 53, 1998. 

23. IBID, page 53. 

NASA/CP—2005-213431 164
24. John A. Zoutendyk, “The Space Calibration of Standard Solar Cells Using High Altitude Balloon 

Flights,” Proceedings of the Fourth Photovoltaic Specialists Conference, page C-4-1, 1964. 

25. B. E. Anspaugh and R. L. Mueller, ”Results of the 2002 JPL Balloon Flight Solar Cell Calibration 

Program,” JPL Publication 02-03, 2002. 

26. IBID, page 20. 

27. IBID, page 22. 

28. IBID, page 26. 

29. IBID, page 1. 

30. Glenroy A. Bowe, Qianghua Wang, James R. Woodyard, Richard R. Johnston and William J. Brown, 

“Investigations to Characterize Multi-Junction Solar Cells in the Stratosphere Using Low-Cost Balloon 

and Communication Technologies,” Proceedings of 16th NASA Space Photovoltaic Research and 

Technology Conference, August 31-Septermber 2, 1999, page 189. 

31. Glenroy A. Bowe, Qianghua Wang and James R. Woodyard, ”Investigations to Characterize Multijunction Solar Cells in the Stratosphere Using Low-Cost Balloon and Communication Technologies,” 

Twenty-Eight IEEE Photovoltaic Specialists Conference Proceedings, 2000, page 1328. 

32. Ali Mirza, David Sant, James R. Woodyard, Richard R. Johnston and William J. Brown, “Report on 

Project to Characterize Multi-junction Solar Cells in the Stratosphere Using Low-Cost Balloon and 

Communication Technologies,” Seventeen Space Photovoltaic Research and Technology Conference, 

2001, page 137. 

NASA/CP—2005-213431 165
Historical Precision of an Ozone Correction Procedure for AM0 Solar Cell

Calibrations

David B. Snyder, NASA Glenn Research Center

Phillip Jenkins, Ohio Aerospace Institute

David Scheiman, Ohio Aerospace Institute

Abstract

In an effort to improve the accuracy of the high altitude aircraft method for calibration of

high band-gap solar cells, the ozone correction procedure has been revisited. The new

procedure adjusts the measured short circuit current, Isc, according to satellite based

ozone measurements and a model of the atmospheric ozone profile then extrapolates the

measurements to air mass zero, AM0. The purpose of this paper is to assess the

precision of the revised procedure by applying it to historical data sets. The average Isc of

a silicon cell for a flying season increased 0.5% and the standard deviation improved from

0.5% to 0.3%. The 12 year average Isc of a GaAs cell increased 1% and the standard

deviation improved from 0.8% to 0.5%. The slight increase in measured Isc and

improvement in standard deviation suggests that the accuracy of the aircraft method may

improve from 1% to nearly 0.5%. 

Introduction

The NASA GRC High Altitude Aircraft Method

has been used to provide the aerospace

industry with solar cells calibrated to orbital

conditions for nearly 40 years. The method

measures the short circuit current, Isc, with an

accuracy of 1% for setting solar simulators to

space conditions during ground-based

measurements (ref.1). However, with the

increased application of high band-gap and

multijunction solar cells in space, it has

become necessary to reassess and revise the

method in order to continue to provide

measurements correctly adjusted to air mass

zero, AM0. 

Aircraft based measurements are made at the

lower edge of the stratospheric ozone layer

(figure 1) where the ozone density is

increasing with altitude. The Langley Plot

method is used to extrapolate flight

measurements to AM0. It assumes that the

adsorption characteristics of the atmosphere

are proportional only to pressure. The

extrapolation will therefore be offset due to

the non-uniformity of the ozone density. This

has been accounted for by increasing the

measured Isc by 1% based on an estimate of

the influence of ozone absorption on silicon

solar cells (ref.1). This procedure must now

be adjusted for higher band gap solar cells. 

Initial work indicated that it was not sufficient

to use the sun angle with the ozone number to

calculate a correction factor (ref. 2) Instead

correcting the measured Isc according to an

ozone profile model, then performing the

Langley extrapolation, looked promising. This

paper applies the technique to historical data

to see how it improves the precision and,

presumably, the accuracy of the flight

measurements.

Figure 1. Ozone profiles for various total

ozone numbers.

0

5

10

15

20

25

30

35

40

0 5 10 15 20 25

Altitude (km)

Ozone Density (d.u./km)

Ozone Density Profile

Mid Latitude

from TOMS Version 7 Standard Ozone Profiles

Lear Max Altitude 15.25 km

Total Ozone Number

225 du

275 du

325 du

375 du

425 du

NASA/CP—2005-213431 166
Procedure

A new ozone correction procedure has

recently been developed (ref 2). The first step

in the new procedure is to use spectral

response data to calculate an ozone

correction factor, Foz, for each cell technology.

The measured Isc is then adjusted by a factor,

1+[O3]*Foz, where [O3] is the Ozone column

along the optical path. Examples of Foz are

shown in Table 1 for various types of solar

cells. Next, based on satellite data available

the day after a calibration flight, an ozone

profile is produced and used to correct the

measured Isc as a function of air mass and

altitude. The corrected Isc is then

extrapolated via the Langley Plot method to

AM0. 

Cell type Foz (/du)

Silicon 29.6x10-6

α Silicon 50.4x10-6

GaAs 42.6x10-6

InGaP top junction 57.2x10-6

GaAs middle junction 26.3x10-6

Ge bottom Junction 2.43x10-6

Table 1. Ozone correction factors for various

solar cell types.

The Foz calculation is based on the

convolution of the ozone absorbed WM0

spectra with the spectral response data for a

solar cell type as illustrated in figure 2. The

ozone absorption region of the spectra that

most affects solar cell performance is from

about 500 nm to 800 nm. This means that

high band gap cells tent to be affected more

than low band gap cells.

Next each Isc measurement during the flight is

corrected using Foz and the ozone along the

optical path by a factor of 1+[O3]Foz. The total

Ozone number, On, for the flight date and

location is available from the Earth Probe

satellite at the web site in reference 3. This is

then adjusted by an ozone density model for

the altitude, f(p),as given by the pressure of

the Isc measurement illustrated as Model 2 in

figure 3. The ozone density model is based of

the TOMS Version 7 Standard Ozone Profile

as described in appendix A of reference 4.

Finally , the optical path is corrected for the

sun altitude, θ, approximately 1/sin(θ). So,

[O3] = On * f(p)/sin(θ).

Results

This paper reports on data from three solar

cells. First, is a silicon solar cell that has been

flown since 1985. It is used for two sets of

data, first it a short term set of data set with 20

flights over a single flying season, second is a

selection of 13 flights over an eight year

period. The second cell examined is a GaAs

cell. Selected flight data from 31 flights over

ten years are examined. Finally, data from an

InGaP top cell which was flown seven times in

Figure 2. Ozone corrected irradiance and

solar cell quantum efficiencies.

0

500

1000

1500

2000

2500

0.2 0.4 0.6 0.8 1 1.2 1.4

Irradiance (W/sq m/um)

Quantum Efficiency

Wavelength (um)

Ozone Corrected Spectra 

 and Cell Quantum Efficiencies

0.0 du

600 du

1200 du

Silicon

GaAs

InGaP

Figure 3. Pressure dependent ozone density

model.

0

0.5

1

1.5

2

2.5

3

3.5

0 100 200 300 400 500 600

Ozone Density (du/mb)

Pressure (mb)

Ozone Density Model

model 1: 225 du

325 du

425 du

525 du

model 2: 225 du

325 du

425 du

525 du

NASA/CP—2005-213431 167
recent flight seasons are presented.

Si cell short term

The silicon solar cell A-161 has flown

frequently on the Lear 25 as a monitor cell

during calibration flights. It flew 20 times

during the 2000-2001 flying season. Using the

1% ozone correction method, the short circuit

current, Isc, measurements averaged

165.60 mA with a standard deviation of

0.81 mA. With the new ozone correction

method the average Isc was 166.46 mA with a

standard deviation of 0.43 mA. Figure 4

shows a comparison of the measurements

plotted against sun angle during the

measurement. The reduction in Isc at lower

Sun altitudes using the 1% correction method,

is interpreted as due to the increased optical

path length through the ozone layer. The new

correction method nearly removes this

dependence. 

Si cell long term

Flight data is also available for A-161 since the

1985 flying season. Figure 5 shows some of

the data over the interval 1985-1993. The

ozone data for this period is available from the

Nimbus 7 satellite (ref. 5). Isc from 13 flights

over that interval averaged 166.0 mA with a

standard deviation of 1.2 mA. Using the new

procedure Isc averaged 166.9 mA with a

standard deviation of 0.8 mA.

GaAs cell long term

A-133 is a GaAs cell that has flown frequently.

Data was examined that was obtained in the

period 1985-1994. Of 31 flights the old

procedure gave a average of 110.16 mA with

a standard deviation of 0.83 mA. The new

ozone correction procedure gives an average

of 111.29 mA with a standard deviation of

0.52 mA. A comparison is shown in figure 6.

InGaP Top Cells

134-5-6 is an InGaP Top Cell, part of a pair of

subcells for a dual junction solar cell. It has

been flown several times over the past few

Figure 4. Variation in Isc(AM0) of a silicon

solar cell as a function of sun angle during a

flight season.

163

164

165

166

167

168

20 25 30 35 40 45

Isc (mA)

Sun Altitude (degrees)

A-161 Short Circuit Current

Comparing Ozone Correction Procedures

2000-03 Flight season

New Ozone Correction

 1% correction

Figure 5. Long term comparison of Isc(AM0)

of a silicon cell as a function of sun angle.

Figure 6. Long term variation in Isc

measurements of a GaAs solar cell.

108

109

110

111

112

113

20 25 30 35 40 45 50

Isc (mA)

Sun Altitude (degrees)

A-133 GaAs Short Circuit Current

Comparing Ozone Correction Procedures/n1985-1994

New Ozone Correction

1% Correction

163

164

165

166

167

168

169

20 25 30 35 40

Isc (mA)

Sun Altitude (degrees)

A-161 Short Circuit Current

Comparing Ozone Correction Procedures

1985-1993

New Ozone Correction

 1% Correction

NASA/CP—2005-213431 168
years. Figure 7 compares results from seven

flights in the 2000-2001 and 2002-2003

seasons. The 1% correction procedure gave

an average of 60.68 mA with a standard

deviation of 0.46 mA. The new method gives

an average of 62.08 mA and a standard

deviation of 0.14 mA. 

Discussion and Conclusion

This paper has compared an old procedure for

correcting Isc measurements for atmospheric

ozone with a new one. Table 2 shows the

changes due to this method. The slight

increase in Isc is expected as the procedure

accounts for the increased optical path due to

low sun altitude. In addition, the correction

increases with band gap. 

Cell Type Isc

change 

Standard

Deviation

1% new

Si short

term 0.52% 0.49% 0.26%

Si long

term 0.54% 0.72% 0.48%

GaAs long

term 1.02% 0.75% 0.47%

InGaP 2.26% 0.76% 0.23%

Table 2. Relative changes is Isc

measurements.

The standard deviation of each set of

measurements also improves by at least one

third because the dependence on sun angle is

nearly removed. This suggests that the

influence of atmospheric ozone has been an

important uncertainty in the aircraft based

measurements. The new correction procedure

appears to produce a significant improvement

in measurements. However, in order to have

complete confidence in the procedure, results

still must be compared to high altitude balloon

measurements.

The new procedure has been used at NASA

GRC for measurements beginning with the

2002-2003 flight season. It is anticipated that

in the 2003-2004 season the capability of

making a spectroscopic measurement of the

ozone profile will refine these calibration

measurements further.

References

[1] P.Jenkins, D. Brinker and D. Scheiman,

“Uncertainty Analysis of High Altitude Air Mass

Zero Solar Cell Calibration “, 26th IEEE PVSC,

V26-206 ,1997, pp.857-860..

[2] D. Snyder, et al, “Ozone Correction for

AM0 Calibrated Solar Cells for the Aircraft

Method “, 29th IEEE PVSC,2002.

[3] <http://toms.gsfc.nasa.gov/eptoms/

 ep.html>

[4] McPeters, R.D. et al, Earth Probe Total

Ozone Mapping Spectrometer (TOMS) Data

Products Users Guide, NASA TP

1998-206895, 1998.

[5] <http://toms.gsfc.nasa.gov/n7toms/ 

 nim7toms.html>

Figure 7. Isc of an InGaP cell as a function of

sun angle.

60

60.5

61

61.5

62

62.5

20 25 30 35 40

Isc (mA)

Sun Altitude (degrees)

134-5-6 InGaP Top Cell Short Circuit Current

 Comparing Ozone Correction Procedures

2000-2003

New Ozone Correction

1% Correction

NASA/CP—2005-213431 169
EVALUATION OF THIN-FILM SOLAR CELLS FOR SPACE APPLICATIONS 

S. H. Liu, E. J. Simburger, J. Matsumoto, A. Garcia III, J. Ross, and J. Nocerino 

The Aerospace Corporation, Los Angeles, CA 90009-2957 

INTRODUCTION 

Photovoltaic devices used on spacecraft have requirements that include high solar-to-electric conversion efficiency with low mass (high specific power), low volume (high specific volume), tolerance of space radiation damage, and stable performance through thermal cycling [1]. Current spacecraft use high-efficiency, single-crystal Si 

and III-V semiconductor solar cells as the source of power generation. With the development of lightweight flexible thin-film solar cells for terrestrial applications, there is an interest in transitioning this technology to space 

applications. Thin-film solar cells have many attractive advantages over crystalline solar cells. The low mass of 

the thin-film solar cells increases the specific power by an order of magnitude from the current level of 65 W/kg[2]. 

Figure 1 compares specific power versus conversion efficiency for various solar cell technologies. The maximum 

specific power achievable with crystalline solar cell technology is in the range of 200 W/kg. If thin-film solar cell 

technology efficiency improves from the current 5%–8% range to a 15%–18% range, the specific power for the 

array would double from the current 200 W/kg to 400 W/kg.

A part of the process of transitioning thin-film solar cell technology to space applications is to assess the solar cell 

performance in a simulated space environment. While in geosynchronous orbit, thin-film solar cells will experience extreme temperatures from –160°C to +100°C within a few minutes after eclipse. The manufacturers of 

thin-film solar cells for terrestrial applications have not measured the temperature coefficient over this temperature 

range. The impact of the operational temperature on the performance of thin-film solar cells must be taken into 

account in the design of the spacecraft power system. In this paper, we present the measured temperaturedependent performance data on thin-film solar cells from the following vendors: Iowa Thin-film Technologies 

(ITFT), Uni-Solar Systems Corp., and MicroSat Systems Inc. 

EXPERIMENTAL 

The experiment consisted of measuring the I-V characteristic of sample thin-film solar cells at temperatures 

between 15°C and 100°C. The temperature coefficients of the ITFT thin-film, dual-junctions amorphous silicon (aSi) on polyimide substrate, Uni-Solar triple-junctions of a-Si and a-Si/Ge on 1-mil stainless-steel substrate, and 

MicroSat Systems’ single junction CuInGaSe2 (CIGS2) on 1-mil stainless-steel substrate solar cells was 

measured. Measurements were conducted on Uni-Solar and MicroSat Systems solar cells as provided by the 

manufacturers. Two changes were made to the ITFT cells. First, a high-conductance silver front contact was 

deposited onto the solar cells by Aerospace rather than a printed ink contact normally applied by ITFT for 

terrestrial applications. The second change is the use of aluminum instead of stainless steel as a method of 

bleeding off static charge buildup during the roll-to-roll deposition process at ITFT. The aluminum layer on the 

back of the solar cell was etched away, exposing the surface of the 2-mil-thick polyimide substrate. Three cells 

from each manufacturer were used in making these measurements. The dimensions of the Uni-Solar cells are 5 

cm × 26.08 cm with an active area of 130.4 cm2 (Figure 2). The ITFT hexagon solar cells have an active area of 

347 cm2

 (Figure 3). MicroSat Systems solar cell dimensions are 7.25 cm × 1.65 cm with an active area of 75 cm2

(Figure 4). 

NASA/CP—2005-213431 170
The temperature coefficient measurements of the thin-film solar cells, within the temperatures range of 15°C to 

100°C, were conducted inside a Plexiglas box that houses a temperature-controlled vacuum plate. The Plexiglas 

housing allows purging of the system with dry nitrogen. The vacuum plate ensures maximum thermal contact 

between the plate and the solar cell, and reduces any thermal gradient in the solar cell. The vacuum plate also 

serves as a thermal reservoir that maintains temperature stability during the performance evaluation. A temperature bath that uses coolant as a heat exchange source controls the temperature of the vacuum plate. 

A Spectrolab X-25 Solar Simulator calibrated for AM0 conditions provided the necessary insolation in performance characterization of the thin-film solar cells. Figure 5 shows the spectra of the sun at AM0, the X-25 Solar 

Simulator, and X-25 with Plexiglas that houses the vacuum plate. 

Table 1. The I-V Characteristics of Thin-Film Solar Cells Provided 

by ITFT, Uni-Solar, and MicroSat Systems at 25°C.

Sample Voc (V) Isc (A) Pmax (W) Eff. (%) 

ITFT #1 5.1414 0.7139 1.6188 3.45 

ITFT #2 4.9329 0.7155 1.4435 3.07 

ITFT #3 5.1481 0.7157 1.5906 3.39 

Uni-Solar #1 2.2220 1.0249 1.4651 8.30 

Uni-Solar #2 2.2153 0.9924 1.4194 8.05 

Uni-Solar #3 2.2181 0.9933 1.4028 7.95 

MicroSat #1 0.5213 2.3933 0.6512 6.42 

MicroSat #2 0.4560 2.5250 0.6356 6.26 

MicroSat #3 0.4892 2.2216 0.5937 5.85 

RESULTS 

Temperature coefficients of the open-circuit voltage (Voc) and short-circuit current (Isc) for each solar cell were 

determined by averaging three separate measurements. The Voc and Isc temperature coefficients were obtained 

from the least-square linear regression of the raw data. The conversion efficiency temperature coefficient was 

determined based on the normalized data. Table 1 shows the I-V characteristics of various solar cells measured 

at 25°C. The results are grouped according to manufacturer. 

Uni-Solar 

 

The temperature dependence of Uni-Solar I-V characteristics is shown in Figures 6–8 and Table 2. Both Voc and 

Isc showed a linear temperature dependence of –0.0080±0.0002 V/°C and 0.000877±0.00002 A/°C, respectively. 

The coefficient of efficiencies are –0.14±0.04 %/°C over the temperature range of 20°C–40°C, –0.35±0.01%/°C 

over the range of 50°C–70°C, and –0.45±0.06 %/°C over the range of 70°C–100°C. 

Table 2. Temperature Coefficients of Uni-Solar a-Si 

Dual-junction Thin-film Solar Cells 

I-V Characteristics Temperature Coefficients 

Voc (V) 

15°C–100°C –0.0080 V/°C 

Isc (A) 

15°C–100°C 0.000877 A/°C 

Conversion 

Efficiency (%) 

20°C–40°C –0.14 %/°C 

50°C–70°C –0.35 %/°C 

70°C–100°C –0.45 %/°C 

Iowa Thin-Film Technology 

NASA/CP—2005-213431 171
The temperature dependence of Iowa Thin-film Technology a-Si solar cells I-V characteristics is indicated in Figures 9–11 and Table 3. Over the temperature range of 15°C to 100°C, Voc showed an average temperature 

coefficient of –0.0133±0.0006 V/°C, and Isc showed a linear dependence of 0.00063±0.00003 A/°C. The temperature coefficients of cell efficiency are 0.14±0.01% over the temperature range of 25°C to 45°C, 

-0.05±0.01%/°C over the temperature range of 45°C to 60°C, and –0.14±0.04%/°C over the temperature range of 

60°C to 100°C. Between 15°C and 25°C, measurements of the cell efficiency showed inconsistent results. Two 

measurements indicated a negative temperature coefficient, but the third one showed otherwise. The cause of 

this discrepancy is unclear, and it will be further investigated in the future. 

Table 3. Temperature Coefficients of Iowa Thin-Film 

Technology a-Si Dual-Junction Thin-film Solar Cells

I-V Characteristics Temperature Coefficients 

Voc (V) 

15°C–100°C –0.0133 V/°C 

Isc (A) 

15°C–100°C 0.00063 A/°C 

Conversion 

Efficiency (%) 

25°C–45°C 0.14 %/°C 

45°C–60°C –0.05 %/°C 

60°C–100°C –0.14 %/°C 

MicroSat Systems 

Figures 12–14 and Table 4 present the I-V characteristics of MicroSat CIGS2 solar cell. The I-V characteristics of 

Voc, Isc, and efficiency showed linear temperature dependence from 15°C to 100°C. Between 15°C and 100°C, 

the Voc, Isc and efficiency temperature coefficients are –0.0024±0.0001V/°C, –0.00090±0.00006 A/°C, and 

-0.69±0.03%/°C, respectively. 

Table 4. Temperature Coefficients of MicroSat 

Systems CIGS2 Thin-Film Solar Cells

I-V Characteristics Temperature Coefficients 

Voc (V) 

15°C–100°C –0.0024 V/°C 

Isc (A) 

15°C–100°C -0.00090 A/°C 

Conversion 

Efficiency (%) 

15°C–100°C –0.69 %/°C 

NASA/CP—2005-213431 172
SUMMARY 

Lightweight, flexible, thin-film solar cells present many promising applications in space power systems. Thin-film 

photovoltaic devices offer attractive advantages in terms of high specific power and high specific volume over traditional crystalline solar cells. The assessment of cell performance must be conducted in simulated space conditions in order to apply commercially available terrestrial thin-film solar cells for space missions. In this paper, we 

have reported on the experimental measurements of temperature coefficients for Voc, Isc, and conversion efficiency in the temperature range of 15°C to 100°C for three different types of thin-film solar cells. This represents 

the initial effort in measuring the temperature coefficients of thin-film solar cells I-V characteristics over the 

extreme temperatures that spacecrafts experience in space. We are in the process of building a vacuum chamber that will allow solar cell temperature to reach as low as –160°C (Figure 15). Results will be presented in 

future publications. Our effort in characterizing solar cell performance allows assessment of current thin-film solar 

cell technology for space applications, and provides a basis for spacecraft design. 

REFERENCES 

1. P. Iles, “Future of Photovoltaics for Space Applications,” Prog. in Photovo.: Res. Appl. 2000; 8:39-51. 

2. N. Dhere, S Ghongadi, M. Pandit, A. Jahagirdar, and D. Scheiman, “CIGS2 Thin-Film Solar Cells on Flexible 

Foils for Space Power,” Prog. in Photovo.: Res. Appl. 2002; 10:407–416. 

NASA/CP—2005-213431 173
Figure 1. Specific power of lightweight solar array with thin 

film, thin crystalline silicon, and crystalline GaAs solar cells 

as a function of solar efficiency. 

Figure 2. Uni-Solar System Corp. a-Si solar cell. 

Figure 3. Iowa Thin-film Technology hexagon a-Si solar cell. 

NASA/CP—2005-213431 174
Figure 4. MicroSat Systems Inc. CIGS2 solar cell. 

Figure 5. Spectra of sunlight at AM0 condition, X-25 Solar 

Simulator, and X-25 Solar Simulator with Plexiglas. 

Figure 6. Open-circuit voltage vs. measurement temperature for a-Si triple-junction Uni-Solar thin-film solar cells. 

NASA/CP—2005-213431 175
Figure 7. Short-circuit current vs. measurement temperature for a-Si triple-junction Uni-Solar thin-film solar cells. 

Figure 8. Conversion efficiency vs. measurement temperature for a-Si triple-junction Uni-Solar thin-film solar 

cells. 

Figure 9. Open circuit voltage vs. measurement temperature for a-Si dual-junction Iowa Thin-film Technology thinfilm solar cells. 

NASA/CP—2005-213431 176
Figure 10. Short-circuit current vs. measurement temperature for a-Si dual-junction Iowa Thin-film Technology 

thin-film solar cells. 

Figure 11. Conversion efficiency vs. measurement temperature for a-Si dual-junction Iowa Thin-film Technology 

thin-film solar cells. 

Figure 12. Open circuit voltage vs. measurement temperature for CIGS2 Micro-Sat Systems thin-film solar cells. 

NASA/CP—2005-213431 177
Figure 13. Short circuit current vs. measurement temperature for CIGS2 Micro-Sat Systems thin-film solar cells. 

Figure 14. Conversion efficiency vs. measurement temperature for CIGS2 Micro-Sat Systems thin-film solar cells. 

Figure 15. New vacuum chamber that will be used in 

future I-V temperature coefficients measurement. 

NASA/CP—2005-213431 178
DEVELOPMENT OF CIGS2 THIN FILMS ON ULTRALIGHTWEIGHT FLEXIBLE LARGE AREA FOIL

SUBSTRATES

Neelkanth G. Dhere, Vivek S. Gade, Ankur A. Kadam, Anant H. Jahagirdar, Sachin S. Kulkarni and 

Sachin M. Bet 

Florida Solar Energy Center, Cocoa, FL 32922, U.S.A. 

Introduction 

The development of thin film solar cells is aimed at reducing the costs for photovoltaic systems [1]. 

Use of thin film technology and thin foil substrate such as 5-mil thick stainless steel foil or 1-mil thick Ti 

would result in considerable costs savings. Another important aspect is manufacturing cost. Current 

single crystal technology for space power can cost more than $ 300 per watt at the array level and weigh 

more than 1 kg/m2

 equivalent to specific power of ~ 65 W/kg. Thin film material such as CuIn1-xGaxS2

(CIGS2), CuIn1-xGaxSe2-ySy (CIGSS) or amorphous hydrogenated silicon (a-Si:H) may be able to reduce 

both the cost and mass per unit area by an order of magnitude. Manufacturing costs for solar arrays are 

an important consideration for total spacecraft budget. For a medium sized 5kW satellite for example, the 

array manufacturing cost alone may exceed $ 2 million. Moving to thin film technology could reduce this 

expense to less than $ 500K [2]. Earlier publications have demonstrated the potential of achieving higher 

efficiencies from CIGSS thin film solar cells on 5-mil thick stainless steel foil as well as initial stages of 

facility augmentation for depositing thin film solar cells on larger (6”x 4”) substrates [3]. This paper 

presents the developmental study of achieving stress free Mo coating; uniform coatings of Mo back 

contact and metallic precursors. The paper also presents the development of sol gel process, 

refurbishment of selenization/sulfurization furnace, chemical bath deposition (CBD) for n-type CdS and 

scrubber for detoxification of H2S and H2Se gases. 

 

Sol Gel Process 

A new sturdy and robust set up for carrying out sol-gel SiO2 coating by dip coating method was prepared at 

FSEC (Figure 1). A structure of adequate height was fabricated for large samples to be raised easily from the 

container (beaker of capacity 4 liters) holding the sol-gel solution. A DC gear motor of shaft diameter = 

0.312”(0.79 cm) with 4 rpm and a pulley assembly for raising and lowering the sample were mounted atop 

the structure. A 12V DC, 1500 mA power supply was used to control the motion of the DC motor. Special 

bushings of various dimensions were designed and fabricated for attaching to the motor for variable 

speeds for carrying out experiments with speeds ranging from 8-16 cm/min. The speed of withdrawal was 

one of the important factors and attempts were made to optimize the speed for improving the quality of 

the coating. Weights attached to the bottom of the foil kept it vertical so that the foil enters the sol-gel 

solution and is withdrawn normal to the sol surface. This prevents the thickness variation during coating. 

After withdrawal of the foil, it was left hanging to dry for a period of 5 minutes. Later the foil was placed in 

an Isotemp Muffle Furnace for sintering along with other foils. A batch of four foils could be heat-treated at 

a time in the furnace in one run. Adequate care was taken to place the foils in a specially prepared fixture 

without bending or stressing. Samples were sintered in air at a temperature of 450oC for a period of 1 

hour. After cooling for approximately 2 hrs the samples were removed from the furnace, rinsed with 

deionized water and dried with compressed N2 gas prior to storing them in a desiccators. Atomic Force 

Microscopy (AFM) was carried out at National Renewable Energy Laboratory to study the morphology 

and roughness of the SiO2 layer on Ti foil substrate at a high resolution. The images were obtained for an 

area of (100 X 100) μm2

. 

NASA/CP—2005-213431 179
Figure 1. Sol gel setup 

AFM images show both the 3-dimensional and normal 2-dimensional images. Figure 2 shows both these 

images for a plain Ti foil surface without any SiO2 deposition. The average roughness value obtained was 

Ra = 543 Å. Figure 3 shows Ti foil coated with SiO2 where the average roughness value was measured to 

be 146 Å [4]. 

Figure 2: AFM images (100 X 100 μm2

) of plain Ti foil, Ra = 543 Å. 

Figure 3: AFM images (100 X 100 μm2

) of Ti foil + 1 thick (0.55 μm) + 2 thin (0.35 μm) SiO2 layer, Ra = 

146 Å. 

NASA/CP—2005-213431 180
CIGS2 thin film was deposited on the foil after cleaning in distilled deionized water and drying with 

compressed N2 gas. Figure 4 shows a stacked layering sequence for the thin film deposition of CIGS2 on 

SiO2 coated titanium foil substrate. Ti was chosen because of its good compromise between cost and 

physico-chemical properties. 

Figure 4: Structure of completed Polycrystalline CIGS2 Thin film Solar Cell.

Mo Back Contact 

Depending on the working gas pressure, residual stresses are developed in refractory metal films 

prepared by magnetron sputtering. Films deposited below transition pressure develop compressive 

stresses, whereas those deposited above the transition pressure develop tensile stresses. Gross stress 

may be determined by visual inspection in that highly compressed films tend to buckle up, frequently in 

zigzag patterns, whereas films under extreme tensile stress develop a system of stress lines that appear 

like scratches. 

Thin flat strips of clean titanium foil of 1cm x 15cm were attached at their ends to a glass substrate 

prior to deposition (Figure 5) using vacuum compatible tape. It was observed that films deposited at high 

power of 300W and low argon pressure of 3 x 10-4 Torr resulted in compressive stresses. The thin foil 

deposited under these parameters showed convex bending (Figure 6). On the other hand films deposited 

at power of 200W and argon pressure of 5 x 10-3 Torr resulted in tensile stresses showing concave 

bending and stress lines (Figures 7-8) [5]. 

Figure 5: Thin strips of flat foil attached to glass substrate prior to deposition. 

 

NASA/CP—2005-213431 181
Figure 6: Films deposited at high power of 300W and low argon pressure of 3 x 10-4 Torr resulting in 

compressive stresses leading to lifting of the flat foil. 

Figure 7: Films deposited at low power of 200W and high argon pressure of 5 x 10-3 Torr resulting in 

tensile stresses leading to lifting of the inverted deposited foil. 

Figure 8: Scratches observed due to tensile stresses on films deposited 200W/ 5 x 10-3 Torr. 

A simple bending foil technique was used to estimate the magnitude of stress present in the foils. 

Loads were applied to the concave and convex region of the foil and weight required to make the foil flat 

similar to its position prior to deposition was measured. The load required and the curvatures in the cycles 

mentioned above are provided in Table I. 

NASA/CP—2005-213431 182
Table I: Loads applied to make foil flat. 

Power / Argon Pressure Height of curvature Loads applied 

300W / 3 x 10-4 0.88 mm 450 mg 

200W / 5 x 10-3 0.48 mm 130 mg 

The applied loads as well as the curvature of the foil were used for estimating the stress. These 

samples were sulfurized in 4% H2S in argon mixture at 475°C for 20 minutes. The reactivity of the 

molybdenum in sulfur/ selenium atmosphere has been reported to occur at temperatures over 600°C. 

However, it was observed that the sample under compressive stress peeled off from several regions and 

turned reddish pink indicating a chemical reaction occurring even at 475°C. On the other hand, the 

sample under tensile stress remained, mostly as it was with slight change in color from brownish to 

brownish red. 

 Foils were deposited with a number of combinations of the cycles of high power/low pressure and 

low power/high pressure to obtain a very flat foil after deposition. It is reported that such stress reversal is 

dependent on energetic bombardment by reflected neutral and/or sputtered atoms. At relatively low 

pressure, the mean free path is high, as a result of which the arriving atoms have higher kinetic energy 

and the resulting film has dense microstructure thereby experiencing compressive stresses. At higher 

working pressure the mean free path is reduced resulting in moderate flux of atoms reaching the 

substrate leaving the film in tensile stress. However, for producing highly efficient 4”x 4” size solar cell, the 

back contact molybdenum has to be fairly free of residual stresses. Therefore, the deposition started with 

low power and high pressure to reduce chemical reactivity. Two compressive stress cycles sandwiched 

between three tensile cycles were deposited in order to lower the stress and to achieve required 

thickness uniformity. 

Thickness Uniformity 

Molybdenum back contact and zinc oxide thin films were deposited on large-area glass substrates 

using optimized parameters of DC or RF sputtering power, working gas pressure, target-to-substrate 

distance, sputtering time and linear substrate movement parallel to the 4” width of the target. 

Figure 9. Thickness distribution profile. 

After depositions, the grid design initially drawn on the substrates together with the overlaying thin film 

was removed by ultrasonic cleaning in isopropanol, thus providing a distinct step for measurements with a 

Dektak3

 thickness profilometer. Thickness variation of ±11% over the central 4”x4” region was observed. 

Series of experiments were carried out for improving thickness uniformity by empirically modifying the 

magnetic field distribution. Experiments were carried out using various permutations and combinations of 

NASA/CP—2005-213431 183
nickel coated soft iron strips arrangements. The strips were replaced by similar-size rubber strips to 

maintain the shape integrity. The optimum configuration resulted in a thickness uniformity of ±3% over 

central 4” x 4” region for molybdenum thin films (Figure 9) while it continued to drop beyond this region 

[6]. 

Based on the continued thickness drop beyond 4”, it was realized that the magnetic field must be 

boosted at the extremities to achieve better thickness uniformity. The sputtering sources were also found 

to produce non-uniform plasma distribution within the racetrack leading to varying plasma intensity glow, 

again confirming the need to modify magnetic array. In the mean time, the manufacturer had modified the 

magnetic array with a better magnetic field distribution and stronger magnetic field at the extremities. Two 

such new magnetron-sputtering sources were procured. The old and new modified magnetic arrays are 

shown in Figures 10 and 11 respectively. 

Figure 10: Older Magnetic Array design 

Figure 11: Modified magnetic array. 

However, in the absence of specific information of their being substantially different, they were 

assumed to be similar to the older sources and no thickness profiling experimentation was carried out 

with them. The new magnetic array design and the necessary additional magnets were obtained from the 

manufacturer for upgrading the remaining three magnetron-sputtering sources. The design introduced 

extra magnets at and near the periphery. The copper body holding the magnetic assembly was machined 

to accommodate the modified assembly. The problem of non-uniform plasma was resolved after 

modifying magnetic assembly. The initial non-uniform plasma was modified into the dumbbell shape of 

the racetrack region with the modified magnetic arrays having boosted magnetic field at extremities 

(Figure 12 a, b). 

NASA/CP—2005-213431 184
Figure 12. a) Non-uniform plasma b) Uniform plasma. 

Figure 13 shows variation of thickness of molybdenum deposited by DC sputtering and zinc oxide by RF 

magnetron sputtering averaged over 2.75” distances along the linear movement direction with the length 

parallel to the 12” dimension of sputtering sources. The thickness uniformities over central lengths were: 

Molybdenum: ±2.24 % over 4”, ±2.40% over 5”, ±2.95% over 6” and zinc oxide: ±2.46 % over 4”, ±3.84% 

over 5”, ±5.60% over 6”. Fortunately the thickness uniformity for ZnO/ZnO:Al is slightly less critical. These 

results pave the way for deposition of large area (6”x4”) thin film solar cells. 

Figure 13. Thickness variation plots Mo (DC sputtering) and ZnO (RF sputtering). 

Selenization/Sulfurization Furnace 

Deposition of metallic precursors on the substrate is followed by sulfurization to get p-type 

CIGS2 absorber layer. A furnace was donated by Shell (formerly Siemens) Solar capable of 

Selenization/Sulfurization of metallic precursors on large 6”x 4” size substrates. The furnace is 

capable of sulfurizing one hundred 4”x4” or sixty-six of 6”x 4” substrates in a single run. The 

furnace in as received condition needed upgrade. The major part of upgrade included; the 

computer, mass flow controllers, valves, thermocouples and cables. The furnace is now fully 

operational and produces CIGS2 absorber layer on routine basis. 

NASA/CP—2005-213431 185
Figure 14. Selenization/Sulfurization furnace

CdS by Chemical bath deposition 

Having augmented the substrate size to 4”x 4”, next step was to design and construct chemical bath 

deposition (CBD) setup for four such substrates. The setup was built using teflon sheets of 1⁄2” and 1/8” 

thickness, and a teflon rods of 1/8” diameter. The setup can hold two1/8” thick glass samples and two 1-5 

mil thick stainless steel or titanium foil substrates. Holes were drilled in the bottom and top teflon plate to 

insert 1/8” diameter rods for holding the two top and bottom sheets assembly. 4 5/8” length of the rods 

allowed 4” clearance between top and bottom plates for placing 4” x 4” substrates. The bottom plate has a 

central circular hole for the stirrer movement. The top plate had openings for insertion of the 

thermometers and pH control sensors. The final assembly is shown in the figure 15. Experiments were 

carried out with the new designed set up, thus scaling up from 1.25” x 1” to 1” x 4” to 4” x 4” as shown in 

Figure 16. Table 2 provides the bath composition for carrying out deposition on 4”x 4” substrates. 

Table II. Chemical bath composition for 4” x 4” substrates 

Sr.No. Type of solution Concentration Quantity (milliliters) 

1 Distilled and Deionized water - 1320 

2 Ammonium Hydroxide (NH)4OH 28 to 30% 309 

3 Cadmium Sulfate CdSO4 0.015 M 181.5 

4 Thiourea 1.5 M 90.75 

Composition of the bath was chosen so that the substrates would be completely immersed inside the 

chemical bath. The bath is prepared in a beaker of 3500 ml capacity for the new set up of 4” x 4” 

substrates to fit snugly [7]. 

 

NASA/CP—2005-213431 186
Figure 15. CDS setup for 4”x 4” substrates 

Figure 16. Increase in substrate sizes from 1.25” x 1” to 1” x 4” to 4” x 4” 

Scrubber 

A scrubber was designed and fabricated to detoxify H2S and H2Se gases coming out of the furnace 

reaction tube. Reaction tube having a volume of 15.3 liters is filled with process gas [N2+H2S (<4%) or N2+ 

H2Se (<4%)] and heated to 500 0

C in a furnace. A mechanical pump pumped the process gas from the 

reaction tube through a mass flow controller. The effluent gas was further diluted with nitrogen by factor 

of 60. The process gas was fed through a venturi that is propelled by nitrogen gas used for dilution. The 

total flow of diluted hydrogen selenide (666 ppm) or diluted hydrogen sulfide (666 ppm) gas was 

calculated to be 132 liters/min equivalent to 4.7 scfm. The diluted process gas flows through the first wet 

scrubber and then through the first packed-bed column. The total height of the packed bed was 

calculated to be 6.9 feet for H2S in order to achieve less than 10 ppm at exhaust. The height of the total 

packed bed was calculated to be 17.14 feet for H2Se to achieve 50 ppb prior to exhaust according to 

Occupational Safety and Health Administration (OSHA). The packed-bed column was split in two. The 

height of the first column was 10.25 feet and the second was 9.25 feet. The extra height was provided to 

accommodate the packed bed column internals. Prior to entering the second packed-bed column, the gas 

passes through the second wet scrubber. The effluent gas from the packed-bed column(s) passes 

NASA/CP—2005-213431 187
through drums containing activated carbon impregnated with KOH and from there will be exhausted to the 

atmosphere [8]. 

The four point CM4 analyzer for H2Se has been installed, to detect the H2Se gas concentration at 

various levels. This CM4 analyzer has been connected to the present building alarm system in FSEC. 

 

 

Figure 17. Inside (left) and outside (right) scrubber assembly including CM4 connections 

Figure 18. Drums containing activated carbon impregnated with KOH and exhaust fans assembly 

Conclusion 

The paper demonstrates the potential of PV Materials Lab at the Florida Solar Energy Center (FSEC) 

for carrying out research and development of CuIn1-xGax

Se2-ySy (CIGSS) thin-film solar cells on large area 

ultralightweight flexible metallic foil for space applications. This facility equipped with large area deposition 

systems, sulfurization/selenization furnace along with CBD setup for CdS deposition and the scrubber unit 

is one of its kinds amongst University Labs and could serve as a nucleus of a small pilot plant for CIGSS 

thin film solar cell fabrication. 

NASA/CP—2005-213431 188
Acknowledgements 

This work was supported by the NASA Glenn Research Center, Contract # NCC 3712, and the 

National Renewable Energy Laboratory, contract no. NDJ-2-30630-03. 

References 

[1] Schock H. W and Pfisterer. F, “Thin Film Solar Cell: The Early Years”, 16th European photovoltaic solar 

energy conference, vol. 1 (2000), pp. 269-274 

[2] S. R. Ghongadi, “CuIn1-xGaxS2 (CIGS2) Thin Film Solar Cells On Stainless Steel Foil Space Power 

Application”, Master’s Thesis (University of Central Florida, Orlando) 2000 

[3] N. G. Dhere, S. R. Ghongadi, M. B. Pandit and A. H. Jahagirdar, “CIGS2 Thin Film Solar Cells On 

Flexible Foils For Space Power”, Proc. 17th Space Photovoltaic Research and Technology Conference, 

Cleveland, OH, Sept. 2001 

[4] V. S. Gade, “Development Of Copper Indium Gallium Disulfide CuIn1-xGaxS2 (CIGS2) Thin Film Solar 

Cells On Large Area Ultralightweight Titanium Foils Coated With Sio2 Barrier Layers”, Master’s Thesis 

(University of Central Florida, Orlando) 2002 

[5] S. M. Bet, “High Voltage Bias Testing Of Thin Film PV Modules, Adhesional Strength And Surface 

Analysis For PV Module Durability And Study Of Back Contact Molybdenum For Thin Film CIGS2 Solar 

Cells”, Master’s Thesis (University of Central Florida, Orlando) 2003 

[6] A. A. Kadam, “Development Of Large Area Copper Indium Gallium Disulfide (CIGS2) Thin Film Solar 

Cells On Stainless Steel Foil For Space Application”, Master’s Thesis (University of Central Florida, 

Orlando) 2002 

[7] C. Guillein, M.A. Martinez, J. Herrero “Accurate Control Of Thin Film CdS Growth Process By 

Adjusting The Chemical Bath Deposition Parameters”, Thin Solid Films 335 (1998) 37-42 

[8] S. S. Kulkarni, “Development Of Scrubber, Optimization Of Deposition Parameters For Large Area 

CIGS2 Solar Cells”, Master’s Thesis (University of Central Florida, Orlando) 2003

 

NASA/CP—2005-213431 189
Cu(InGa)Se2 SOLAR CELLS ON A FLEXIBLE POLYMER WEB

Robert Birkmire, Erten Eser, Shannon Fields, William Shafarman

Institute of Energy Conversion

University of Delaware, Newark, DE 19716

Introduction

Elemental in-line evaporation on glass substrates has been shown to be a viable process for the large-area

manufacture of CuInSe2-based photovoltaics, with demonstrated module efficiencies as high as 12.7% [1].

However, in certain applications such as space power and mobile terrestrial power sources, lightweight, flexible

modules are preferable. In addition, flexible substrates allow the thin film to be deposited in a roll-to-roll

configuration which may have significant manufacturing advantages. These considerations motivated the use of

high-temperature polymer substrates [2], which resulted in a highest efficiency of 12.8% for a laboratory device

fabricated by a lift-off process [3]. Unlike metal foils, which are also lightweight and flexible, polymer substrates

are (a) electrically insulating, simplifying monolithically-integrated module fabrication, (b) inert in the Cu(InGa)Se2

deposition environment, and (c) are not a source of impurities diffusing into the growing film. However, while

possessing these advantages, polymer substrates have a potential disadvantage in their limited maximum

operating temperature and their effect on adhesion between the Cu(InGa)Se2 film and the Mo back contact. It

should also be mentioned that Na incorporation into the Cu(InGa)Se2 film, which has been shown to increase

efficiency [4], would be problematic for both metal foils and polymer substrates. The Institute of Energy

Conversion (IEC) has modified its in-line evaporation system [5] from deposition onto glass substrates to roll-toroll deposition onto polyimide (PI) film. The addition of this capability has allowed the investigation of key issues

in the deposition of large-area Cu(InGa)Se2 films on flexible substrates. In particular, this paper gives a brief

description of the in-line system and the theoretical model simulating the film growth process and presents

experimental results addressing Cu(InGa)Se2 adhesion to the Mo back contact, large-area uniformity, and device

performance.

Experimental

Cu(InGa)Se2 films were deposited in a multi-source in-line evaporation system onto a 6-inch wide

polyimide/Mo web provided by Global Solar Energy. A detailed

description of the system shown in Figure 1 is given in Reference

5. The system consisted of Cu, Ga, and In sources each having

two nozzles with Se delivered globally via a sparger. The

substrate web was heated from the backside by two sequential

platens. The deposition zone was limited to a 5"x15" area defined

by an aperture just below the web. Cu and In evaporation were

controlled by in-situ atomic absorption spectroscopy (AAS) using

the ATOMICASTM system by Sycon Instruments, while Ga

evaporation was controlled by temperature. The internal layout of

the system is shown schematically in Figure 2. This configuration

where sources are placed sequentially gives, nevertheless,

1. Fig. 1. View of the roll-to-roll deposition system.

NASA/CP—2005-213431 190
uniform Cu distribution through the thickness of the film due to the high diffusivity of Cu atoms. The distribution of

Ga relative to In, however, will change as the film grows. This results in a band gap gradient in the film that can

be engineered, giving another level of control in optimizing the performance of the devices based on these films.

 

Fig. 2. Schematic layout of the web substrate, sources and AAS beams.

Two growth sequences were used to deposit Cu(InGa)Se2 films: 1) The web was moved through a Cu-Ga-In

source sequence at a web speed of 0.75”/min. at a web temperature between 400 to 450°C. This yielded a 1.8

μm Cu(InGa)Se2 film in 20 min. (0.09 microns/min). With this source sequence, the Cu(InGa)Se2 film initially

grows Cu-rich, followed by a Cu-deficient growth stage [5]. 2) The web was initially coated with a Ga-In-Se film

deposited by moving the web through an In-Ga source sequence at a substrate temperature of approximately

350°C. Then, the Cu(InGa)Se2 film was deposited as in sequence 1. Typical thickness of the final film was

between 1.8 to 2.5 microns.

The composition of the Cu(InGa)Se2 film was determined by energy dispersive spectroscopy (EDS) at 20 kV

accelerating voltage. The structure of the Ga-In-Se precursor and Cu(InGa)Se2 films were characterized using

θ/2θ X-ray diffraction (XRD), or glancing incidence XRD, (GIXRD) at an incident angle of 1o. In some cases,

composition depth profiles were obtained by auger electron spectroscopy (AES).

Devices were 0.56 cm2

 in area, and were fabricated using a PI/Mo/Cu(InGa)Se2/CdS/ZnO/ITO/Ni-Al device

structure. After initial J-V characterization under 100mW/cm2

 AM1.5 illumination, devices were sometimes heat

treated in ambient air at 200°C and re-tested. Heat treatment times ranged from 2 to 150 minutes.

Results

Film Growth Model

The film growth model developed considers only the transport of metal atoms from the sources on to the

substrate and assumes unity sticking coefficients and instantaneous chemical reaction on the substrate surface

with Se atoms. The problem is then defined, as in Figure 3, by the effusion rate

of the metal atoms from the source nozzles and by the arrival flux of these

atoms on a unit area of the moving web. Kinetic theory of gases gives for the

evaporative effusion rate from the surface of the melt of area Asurf,

F = M

2πRT

⎡ 

⎣ ⎢ 

⎤ 

⎦ ⎥ 

1

2

psat ( ) ( ) T − p1 Asurf Eq. 1

The theory of gaseous flow through orifices, in the free molecular regime gives,

F = 2πΓ2K M

2πRT

⎡ 

⎣ ⎢ 

⎤ 

⎦ ⎥ 

1

2

( ) p1 − p2 Eq. 2

for the two nozzle sources used.

Fig. 3. Geometry of the effusion

flux onto a substrate.

f θ

θ

F

2Γ

p2

p1

NASA/CP—2005-213431 191
These effusion rates are set equal due to the conservation of mass solving for p1, the internal pressure in the

sources. Assuming that flux intensity profile is of the form acosn

θ, one finds by setting the integrated flux over a

hemispherical surface to be equal to the total effusion that

f = ( ) n +1 F

2πr 2 cosn+1 θ Eq. 3

In the present case, the exponent n was determined experimentally by measuring the thickness of Cu films

deposited on glass slides located on a hemispherical ring placed right over one of the nozzles. Figure 4 below

gives the normalized Cu thickness for a number of effusion rates spanning the range from 0.15 g/hr to 23.2 g/hr

from the nozzle. As can be seen, n = 3 gives an acceptable fit to data.

0.0

0.2

0.4

0.6

0.8

1.0

-90 -60 -30 0 30 60 90

Azimuthal Angle (°)

Normalized Thickness

Data from

0.15 g/hr/nzl

to

23.2 g/hr/nzl Cos3

θ

 

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0 0.2 0.4 0.6 0.8 1

Normalized thickness

Ga/Ga+In

AES depth profiling

Model calculation

Fig. 4. Flux intensity profile from normalized thickness Fig. 5. Ga/Ga+In ratio from Auger depth profiling data

 measurements and cos3

θ distribution. and from the film growth model calculation.

The Ga gradient in a typical Cu(InGa)Se2 film has been measured by AES and is shown in Figure 5, along

with the profile calculated from the model. Normalized thickness 0 indicates the Mo/Cu(InGa)Se2 interface, and 1

indicates the film surface. Although there is some difference between the model and the experimental data at the

Mo interface, this difference becomes negligible in the top 60% of the film. Figure 5 also indicates that the Ga

gradient in these films is almost linear. For this particular Cu(InGa)Se2 film Ga ratio changes from 55% at the Mo

interface to 18% at the surface corresponding to a band gap variation from 1.33 eV to 1.12 eV [6].

Adhesion and Crystallographic Structure

Cu(InGa)Se2 films deposited on glass substrates by sequence 1 and characteristics of devices fabricated on

them were previously reported [5]. This single-layer deposition process produced well-adherent films with good

device efficiencies on Mo-coated glass. Films deposited

directly on the polyimide/Mo web using sequence 1, however,

delaminated during subsequent handling especially during

CdS film growth by the chemical cath deposition method. This

situation was resolved by depositing a Ga-In-Se precursor

layer prior to Cu(InGa)Se2 deposition (i.e. sequence 2). These

films showed no delamination when subjected to a 40 oz/inch

tape test as well as during cell fabrication processes.

XRD measurements show that the Cu(InGa)Se2 films

deposited by sequence 1 had (112) preferred orientation at

~400 °C web temperature and tend to be more randomly

oriented at higher web temperatures. The Cu(InGa)Se2

1. Fig. 6. GIXRD spectrum of the precursor layer.

70120

0

25

50

75

100

20 30 40 50 60 70 80

2θ (°)

Norm. Intensity

Ga2Se3

(131)

Mo

(110)

Ga2Se3

(312/116)

Mo

(211)

NASA/CP—2005-213431 192
films deposited on the Ga-In-Se layer were randomly oriented independent of web temperature. The structure of

the precursor layer was determined by the GIXRD technique and its composition by AES depth profiling. The

GIXRD spectrum of Figure 6 shows that the structure is consistent with that of β-Ga2Se3 (JCPDS 44-1012) with

the peaks shifted to a higher d-spacing due to the presence of indium atoms at substitutional sites in the lattice.

AES indicates an Ga/(Ga+In) ratio of 0.8 identifying the precursor layer to be β-(Ga1-xInx)2Se3.

Uniformity and Device Performance

The cross web thickness uniformity was evaluated by estimating film thickness from SEM cross-sections and

was found to have better than ±3% standard deviation, well within the accuracy of the measurement.

Table I. Film composition over 5 feet x 6" of web. The compositional uniformity was evaluated by an

18-point EDS characterization of a single run of 5 feet

in length and the results are summarized in Table I.

The standard deviation of about 1% for the atomic

concentration and ratios are within the accuracy of

the EDS measurement. Composition controllability

over 16 consecutive runs had a standard deviation of

0.062 for Cu/(Ga+In) and 0.025 for Ga/(Ga+In). The

Cu(InGa)Se2 composition and film thickness data

indicate acceptable uniformity in film properties over

the web surface for a given deposition.

Device uniformity for Cu(InGa)Se2 films in a

single run over a 5 foot length of web are shown in

Table II for 18 devices fabricated along the centerline

of the web. The average Voc was 0.529 V with a

standard deviation of only 0.005 V and average

efficiency was 10%. For this run the Cu/(Ga+In) ratio

was 0.82 and the Ga/(Ga+In) 0.40. Cu(InGa)Se2

film thickness was measured to be 1.9 μm. A 2-D

Table II. Device uniformity over the centerline of the web. device efficiency survey of the web is given in Table

III where each entry is the average of 4 devices over

1"x1" sample. Device efficiencies averaged

10.0%±0.4% along the centerline of the web, and

8.8%±1.0% and 8.4%±0.9% along either edge. The

data show a clear pattern of higher efficiencies along

the centerline, and also a trend of improving

performance and cross-web uniformity down the web.

Figure 7 shows the efficiency of the best device

from each run obtained for 21 different depositions of

5 to 6 feet. The run-to-run composition variation

observed in the figure suggests that the effusion

control technique may require some refinement.

However, it should be noted that the sensitivity of

device efficiencies, especially after heat treatment, on

the Cu/(Ga+In) ratio is quite low, which mitigates

deficiencies in the effusion control scheme as

currently implemented. In fact the data show that the

process is able to produce high-quality material over

a wide range of Cu/(Ga+In) ratios from 0.67 to 0.96.

The best efficiency was obtained for Cu/(Ga+In) = 0.70 and Ga/(Ga+In) = 0.26. After a 150-minute heat treatment

the efficiency was 11.5%, reaching 12.1% with a 1200 Å MgF2 anti-reflection coating (see Fig. 8).

The device parameters currently limiting performance are open-circuit voltage and fill factor is probably due to

the fact that substrate temperature is limited to 400°C to 450°C because of the polyimide substrate. However,

Cu In Ga Se Cu/III Ga/III

X Y (at%) (at%) (at%) (at%)

0.0 10 23.3 20.0 7.7 49.0 0.84 0.28

-1.5 25 24.1 19.9 7.8 48.3 0.87 0.28

1.5 25 23.8 19.7 7.7 48.8 0.87 0.28

0.0 30 23.9 19.9 7.2 49.0 0.88 0.26

-1.5 35 23.9 19.7 7.6 48.8 0.88 0.28

0.0 35 23.7 19.8 7.6 48.9 0.86 0.28

1.5 35 24.1 19.6 7.5 48.9 0.89 0.28

0.0 40 24.3 19.6 7.3 48.9 0.90 0.27

0.0 45 23.6 19.6 7.2 49.6 0.88 0.27

-1.5 45 23.8 19.5 7.3 49.3 0.89 0.27

1.5 45 24.2 19.5 7.3 49.1 0.91 0.27

0.0 50 23.9 19.7 7.5 49.0 0.88 0.28

0.0 55 24.2 20.0 7.5 48.3 0.88 0.27

0.0 60 23.6 19.6 8.1 48.8 0.85 0.29

1.5 65 24.5 19.5 8.0 48.0 0.89 0.29

-1.5 65 24.4 19.5 7.6 48.5 0.90 0.28

0.0 65 24.4 19.8 7.5 48.3 0.89 0.27

0.0 70 24.4 19.7 7.6 48.2 0.89 0.28

Average 24.0 19.7 7.5 48.7 0.88 0.28

Stdev (%) 1.4 0.9 3.3 0.9 1.9 2.5

Position (")

Position Cell Voc Jsc FF Eff

(") # (V) (mA/cm2) (%) (%)

20 3 0.517 29.6 56.7 8.7

4 0.533 30.2 63.0 10.2

30 1 0.527 30.6 60.1 9.7

2 0.521 31.0 60.1 9.7

3 0.534 30.2 62.4 10.1

4 0.532 30.3 60.8 9.8

40 1 0.529 30.7 62.5 10.1

2 0.528 30.6 62.6 10.1

3 0.532 30.9 62.6 10.3

4 0.528 30.8 62.7 10.2

50 1 0.528 31.0 60.5 9.9

2 0.533 30.9 62.2 10.2

3 0.529 30.4 60.1 9.7

4 0.533 30.8 63.0 10.4

60 1 0.529 30.4 61.1 9.8

2 0.526 31.0 61.6 10.1

3 0.535 30.4 61.6 10.0

4 0.531 31.0 62.4 10.3

Best 0.533 30.8 63.0 10.4

Avg 0.529 30.6 61.4 10.0

Stdev (%) 0.86 1.2 2.6 3.9

NASA/CP—2005-213431 193
incorporation of Na into the Cu(InGa)Se2 film, using a separate source, would contribute to improving the

performance.

Table III. 2-D efficiency survey of 5' of web.

Location 20” 30” 40” 50” 60” Cumulative 

-1.5” 9.0 ± 0.4 8.2 ± 1.6 9.1 ± 0.3 8.8 ± 1.0 

0 9.4 ± 1.1* 9.8 ± 0.2 10.2 ± 0.1 10.0 ± 0.3 10.0 ± 0.2 10.0 ± 0.4 

+1.5” 8.2 ± 0.4 8.2 ± 1.5 9.0 ± 0.2 8.4 ± 0.9 

Cumulative 8.7 ± 2.1 8.8 ± 1.5 9.4 ± 0.5 9.2 ± 1.0 

-40

-30

-20

-10

0

10

20

30

40

-0.4 -0.2 0.0 0.2 0.4 0.6

Volts (V)

Current Density (mA/cm2)

Light

Dark

70146.42-004

Fig. 7. Best device efficiency as a function of Fig. 8. J-V characteristics of the best device with a

 Cu/(Ga+In) ratio over 21 depositions. MgF2 anti-reflection coating. Jsc = 35.4 mA/cm2,

Voc = 0.512 V, FF = 66.6 %, Eff = 12.1 %.

Conclusion

A pilot process for depositing uniform, adherent Cu(InGa)Se2 films onto a continuous polyimide/Mo web has

been demonstrated. Average device efficiencies are better than 9%, with a highest efficiency of 12.1%. Device

uniformity in the range of 10% has been demonstrated. The process also demonstrates a high tolerance to

Cu/(Ga+In) variation. In summary, this study lends support to the feasibility of CuInSe2-based photovoltaics on

polyimide, and has identified both fundamental and engineering issues to further their realization.

References

1. Powalla, M. and B. Dimmler, Thin Solid Films 361-361, 540 (2000).

2. Basol, B.M., V.K. Kapur, C.R. Leidholm, A. Halani, and K. Gledhill, Solar Energy Materials and Solar Cells 43,

93 (1996).

3. Tiwari, A.N., M. Krejci, F.-J. Haug, and H. Zogg, Progress in Photovoltaics: Research and Application 7,

(1999), pp. 393-397.

4. Shafarman, W., Stolt, L., Chapter 13 in Handbook of Photovoltaic Science and Engineering, (A. Luque and S.

Hegedus, eds.), Wiley and Sons, Chichester UK, 2003.

5. Hanket, G.M., P.D. Paulson, U. Singh, S.T. Junker, R.W. Birkmire, F.J. Doyle III, E. Eser, and W.N.

Shafarman, 28th IEEE Photovoltaic Specialists Conf., 2000.

6. Paulson, P.D., R.W. Birkmire, and W.N. Shafarman, J. Appl. Phys., 94, 879 (2003)

NASA/CP—2005-213431 194
WIDE-BANDGAP CIAS THIN-FILM PHOTOVOLTAICS WITH TRANSPARENT BACK CONTACTS FOR NEXT 

GENERATION SINGLE AND MULTI-JUNCTION DEVICES 

Lawrence M. Woods, Ajay Kalla, Damian Gonzalez and Rosine Ribelin 

ITN Energy Systems Inc., Littleton, CO, 80127 

INTRODUCTION 

Future spacecraft and high-altitude airship (HAA) technologies will require high array specific power (W/kg), which 

can be met using thin-film photovoltaics (PV) on lightweight and flexible substrates.(1) It has been calculated that 

the thin-film array technology, including the array support structure, begins to exceed the specific power of 

crystalline multi-junction arrays when the thin-film device efficiencies begin to exceed 12%.(2) Thin-film PV 

devices have other advantages in that they are more easily integrated into HAA’s, and are projected to be much 

less costly than their crystalline PV counterparts. Furthermore, it is likely that only thin-film array technology will 

be able to meet device specific power requirements exceeding 1 kW/kg (photovoltaic and integrated 

substrate/blanket mass only). Of the various thin-film technologies, single junction and radiation resistant CuInSe2

(CIS) and associated alloys with gallium, aluminum and sulfur have achieved the highest levels of thin-film device 

performance, with the best efficiency, reaching 19.2% under AM1.5 illumination conditions and on thick glass 

substrates.(3) Thus, it is anticipated that single- and tandem-junction devices with flexible substrates and based 

on CIS and related alloys could achieve the highest levels of thin-film space and HAA solar array performance. 

ITN Energy Systems, Inc. (ITN) intends to meet the requirements for future high-altitude airship and spacecraft 

power by developing highly efficient single-junction and two-terminal monolithic tandem (multi-junction) PV cells 

using low-cost thin-films on lightweight flexible substrates. ITN will achieve this goal using thin-films of highefficiency and radiation resistant CIS partnered 

with bandgap-tunable CIS-alloys with Ga (CIGS) 

and/or Al (CIAS) on novel lightweight and flexible 

substrates. ITN is currently developing a widebandgap CIAS top cell on infra-red transparent 

back contacts and polyimide substrates to 

increase the single-junction performance over 

existing thin-film CIGS on thin metal foil. An 

illustration of this configuration is given in Figure 1. 

As a single-junction material, a device using a 

wide-bandgap CIAS solar absorber offers several 

performance advantages over standard lowbandgap CIGS devices: it is better tuned to the 

optimum bandgap for the AM0 solar spectrum; it is 

better suited for high voltage applications; it 

undergoes lower resistive losses for cells and 

modules; it transmits more unused infra-red light 

for lower temperature operation; and it produces 

higher power during high-temperature applications 

due to lower temperature coefficients. A widebandgap single junction with transparent back 

contacts also enables potential backside power 

generation and is applicable as the top cell and 

CuInAlSe2 (p-type absorber)

CdS (n-type window) ZnO/ITO (transparent top contact)

Polyimide (IR transparent substrate)

TCO (IR transparent back contact)

Thin IR

transparent 

contact

layer

IR Light

Incoming

White Light

Figure 1 – Cross-sectional (not to scale) illustration of widebandgap CIAS device with IR transparent back 

contacts and substrate. 

NASA/CP—2005-213431 195
transparent interconnect components in a monolithic tandem device. This paper discusses the current 

performance, testing, and problems of the wide-bandgap single junction CIAS device with transparent back 

contacts. 

 

WIDE-BANDGAP CIAS ALLOY THIN-FILM DEVICES 

The electronic quality of the CuIn1-xGaxSe2 (CIGS) material has been shown to degrade for x > 0.3, thus the 

achievement of high-efficiency wide-bandgap CIGS devices is not as easy as simply increasing the Ga alloying in 

the CIGS material. The best CGS devices (x=1 or bandgap equal to 1.67 eV) were recently fabricated and tested 

at 9.53% under AM1.5 illumination (Reference: NREL report, “Polycrystalline Thin Film High Performance: NREL 

In-house Activities,” March/April 2003 bimonthly); a good result but far from the 19.2% efficiency of the lowbandgap material. One approach to potentially overcome the bandgap limitations of CIGS devices is to alloy the 

CIS with aluminum instead of gallium, thus forming thin-film CuIn1-xAlx

Se2 (CIAS). This approach for thin-film 

photovoltaic devices was initiated by ITN in conjunction with the Institute of Energy Conversion (IEC) nearly four 

years ago, and is currently under continued development at ITN for large-area (production-like) CIAS depositions. 

The potential advantage of CIAS is that, for a given wide bandgap, much less Al alloying is needed compared to 

alloying with Ga,(4) and thus potentially avoiding defect formation associated with higher alloy (Ga or Al) content 

materials.(5,6) Indeed, recent information about CIAS deep defect density has indicated that Al alloying has more 

of an effect on the band tail defect distribution, rather than deep defects (Reference: Heath et al., Spring MRS 

2003, to be published). R&D CIAS devices with efficiencies of about 11% have been fabricated with bandgap • 

1.5 eV (x = 0.5). Furthermore, excellent devices (eff. ≈ 17% under AM1.5) have recently been achieved using 

small amounts of Al alloying,(7) and with low substrate temperatures (450 oC) during CIAS co-evaporated 

deposition.(8) This good performance capability for this material system at the lower substrate temperatures is 

important for the temperature-limited polyimide substrates and for reducing the stress to the bottom cells in 

monolithic tandem devices. For CIAS with more top-cell-optimal bandgaps of around 1.65 eV, the best devices to 

date are 7.8%, but without a performance enhancing anti-reflective top coat, and is thus similar to the CGS 

performance mentioned above.(8) All these references to CIAS device performance were measured on devices 

with CIAS deposited without any bandgap tailoring vs. depth, which is typically associated with the highest 

performance devices. More latitude with bandgap tailoring is available for the wide-bandgap materials using the 

CIAS alloy system, given its endpoint bandgap of 2.7 eV (CAS), compared to 1.68 eV for CGS in the CIGS alloy 

system. ITN’s CIAS device performance, from the large area and moving substrate deposition system (production 

like), show similar performance trends with Al content as stated above, and are currently achieving about 2/3rds the 

efficiency of the best laboratory (small area, stationary substrate) devices. Some light IV curves from a series of 

CIAS devices with variable Al-content made using the large-area, moving substrate system at ITN are shown in 

the left side of Figure 2. 

30424-1

-25

-20

-15

-10

-5

0

5

10

15

20

25

-0.2 0 0.2 0.4 0.6

Voltage (V)

Current Density (mA/cm^2)

Cell C9

Cell B8

Cell C7

Cell D6

Cell C5

Sample Number W@V=0 Al ratio Cu ratio

30424-1-C5 1.55 0.52 0.56

30424-1-D6 1.45 0.51 0.57

30424-1-C7 1.21 0.48 0.67

30424-1-B8 1.02 0.44 0.65

30424-1-C9 0.74 0.40 0.68

0.8

0.85

0.9

0.95

1

1.05

-0.25 -0.15 -0.05 0.05 0.15 0.25

Voltage (V)

(Jillum - Jdark)/JL

Cell C9

Cell B8

Cell C7

Cell D6

Cell C5

Figure 2 –Light IV curves (left) and voltage dependence of the normalized current difference (right) of 

small-area (1 cm2

) wide-bandgap CIAS devices from the large-area, moving substrate deposition 

system. Device composition and zero bias depletion width (W@V=0) are given in the inset 

tables. 

NASA/CP—2005-213431 196
Table 1 - AM1.5 Light IV parameters for small-area (1 cm2

) wide-bandgap CIAS devices shown in Figure 2. 

Device Eff. (%) Voc (V) Jsc (mA/cm2

) Fill Factor Rs (ohms) Rsh (ohm-cm) Diode Factor

30424-1-C5 4.66 0.504 19.71 0.47 4.01 227.00 4.36

30424-1-D6 5.65 0.511 20.45 0.54 3.22 255.00 3.68

30424-1-C7 6.41 0.526 21.81 0.56 3.34 166.00 3.07

30424-1-B8 6.61 0.532 22.15 0.56 3.33 107.00 3.14

30424-1-C9 6.47 0.535 21.51 0.56 3.55 174.00 3.05

The IV curves indicate an apparent high series resistance problem that becomes worse when the Al-ratio exceeds 

0.5. Closer examination shows that the series resistance (Rs) as determined using the intercept of a plot of dV/dJ 

vs J-1, given in Table 1, does not change much with Al content. Furthermore, a plot with the voltage dependence 

of the normalized current difference (right side of Figure 2) shows significant voltage dependence in all devices, 

especially in the device with the highest Al-content. Thus, the IV curve features are partially a result of voltagedependent current collection. In addition, there is a decrease in the open-circuit voltage (Voc) for these higher Al 

content devices, instead of steadily increasing with increasing Al content. This is accompanied by an increased 

diode factor (A, in Table 1) and decreasing FF with increasing Al-content. Using the standard solar cell model, 

the series of measured light IV results can only be fitted by increasing the diode saturation current and diode 

factor with increasing Al-content, and thus indicates changes in the CIAS electronic properties with Al-content. 

Indeed, the zero bias depletion widths from Capacitance-Voltage (CV) testing, shown in the figure inset (in units 

of microns), indicate an increase in the junction depletion width with increasing Al-content. This indicates a 

decrease in the net doping density near the junction with increasing Al-content. Additional data on the zero bias 

depletion depth versus CIAS composition are shown in Figure 3, and confirm the correlation with Al content. In 

this figure the data is color-coded for different Cu-ratio ranges, so that the effect of Cu-ratio on the depletion width 

can be separated out. For this analysis, the depletion widths were calculated using a composition-dependent 

dielectric constant, εr

(x), which was estimated as linearly decreasing with Al-content between εr(0)=11.5 and 

ερ(1)=8.8, representing the respective CIS and CAS endpoint ternaries. This adjustment works to counteract the 

apparent trend, thus without it, the depletion width dependence on Al-content would have shown a higher 

correlation. Similarly, a reduction in depletion width from Cu-ratios around 0.63 is also apparent as shown in 

Figure 3, but is not as strongly correlated as the Al-ratio. Here also the data is color coded for different Al-ratio 

ranges, so that the effect of Al-ratio on the depletion width can be separated out. 

 

Depletion Width vs. Al Ratio

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

2.00

0.30 0.40 0.50 0.60 0.70 0.80

Al Ratio

Depletion Width

Cu Ratio (0.45-0.52)

Cu Ratio (0.55-0.60)

Cu Ratio (0.61-0.68)

Cu Ratio (0.69-0.74)

Depletion Width vs. Cu Ratio

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

2.00

0.40 0.45 0.50 0.55 0.60 0.65 0.70 0.75 0.80

Cu Ratio

Depletion Width

Al Ratio (0.33-0.36)

Al Ratio (0.40-0.44)

Al Ratio (0.48-0.53)

Al Ratio (0.54-0.62)

Figure 3 - Zero Bias depletion width versus Al-ratio (left) and Cu-ratio (right) for several devices. Data is 

adjusted for linearly vary dielectric constant with Al-content and color coded to bracket certain 

Cu-ratio ranges (left) and Al-ratio ranges (right). 

NASA/CP—2005-213431 197
Owing to the lower net doping level in the junction region, the high Al and low Cu content (highly compensated) 

CIAS devices may have a more severe surface inversion (n-type) layer, a feature now commonly associated with 

low bandgap CIGS devices. This inversion layer, with an electron barrier at the CdS/inversion layer interface, 

could lead to the observed voltage dependent current collection and the kink feature sometimes observed in IV 

curves. A barrier in the conduction band at the CdS/inversion layer interface is possible with increasing 

conduction band maximum in the CIAS and an n-type inversion layer. This possibility is supported by the high 

sensitivity of the voltage dependent current collection to air annealing (surface sensitivity), which may reduce the 

inversion layer thickness and doping in high Al content CIAS devices. An IV curve example of this sensitivity is 

shown in Figure 4.

30411-2-C1

-25

-20

-15

-10

-5

0

5

10

15

-0.2 0 0.2 0.4 0.6

Voltage (V)

Current Density (mA/cm^2)

Light IV

Air Anneal Light IV

Sample Number W@V=0 Al ratio Cu ratio

30411-2-C2 1.46 0.41 0.50

30411-2-C2 Anneal 0.79 0.41 0.50

Figure 4 – Example IV’s showing the effect of air anneal on high bandgap CIAS devices. 

As expected, the short-circuit current (Jsc) decreases with increasing Al-content (bandgap). The quantum 

efficiency (QE) measurements shown in Figure 5 confirm the increasing bandgap with Al content, as indicated by 

the decreasing long wavelength cut-off with increasing Al content (left side of Figure 5). The QE measurements 

also indicate wavelength dependent collection with reduced collection at longer wavelengths and significant subbandgap collection. These characteristics worsen with increasing bandgap (Al-content), as does the maximum 

QE, and is easier to see when the QE curves are aligned near their long-wavelength inflection points as shown in 

the right side of Figure 5.

Sample 30424-1 0

10

20

30

40

50

60

70

80

90

400 500 600 700 800 900 1000 1100 1200

Wavelength (nm)

QE (%)

Cell C9

Cell B8

Cell C7

Cell D6

Cell C5

Sample Number W@V=0 Al ratio Cu ratio

30424-1-C5 1.55 0.52 0.56

30424-1-D6 1.45 0.51 0.57

30424-1-C7 1.21 0.48 0.67

30424-1-B8 1.02 0.44 0.65

30424-1-C9 0.74 0.40 0.68

Sample 30424-1 

0

10

20

30

40

50

60

70

80

90

A.U.

QE (%)

Cell C9

Cell B8

Cell C7

Cell D6

Cell C5

Sample Number W@V=0 Al ratio Cu ratio

30424-1-C5 1.55 0.52 0.56

30424-1-D6 1.45 0.51 0.57

30424-1-C7 1.21 0.48 0.67

30424-1-B8 1.02 0.44 0.65

30424-1-C9 0.74 0.40 0.68

Figure 5 - QE’s (right) of small-area (1 cm2

) wide-bandgap CIAS devices from the large-area, moving 

substrate deposition system. Device composition and zero bias depletion width (W@V=0) are 

given in the inset tables. QE’s on left are the same devices but adjusted to line-up the long 

wavelength inflection point. 

NASA/CP—2005-213431 198
Devices with a depletion width extending almost completely through the CIAS thickness (higher Al content) do not 

give the expected better long wavelength collection efficiency, and indicate an increase in junction and bulk 

recombination. 

TRANSPARENT BACK CONTACTS 

Reverse bias capacitance-voltage profiling (not shown) of CdS/CIAS devices with standard opaque Mo back 

contacts, reveals that the net bulk doping density is decreasing with increasing Al-content and correlates with the 

zero bias depletion widths discussed above. The lowest values are measured to be in the low 10-15 cm-3 range. 

This could affect the back contact as the native bulk doping level may not be sufficient for a good low resistance 

contact or to get appreciable majority carrier thermally assisted tunneling and thermionic emission over the back 

contact Schottky barrier.(9) The majority carrier current is likely to be further impeded by an increasing Schottky 

barrier energy with increasing Al-content, due to a decreasing CIAS valance band maximum energy.(10) This 

could be the reason for the high series resistance measured in the devices above. High-efficiency CIGS devices 

have order of magnitude higher doping levels than those measured on the CIAS devices herein, and high 

bandgap CGS devices do not show back contact related problems. This is presumably due to increasing net 

acceptor doping with increasing Ga content, and a valance band maximum energy that does not change much 

with increasing Ga content. These results lead us to believe that the high series resistance observed in the high 

Al-content CIAS devices is due in part to the back contact. Thus, the low doping densities in high Al-content CIAS 

could further complicate the ability to make low resistance transparent back contacts. 

Several transparent conducting oxides (TCO’s) have been tested as IR transparent replacements to the standard 

opaque Mo back contacts on high- and low-bandgap CIGS devices. Indium Tin Oxide (ITO) was demonstrated to 

make good devices with low-bandgap CIGS (low Ga, Eg = 1.1 eV), while additionally showing good durability to 

the top cell deposition. However, high-bandgap CIGS or CIAS devices (Eg > 1.35 eV) do not make good devices 

directly on ITO. Other TCO’s such as SnO2 and ZnO did not fair well either, even with the low-bandgap CIGS, 

indicating that the type of TCO is important to the transparent back contact. Theoretically, this could be due to the 

decreasing valance band maximum with increasing bandgap, as previously discussed, and a lack of sufficient 

interface states or mid-gap states in the TCO for tunneling or thermally emitted holes followed by recombination 

processes. The energy band diagram of the back contact “Schottky” barrier at the p-type CIAS and degenerate ntype TCO interface, and the majority carrier tunneling and recombination process is depicted in Figure 6 below. 

These results suggest that an interface layer is necessary for wide-bandgap devices using TCO back contacts. 

++

++

P-type 

CIAS

N-type

TCO

+

EC

EV

EF

TCO Intra-gap

Defect States } Edefect EC

EV

EF

CIAS/TCO Band Diagram

Tunneling

Recombination

Legend

+ Hole

Electron

Figure 6 - Energy band diagram of the back contact “Schottky” barrier at the p-type CIAS and degenerate 

n-type TCO interface, and showing the majority carrier tunneling and recombination process. 

NASA/CP—2005-213431 199
Since the earlier CIAS and CIGS device testing with TCO back contacts, ITN has developed an interface layer to 

the TCO’s. The normalized transmission data for this interconnect is shown in Figure 7 below, and was tested to 

have an average transmission of about 71.5%, weighted to the AM0 solar irradiance in the range of 750 nm to 

1300 nm. 

0.40

0.50

0.60

0.70

0.80

0.90

1.00

500 700 900 1100 1300 1500

Wavelength (nm)

T/[(1-R)*Tsub)]

Thin ITO

Thin ITO + Interface Layer

Detector Change

Figure 7 – Normalized Transmission (substrate divided out) of a thin ITO layer, and a thin ITO layer plus a 

top interface layer. 

This level of transmission enables approximately 33% of the useable AM0 solar power to be transmitted to the 

bottom cell in a tandem configuration when assuming a top cell bandgap of 1.65 eV or 750 nm, or to be 

transmitted out of the device for lower temperature operation from a wide-bandgap single junction device. In 

addition, this back contact would enable over 70% transmission of the visible spectrum to pass through, in the 

event of bifacial illumination. 

The same semi-transparent back contact structure was used in the fabrication of CIAS devices (bandgap ≈ 1.35 

eV) and efficiencies near 11% were obtained. These efficiencies were comparable to the performance of the 

controls using the standard opaque molybdenum back contacts. The results are shown in Table 2 below. 

However, the device results have not yet tested for compatibility with wider-bandgap CIAS (Eg > 1.5 eV). 

Table 2 – CIAS device with transparent back contact and Mo controls (Al/[Al+In] = 0.35, Cu/[Al+In] = 0.69, 

Tsub = 450°C) 

SUMMARY 

Significant challenges exist in the development of high-efficiency wide-bandgap devices, but this development is 

still in its infancy and some possibilities for improvement exist. The wide-bandgap CIAS device performance 

 Glass Substrate/ Voc Jsc FF η 

 Back Contact (Volts) (mA/cm2

) (%) (%) 

 SL/Mo 0.58 29.4 63.5 10.9 

 7059/Mo 0.56 28.8 65.7 10.7 

7059/Mo/TCO/Thin TIL* 0.59 27.8 64.9 10.7

* TIL - transparent interface layer

NASA/CP—2005-213431 200
without any bandgap tailoring to date has not yet proven to be any better than the CIGS performance at the same 

bandgap.(8) This equivalence supports some recent theories about deep defects becoming more effective by 

moving toward mid-gap with increasing bandgap.(11) However, there are enough differences between the two 

wide-bandgap CIS alloys to suggest the possibility of different mechanisms limiting their respective performance. 

These include differences in the intra-gap defect distribution, net native doping, and valance and conduction band 

maximum energies, all of which can have implications for device performance. The test results presented herein 

suggest that improvements to the CIAS near surface properties could improve the high Al content devices from 

the large-area, moving substrate deposition system. This could involve modified CIAS deposition near the 

surface or some post deposition surface treatments. The test results also suggest that a back contact interface 

layer may be needed to help improve the back contact. Other device test results with CuGaSe2 (Eg = 1.67 eV) 

solar absorbers show dominant interface recombination, indicating that better matched alternative heterojunction 

partners or surface treatments may be needed for the wide-bandgap devices.(12) 

Finally, the potential for obtaining good transparent back contacts to wide-bandgap devices is promising, with 

demonstrated IR transparency greater than 60% on some back contact structures and device efficiencies that are 

comparable to standard Mo back contacts, but there is still much room for improvement. 

ACKNOWLEDGEMENTS 

ITN would like to acknowledge the Institute of Energy Conversion for collaborative work performed on CIAS on 

TCO results. The results discussed herein were obtained by work funded by: NIST/Dept. of Commerce, Advanced 

Technical Program, Contract No. 70NANB8H4070; and NASA, Contract No.’s NAS3-02027 and NAS3-02170. 

REFERENCES 

 

1

 S.G. Bailey, A.F. Hepp, and R.P. Raffaelle, “Thin Film Photovoltaics for Space Applications,” Proc. Of the 36th

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Cells, J. Appl. Phys., 91, pg. 10153, (2002). 5

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6

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Jersey, pp. 324-332, (2001). 

7

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Cu(InAl)Se2 Thin Films, Appl. Phys. Lett., 81,(7), pp. 1350-1352, (2002). 8

 W. Shafarman, S. Marsillac, P. Paulson, M. Haimbodi, T. Minemoto, and R. Birkmire, “Material and Device 

Characterization of Thin Film Cu(InAl)Se2 Solar Cells, Proc. of the 29th IEEE PVSC, pp. 519-522, (2002). 9

 A. Fahrenbruch and R. Bube, “Fundamentals of Solar Cells – Photovoltaic Solar Energy Conversion,” Academic 

Press, London, UK, 1983, pg. 169.

10 S. B. Zhang, S.H. Wei, and A. Zunger, “A Phenomenological Model for Systematization and Prediction of 

Doping Limits in II-VI and I-III-VI2 Compounds,” J. Appl. Phys., 83(6), pp. 3192-3196, (1998). 11 J. Heath, J. Cohen, W. Shafarman, D. Liao, and A. Rockett, “Effect of Ga Content on Defect States in CuIn1-

x

GaxSe2 Photovoltaic Devices,” Appl. Phys. Lett., 80(24), pp.1-3, (2002). 12 J. Malmstrom, J. Wennerberg, M. Bodegard, and L. Stolt, ”Influence of Ga on the Current Transport in 

Cu(In,Ga)Se2 Thin Film Solar Cells,” Proc. of the 17th European Photovoltaic Solar Energy Conference and 

Exhibition, Munich, Germany, (2001).

NASA/CP—2005-213431 201
DEVELOPMENT OF A THIN FILM SOLAR CELL INTERCONNECT FOR THE POWERSPHERE CONCEPT 

Edward J. Simburger, James H. Matsumoto, Thomas W. Giants, Alexander Garcia III, Simon Liu 

The Aerospace Corporation, El Segundo, CA 90245 

Dr. Suraj P. Rawal, Alan R. Perry, Craig H. Marshall, 

Lockheed Martin Corporation, Denver, CO 80125 

John K. Lin, Stephen E. Scarborough 

ILC Dover Incorporated, Dover, DE 19946 

Henry B. Curtis, Thomas W. Kerslake, Todd T. Peterson 

NASA Glenn Research Center, Cleveland, OH 44135

Introduction

Dual junction amorphous silicon (a-Si) solar cells produced on polyimide substrate have been selected as the 

best candidate to produce a lightweight solar array for the PowerSphere program. The PowerSphere concept 

features a space-inflatable, geodetic solar array approximately 0.6 meters in diameter and capable of generating 

about 20W of electrical power. Trade studies of various wiring concepts and connection methods led to an 

interconnect design with a copper contact that wraps around the edge, to the back of the solar cell. Applying 

Plasma Vapor Deposited (PVD) copper film to both sides and the edge of the solar cell produces the wrap around 

contact. This procedure results in a contact pad on the back of the solar cell, which is then laser welded to a flex 

circuit material. The flex circuit is constructed of copper in a custom designed routing pattern, and then 

sandwiched in a Kapton® insulation layer. The flex circuit then serves as the primary power distribution system 

between the solar cells and the spacecraft1,2,3. Flex circuit material is the best candidate for the wiring harness 

because it allows for low force deployment of the solar cells by the inflatable hinges on the PowerSphere. An 

additional frame structure, fabricated and assembled by ILC Dover, will reinforce the wrap around contact-flex 

blanket connection, thus providing a mechanically robust solar cell interconnect for the PowerSphere 

multifunctional program4,5. The PowerSphere team will use the wraparound contact design approach as the 

primary solution for solar cell integration and the flex blanket for power distribution. 

Use of the wrap around contact will yield three key benefits noted below. 

1.) Utilizing the back of the solar cell for PVD copper electrical traces due to the insulation properties of 

polyimide. This multifunctional benefit reduces mass by removing what would otherwise be additional flex 

circuit material. Future advancements in this area could result in removal of the entire flex material from 

Powersphere solar array system creating an integral wiring harness. 

2.) The use of the wrap around contact physically locates both the top and bottom contact on the same side of 

the solar cell. This provides fabrication process simplification by laser welding both contacts of the solar cell 

to the flex circuit at the same time thus reducing costs. 

3.) After the wrap around contact PVD process is complete a protective Tefzel® film is applied over the solar cell 

and contact. The Tefzel® film is applied to the front of the solar cell to increase the thermal emissivity to 0.8, 

which limits the peak temperature of the cell from exceeding 80°C. It also improves the robustness of the 

solar cell and mitigates possible risk of damage from a handling aspect. By utilizing the back of the solar cell 

as the contact, the protective Tefzel® film will not impede on the laser welding process. 

NASA/CP—2005-213431 202
Requirements

Developments of a solar cell interconnect to be used for a Low Earth Orbit (LEO) application was driven by 

several specific requirements. It must be able to withstand the extreme environmental conditions that are present 

including Atomic Oxygen (AO), Ultraviolet light (UV) and rapid thermal cycling. Exposure to these conditions will 

cause degradation of the solar cell interconnect and possibly complete failure of the joint. 

Additional requirements that were derived for the interconnect joint are described below. 

1.) Mechanically robust in order to survive loads incurred during launch. 

2.) Adequate sizing to support electrical loads. 

3.) Ability to be scaled to support different sizes of solar cells. 

4.) Easily repaired if interconnect is damaged. 

Background

Several trade studies were undertaken to investigate the successful fabrication of solar cell interconnect 

architecture. Various methods of joining the a-Si solar cell to a flexible Kapton® blanket including joint 

architecture, materials and fabrication methodology were considered. 

Prior to starting the trade studies a lap joint configuration was selected as the best candidate to verify all of 

the requirements to evaluate the robust solar cell interconnect. Also, methods that both electrically and 

mechanically joined the solar cell to the flex circuit material received the highest priority. 

The architecture trade study parameters included the overall geometry, mechanical and electrical 

configuration of the interconnect joint. Several methods to electrically and mechanically join the materials were 

considered. Conventional methods such as conductive epoxies, solders and conductive tapes were quickly 

eliminated due to lack of confidence in the survivability at the temperature extremes. Additional methods that 

were considered to provide the electrical connection include drilled vias, welded wire or foil and PVD copper film. 

Based on the trade studies, PVD copper film was selected as the leading candidate to provide the electrical 

interface. Pressure Sensitive Adhesive (PSA) was used to mechanically join the solar cell to the flex blanket. 

The PVD copper film was applied over the lap joint to produce the coupon shown in Figure 1. 

Upilex® Substrate

Simulated Flex

PSA Under Contact

Top Contact

PVD Copper Trace

Figure 1- PVD Copper/PSA Coupon 

This concept proved to be challenging due to the high out gassing of the adhesive during the vacuum 

deposition process and relatively large step between the cell and flex circuit. Thermal mismatch between the 

materials was also a concern and this concept was abandoned. 

As a result of the previous coupon preparation an alternate PVD process was considered. This concept 

utilized the PVD copper film applied to the solar cell contact that was laser welded to the flex circuit. The copper 

film was applied to the top, edge and bottom of the solar cell thus routing the generated electrical load to the back 

of the solar cell. This wrap around contact is achievable by sputtering the copper film unlike other deposition 

NASA/CP—2005-213431 203
processes that provide line of site application only. Having both contact located on the back of the solar cell 

provides many benefits that are previously noted. A cross section of wrap around concept is shown in Figure 2. 

Several coupons were produced using the wrap around design. 

Upilex®

Kapton® Flex Laser Welds

Wraparound Contact

Solar Cell

Copper Trace

Top Contact

Laser Scribe

Back Contact

Figure 2- Cross Section of Wrap Around Contact 

Figure 3 shows a magnified view of the initial laser welds that were produced indicating a good laser weld 

between the back contact and the PVD copper film. The picture shows the Kapton® side of the flex circuit 

where the laser makes first contact. 

Upilex® Substrate 

w/PVD Copper 

(Farside)

Flex Blanket

Laser Weld

Figure 3- Laser Weld Coupon 

Once the fabrication process was established, four coupons were produced using representative materials. 

Two of the coupons had an additional ILC Dover mock frame installed to simulate actual conditions. The coupons 

were subjected to rapid thermal cycling at NASA Glenn Research Center to simulate a LEO environment. 

Pretest and posttest resistance measurements between the simulated flex circuit and solar cell were documented 

to evaluate the electrical continuity and integrity of the welded coupons. During the thermal cycle testing all but 

one of the coupons survived with the one failure attributed to material handling. The posttest resistance 

measurements showed that there was little degradation of the laser-welded solar cell interconnect. Figure 4 

shows both types of coupons, with and without the simulated frame. 

NASA/CP—2005-213431 204
Simulated Solar Cell

PVD Copper Film

Flex Blanket

Laser Welds

Simulated Frame

Figure 4- Unframed and Framed Weld Coupons 

A thermograph inspection was also performed to verify adequate electrical sizing of the laser welds when 

subjected to 1 amp of current. The white areas around the laser welds in Figure 5 represent the coolest area and 

confirmed that the weld sizing was done correctly. 

Laser Welds

Flex Blanket

PVD Copper Traces

Simulated Solar Cell

Figure 5-Laser Weld Thermograph 

Design Improvements

Once the wrap around concept proved to be a viable approach for fabricating a robust interconnect, five 

improvements in the weld approach were investigated. The first improvement attempted was to laser weld 

completely through the flex circuit and solar cell. Visual observation of the plasma exiting the weld was used as 

initial verification of a successful weld. This can only be accomplished on the edge of the solar cell where the 

contact pad was isolated from the active solar cell area and eliminate possible shorting of the cell. Testing has 

shown that this change has significantly improved the quality of the weld. 

A second improvement to the design was to decrease the PVD copper thickness on the wrap around contact. 

This was attempted to reduce the residual stresses in the copper film, which caused the substrate to curl and can 

potentially cause possible delaminations. However, sufficient copper thickness must be present to provide 

adequate material for the weld. The PVD copper thickness was optimized to meet all of the previously noted 

requirements. An analysis of the electrical sizing was also accomplished to verify that the reduction in film 

thickness was still adequate to carry solar cell power. 

The third improvement was to apply a titanium thin film binder layer prior to the PVD copper film. This helped 

to reduce the delaminations of the copper from the solar cell substrate during the welding process. The 

delaminations occurred due to localized shrinkage of the weld material. The future binder material will use 

chromium at the base layer that has previously demonstrated superior properties to that of titanium. 

NASA/CP—2005-213431 205
The laser welding process provides both the mechanical and electrical joint, however the mechanical strength 

of the weld is weak suggesting the need for a adequate structural support such as Kapton® tape. Tape was used 

to provide the clamping force during the laser welding process. The tape also helped to reinforce the solar cell 

interconnect to minimize failures due to material handling during the solar cell frame installation. One additional 

benefit of the tape is that it provides more intimate contact of the copper during the welding process, which 

produces better welds. 

The last improvement that was investigated was to increase the spacing or pitch of the laser welds. This was 

attempted to reduce some of the localized delaminations of the PVD copper from the substrate in the heataffected zone of the weld area. By increasing the pitch of the welds the issue of delaminations was eliminated. 

Decreasing the amount of welds also decreases the amount of current carrying capacity. The electrical load 

analysis was reevaluated to verify that the amount of welds was adequate to support the electrical loads. 

Flight Experiment

Building upon the successful development of the solar cell interconnect a flight experiment will be used to 

validate the wrap around contact and associated PowerSphere technologies. All of the design improvements 

were incorporated into the Material International Space Station Experiment 5 (MISSE 5) flight experiment. The 

experiment consisted of three a-Si solar cells laser welded to a flex circuit and embedded in a composite frame as 

shown in Figure 6. 

Flex Blanket

Egress Wires

Composite Frame

a-Si Solar Cell

Wrap Around Contact 

(Embedded)

Figure 6- MISSE 5 Flight Experiment 

The MISSE 5 experimental coupon will be mounted to the exterior of the International Space Station for an 

approximately one-year exposure. On orbit IV curves will be generated from the experiment to better understand 

the overall exposure to a LEO environment. The ability to demonstrate new technologies operating in the actual 

environmental conditions will prove to be an invaluable tool for future designs. Once the experiment is returned 

the materials will be evaluated and the effects of the LEO environmental exposure will be documented. 

Significant improvements have helped to create a successful design and fabrication process for the 

PowerSphere solar cell interconnect. Fabrication of the MISSE 5 flight experiment helped to improve upon the 

initial concept and create an optimized design as shown in Figure 7. 

NASA/CP—2005-213431 206
Back Contact

Upilex®

Kapton® Flex

Laser Weld

Solar Cell Wraparound Contact

Copper Trace

Top Contact

PVD Copper

Chromium

Kapton® Tape

Kapton® Tape

Weld Direction

Laser Scribe

Figure 7- Optimized Interconnect Design Cross Section 

In the optimized design the laser first makes contact with the Kapton® and then fully penetrates all layers with 

the plasma exiting the Kapton® tape on the opposite side. Once the solar cell interconnect has been welded, a 

frame will encapsulate the weld area and further reinforce the joint. A patent application has been filed with the 

United States patent office for the solar cell contact design co-developed by Aerospace Corp. and Lockheed 

Martin Corp. 

Summary and Conclusions

The significant advancements in solar cell interconnect design and fabrication has provided a valuable 

building block for the PowerSphere program. Lessons learned from preliminary concepts, MISSE 5 flight 

experiment and thermally cycled coupons have provided a robust design. Future testing of the complete 

engineering design unit assembly will verify that the wrap around design is an effective concept for both 

electrically and mechanically joining a-Si solar cells to flex circuit. 

Acknowledgment

This work was supported by NASA contract NAS3-01115. 

References

[1] A. Prater, E.J. Simburger, D.A. Smith, P.J. Carian, and J.H. Matsumoto, The Aerospace Corporation, “Power 

Management and Distribution Concept for Microsatellites and Nanosatellites,” Proceeding of IECEC, August 1-5, 

1999. 

[2] D.G. Gilmore, E.J Simburger, M.J. Meshishnek, D.M. Scott, D.A. Smith, A. Prater, J.H. Matsumoto, and M.L. 

Wasz, The Aerospace Corporation, “Thermal Design Aspects of the PowerSphere Concept,” Proceedings of 

Micro/Nano Technology for Space Application Conference, April 11-15, 1999. 

[3] E.J. Simburger, J.H. Matsumoto, T.W. Giants, M. Tueling, J. Ross, J.K. Lin, C. Knoll, S. Rawal, A.R. Perry, C. 

Marshall, D.M. Barnett, T.T. Peterson, T.W. Kerslake, and H. Curtis, “Development of Flex Circuit Harness for the 

PowerSphere Concept,” 29th Photovoltaics Specialist Conference, IEEE, New Orleans, LA, May 20-24, 2002. 

[4] E.J. Simburger, J.H. Matsumoto, D.A. Hinkley, D.G. Gilmore, T.W. Giants, and J. Ross, “Multifunctional 

Structures for PowerSphere Concept,” 42nd AIAA/ASME/ASCE/ASC Structures, Structural Dynamics, and 

Materials Conference & Exhibit, Seattle, Washington, April 16-19, 2001. 

[5] E.J. Simburger, J.H. Matsumoto, J.K. Lin, C.Knoll, S. Rawal, A.R. Perry, D.M. Barnett, T.T. Peterson, T.W. 

Kerslake, and H. Curtis, “Development of a Multifunctional Inflatable Structure for the PowerSphere Concept,” 

AIAA-2002-1707 43rd AIAA/ASME/ASCE/ASC Structures, Structural Dynamics, and Materials Conference & 

Exhibit, Denver, Colorado, April 22-25, 2002. 

NASA/CP—2005-213431 207
RESEARCH TRENDS TO ENHANCE THE RADIATION TOLERANCE OF Si SPACE SOLAR 

CELLS 

A. Khan and Al-Khatib 

 Department of Electrical and Computer Engineering, 

University of South Alabama, Mibile, AL 36688 

M. Yamaguchi 

 Toyota Technological Institute, Nagoya, Japan 

T. Abe 

 Shin-Etsu Handotai, Gumma, Japan 

M. Imaizumi and S. Matsuda 

 National Space Development Agency of Japan, Tsukuba, Japan 

1. Introduction 

Previously, we developed an analytical model [1], which could be adequately used for 

improving the radiation tolerance of silicon space solar by varying the base carrier concentration 

[Figure 1]. However, the optimization of the base carrier concentration reduces the pre-radiation 

minority-carrier diffusion length in the base , therefore there have to compromise between beginning 

of the life (BOL) and end of the life (EOL) performance of the space solar cell.

 The fact remain that achieving still higher radiation tolerance require that use of alternatives 

approaches. The interaction between vacancies, self-interstitials (Sii

), impurities, and dopants in Si 

leads to formation of undesirable point defects such as recombination and compensator centers 

which effect performance of the solar cells in space. In our previous studies [2, 3], the EC-0.18-

0.21eV center is found to especially plays a dominant role in carrier removal, which is related to 

interstitial-boron and interstitial-oxygen complex (Bi

-Oi) [4,5]. This implies that using different dopants 

and less oxygen crystals such as FZ-grown Si could reduce carrier removal effect. Furthermore, Gadoped Czochralski (CZ) grown Si solar cells show less light-irradiation degradation as compared to 

boron-doped CZ-grown cells [6]. This leads to the fact that Ga-doped CZ-grown solar cells may have 

more potential compared to B-doped CZ grown Si solar cells, In addition, understanding the origins of 

radiation-induced deep levels in Si, especially interaction of Si interstitials with impurities in Si, is very 

important to further improve radiation-tolerance of Si space solar cells. 

No effort has apparently yet been made to limit the effect of carrier removal by deliberately 

doping with different dopants in order that the defects generated should be less electrically active In 

this study, we first present the results of a comparison of radiation-induced defects in n+

-p-p+ CZgrown Si diodes doped with gallium or boron ranging in resistivity from 0.33-10 Ω cm. Next, we will 

NASA/CP—2005-213431 208
Fluence (cm-2)

0

2

4

6

8

10

12

14

16

Optimum Base Carrier Concentration (cm-3)

AM0 Optimum Conversion 

Efficiency (%)

1014 1015 1016 1017

0

1014

1015

3 x 1015

 1016

3 x 1016

 1017

1018

 

Fig. 1. Variation of the optimum base carrier concentration and efficiency as a function of 

1-MeV electron fluence from 1014 cm-2 to 1018 cm-2. 

give a detailed account of the impact of impurities, such as gallium, boron, carbon and oxygen on

radiation-induced defects (i.e. Bi-Oi, Ci-Oi, V-V) and hierarchy of defect reaction in silicon. 

2. Experimental details 

The fundament of the current investigation of impurity effects on radiation-induced defects in

Si is a set of wafers grown by Shin-Etsu Handoti, Japan. The different types of boron and gallium 

doped wafers grown either with the Czochralski (CZ), magnetic Czochralski (MCZ) or float zone (FZ)

method with doping and oxygen concentrations were used. The n+

/p/p+ solar cells structure was

similar to the structure used in previous studies. Area of 2 x 10-6 ~ 4 x 10-6 m2 was defined in order to

keep the capacitance of the device within the range of the capacitance meter used for C–V

measurements. The carrier concentration was also measured using Hall effect measurements at 

room temperature. The samples were irradiated in the open circuit condition with 1 MeV electrons and

10 MeV protons using a Cockcroft-Walton accelerator and Tandem accelerator, respectively at Japan 

Atomic Energy Research Institute (JAERI), taking care to avoid sample heating by appropriate control 

of the beam current and water cooling of the sample mount. Additionally, the electron and proton

beam was scanned in order to insure a uniform irradiation over the whole surface of the sample. 

3. Results and Discussion 

A. Suppression of recombination and compensator centers by Ga impurity 

Figure 2 shows a minority-carrier emission (injection pulse) spectra of 9 and 10 Ωcm, B and

Ga-doped CZ-grown silicon irradiated to a fluence of 3 x 1016 electrons cm-2. Comparison with

published data shows that the hole levels EV+0.36 eV is most probably due to the interstitial-carboninterstitial-oxygen complex (Ci

-Oi) [2, 3, 5]. The electron level EC-0.18 eV is one of the highest

concentration defects that has been commonly observed only in B-doped CZ grown Si samples,

while this level is found to be totally absent in Ga-doped CZ-grown samples (Fig. 2). The absence of

this level in Ga-doped Si gives support that this center in B-doped Si is related to Bi-Oi [ 3, 6] complex. 

As we have previously demonstrated that the hole center EV+0.36 eV acts as recombination 

center and has a dominant role to control lifetime is Si solar cells [2]. It is worth noting that the overall

concentration of EV+0.36 eV is strongly suppressed in Ga-doped samples as compared to B-doped

samples, as shown in Fig. 2. This implies that the minority carrier lifetime degradation can be

improved by using Ga as the dopant in Si space solar cells. 

NASA/CP—2005-213431 209
0 50 100 150 200 250 300

0.11 eV

0.1 7 e V

0.1 8 e V

0.3 6 e V

G a : 9 Ω c m

B : 10 Ω c m

DLT

S Signal (a. u.)

Temperature (K)

Fig. 2. Comparison of the minority carrier DLTS spectra measured for boron-or 

gallium –doped CZ-grown Si at same gallium and boron doping levels, 

irradiated with 1-MeV 3 x 1016 electrons/cm2

. 

One of the most interesting and technologically important feature of our work is the

disappearance of the dominant donor like electron level EC-0.18 eV in Ga-doped CZ- grown samples.

As we have discussed in previous studies [2], this level acts as a compensator center, which is 

positively charged before electron capture. The concentration of this level is about 60% of the change

in carrier concentration after heavy fluences and therefore populous enough to be the dominating 

influence on device performance. This implies that carrier removal effects can be partially offset by 

using gallium as dopant instead of boron. 

B. Carrier concentration analysis: using different dopants species 

Our detailed carrier concentration analysis show that carrier removal effects can be partially offset 

by using gallium as dopant instead of boron. For low fluences, the change in carrier concentration 

can be expressed by the following expression: 

p0 − pφ ≈ Rcφ (1) 

Where po and pφ are the carrier concentrations before and after irradiation with fluence φ of 1 MeV

electrons and RC is the carrier removal rate. However, for high fluences it can be approximated by, 















 − ≈

0

0 exp

p

R

p p Cφ

φ (2) 

NASA/CP—2005-213431 210
Equation (2) is equivalent to Equation (1) in the range of low fluence. The changes in carrier

concentration determined from C–V and Hall effect measurements as a function of 1 MeV electron

fluence, with pre-irradiation carrier concentration p0 of 1.15 x 1015 cm-3 are shown in Fig. 3. Fitting of

1016 1017 1018 1012

1013

1014

1015

1016

RC=0.15 cm-1 (best fit)

Si:B

Fig. 4. Calculated and experimental changes in the carrier concentration of 

p-Si boron doped crystals with initial carrier concentrations of 1.15 x 1015 

cm-3, as a function of 1-MeV electron fluence. 

Fig. 3. Calculated and experimental changes in the carrier concentration of p-Si 

gallium doped crystals with initial carrier concentrations of 1.15 x 1015 cm-3, as a 

function of 1-MeV electron fluence. 

Si:Ga

Carrier concentration (cm-3

)

1 MeV electron fluence (cm-2)

experimental data

 RC = 0.04 cm-1 (best fit)

 RC = 0.02 cm-1

 RC = 0.06 cm-1

NASA/CP—2005-213431 211
Eq. (2) gives the value of RC = 0.04 cm-1 that is almost four times lower than that for boron doped Si

(0.15 cm-1), as shown in Fig. 4 with initial carrier concentration of 1.51 x 1015 cm-3. This is a highly

technologically interesting new result on the properties of the gallium as dopant in Si. Furthermore,

this result is consistent with deep level transient spectroscopy (DLTS) analysis (Fig. 2 ), which

showed that introduction of gallium as dopant in Si provides a simple method for selectively bleaching

the dominant compensating center (Bi

-Oi). 

The damage coefficient KL determined from the spectral response, by fitting the numerical

simulation, is 8 x 10-11 for 10 Ωcm boron doped Si solar cells and 4 x 10-11 for gallium doped Si solar 

cells. These values are consistent with the degradation behaviors of the solar cells. This is further 

supporting evidence that gallium doped silicon solar cells showed better radiation tolerance compare

to boron doped silicon solar cells. 

C. Annealing behavior of radiation-induced defects in Ga-doped Si 

Isochronal thermal annealing was performed on gallium doped silicon samples after exposing

them with the same fluence of 3 x 1016 cm-2 1 MeV electron to investigate (i) the effects, if any, of the

presence of gallium as dopant on the thermal stability of the radiation-induced defects (ii) to monitor 

the evolution of new defect following annealing of first order point defects, such as vacancy-interstitial

and their complexes with impurities. These three aspects together would be expected to provide a

comprehensive comparative picture of the radiation-induced defects in boron and gallium doped

silicon, which can, in turn shed light on the possible formation of their complexes. 

Isochronal annealing was performed for 15 minutes in steps of 25-30°C in nitrogen ambient,

each followed by DLTS measurements. The concentration data obtained for the various deep levels

from these experiments are plotted in Fig. 5 as a function of the isochronal annealing temperature up

to 400 °C. 

Comparison with the isochronal annealing curve data for the radiation-induced levels in the 

boron doped samples [3] show that gallium not only strongly suppressed recombination centre

(EV+0.36) eV but also annealing curve are found to show different anneal-out characteristics to that in

100 200 300 400

1012

1013

1014 Si:Ga

Defects concentration (cm-3

)

 EV+0.17 eV

 EV+0.36 eV

 EC-0.14 eV

 EC-0.39 eV

 EV+ 0.55 eV

 EV+0.18 eV

Annealing temperature (oC)

Fig. 5. Concentration of the annealed-out and annealed-in deep levels in 1 

MeV ( 3 x 1016 cm-2 ) electron irradiated Ga-doped Si diodes and space 

solar cell as a function of isochronal annealing temperature. Each annealing 

step was 15 min duration.

NASA/CP—2005-213431 212
boron samples. One important difference in the reverse annealing behaviour (300 o

C-350 oC), which

is negligible in gallium doped samples and significant in boron doped samples [3]. This is further 

supporting evidence to the fact that reverse annealing (an increase in concentration) of Ci

-Oi

complex (EV+0.36 eV) in boron doped samples is due to an increase in the concentration of Bi

-Cs 

complex [6] at EV+0.30 eV following annealing of donor like Bi-Oi center (EC-0.18 eV), which acts as 

a minority trap in boron doped silicon. It is worth noting that neither of these levels ( Bi-Cs, Bi-Oi) has

been observed in gallium-doped samples, consequently reverse annealing [3] is almost absent in

gallium-doped samples. The EV+0.36 eV level both in boron doped samples and gallium doped

samples remains almost unchanged up to 400 oC annealing step as the Ci-Oi complex is the one of 

the most stable complexes of all the observed interstitials related deep levels in silicon. 

The annealing behaviour of EV+0.17–0.18 eV level which is associated divacancy in galliumdoped samples is found to show almost identical anneal-out (375-400 o

C) characteristics of this level

to that in the boron-doped samples, showing that different dopant does not alter its characteristics. 

As seen in Fig. 5, the dominant level to anneal in earliest in gallium-doped samples is

EV+0.18 eV. It shows up at 150 oC and raises two orders of magnitude with the increase in annealing

temperature, reaching a maximum in concentration (1 x 1014 cm-3) at 325 oC. Thereafter, it suddenly

anneal-out as shown by a sharp decline in its concentration (Fig. 5). 

3. Further Discussion 

A. tentative identification of the recombination and compensator centers 

The identification of the dominant hole level EV+0.36 eV observed in our electron and proton

irradiated samples has been less straightforward. However, our comprehensive study throws new 

light on the formation mechanism of this defect and fixed its identification as Ci-Oi complex. A

summary of the boron concentration dependence of the introduction rates of EV+0.36 eV defect

together with other room temperature stable defects in 10 MeV proton irradiated p-type silicon is

shown in Fig. 6. In all these Czochralski (CZ) grown samples the concentration of interstitial oxygen

(Oi

) is on the order of ~9 x 1017 cm-3. 

1015 1016 1017

0.01

0.1

1

10

100

Oi ~ 9 x 1017 cm-3

C ~ 5 x 1015 cm-3

10 MeV proton

Φ = 3 x 1013 cm-2

Introduction rate (cm-1

)

Boron concentration (cm

3

)

Bi-Bs (0.30 eV)

Ci-Oi (0.36 eV)

Bi

-Oi (0.18 eV)

Fig. 6. Variation in the introduction rates of the interstitial related 

defects in 10 MeV proton irradiated p-Si as a function of 

background impurity concentration. 

NASA/CP—2005-213431 213
In the first experiment we investigated the defects introduction rates as a function of boron

concentration as shown in Fig. 6. The introduction rate of EV+0.36 eV defect decrease linearly as the

boron concentration (Bs) increases (Fig. 6). It is important to point out here, in this concentration

region, the electron level EC-0.18eV is increasing, being consistent with its identification (Bi

-Oi). This

implies that EV+0.36 eV is in direct competition with the production of Bi-Oi. Higher introduction rates 

of Bi-Oi strongly suppressed the production of the Ci-Oi (EV+0.36) defect. 

In order to provide the additional information about the nature of the defects the influence of 

interstitial oxygen has been studied quantitively, CZ, FZ, and MCZ grown samples of almost same

boron concentration were chosen. The EV+0.36 eV introduction rate increases with increasing Oi as

An analogous investigation was performed for the influence of carbon on the introduction rates 

shown in Fig. 7. The conclusion is clearly justified that oxygen is one of the component of this defect. 

of E

rved to have one of the highest concentrations in our low 

dose

1016 1017 1018 1019 10-2

10-1

100

101

10 MeV proton 

Φ = 3 x 1013 cm-2

Boron concentration = 2~3 x 1015 cm-3 

Introduction rate (cm-1

)

Oxygen concentration (cm-3)

Bi-Oi (0.18 eV)

Ci

-Oi (0.36 eV)

Fig. 7. Variation of the introduction rates in boron-doped silicon ( irradiated with

10 MeV proton) as a function of interstitial oxygen concentration. 

V+0.36 eV defect. Now samples with interstitial oxygen concentration between 6.2 to 9.4 x 1017

cm-3 and boron concentration between 1.2–1.9 x 1016 cm-3 were selected. There is approximately

linear relationship between the EV+0.36 eV defect and carbon concentration as presented in Fig. 8. 

These observations that carbon and oxygen are the major components of the EV+0.36 eV defect in

electron and proton irradiated silicon fixed its identification as Ci

-Oi complex. It is worth noting that 

with increasing carbon concentration EC-0.18 eV (Bi

-Oi ) is decreasing due to direct competition with 

the production of Ci

-Oi (EV+0.36) complex. 

This defect (EC-0.18-0.20 eV) was obse

samples up to low/intermediate fluence. As presented in Fig. 6 the introduction rate of minority

level (EC-0.18 eV) for 10 MeV proton irradiated CZ-grown samples in which oxygen content is 

constant at ~9 x 1017 cm-3 increases linearly with boron concentration at lower doping level

(~1016 cm-3 Obviously, the boron concentration plays an important role in the formation of this center.

Kimirling and Mooney [6, 7] have also observed this phenomenon in 1 MeV electron irradiated

samples. At higher boron doping level there is competition between EC-0.18 eV and Bi-Bs (Fig. 6) and 

consequently introduction rates of EC-0.18 eV center decrease following evolution of Bi

-Bs [6]. Figure 

7 also show that the introduction rates of EC-0.18 eV also increases linearly with oxygen

concentration for intermediate boron level. The linear dependence of the introduction rate on both 

boron and oxygen firmly justify the hypothesis that boron and oxygen are the major component (Bi-Oi) 

of this compensator center (EC-0.18 eV) in silicon. Moreover, the absence of this level in Ga-doped 

NASA/CP—2005-213431 214
silicon demonstrates that this center in boron-doped silicon is related to Bi

-Oi complex. This 

.

As eads to

stron

 with

sub

identification is consistent with the previous speculations [6, 7] and fixed its identification as Bi-Oi

complex.

B Possible reasons behind the different role of boron and gallium in Si 

t in silicon l

diation interchanges its location

→ → + → −

n energy of replacing Sii

 with impurities is thought to be due to atomic size

rder to understand the effects of Ga atoms upon ge

we have discussed in previous sections the presence of gallium as dopan

g suppression of the recombination center EV+0.36 eV and complete removal of the dominant

compensator center EC-0.18 eV (Bi-Oi). This implies that boron has a strong affinity for pairing with

oxygen in p-type silicon like well known reported boron iron complex (Bs-Fei) [8]. The possible

interpretations for this new observation are discussed below. 

It is known that Si self-interstitial (Sii) produced by ra

stitutional boron (Bs), carbon (Cs) [9,10] and gallium (Gas), results in interstitial B (Bi), C (Ci) and

Ga (Gai), and forms complex defects such as Bi-Oi, Ci-Oi and Gai-Oi by combining with interstitial 

oxygen (Oi

) as follows: 

Sii + Bs → Bi → Bi + Oi → Bi − Oi

Sii + Cs → Ci → Ci + Oi → Ci − Oi

Sii + Gas Gai Gai Oi Gai Oi

The activatio and

cha

neration 

1015 1016 1017 1018

10-2

2

1

Boron ~ 1.2 1016 -1.9 x 1016 cm-3

Oi ~ 6 x 1017- 9 x 1017 cm-3

Introduction rate (cm-1

)

Carbon concentration (cm-3)

 Bi-Oi

 Ci-Oi

Fig. 8. Influence of carbon concentration on the defects productions in Si 

samples irradiated with 10 MeV proton. 

rge state of impurities. In fact, the size of the tetrahedral covalent radius is different for B

(0.088nm), C (0.077nm), Si (0.117nm) and Ga (0.126nm). Therefore, the Ga atom with larger size is 

thought to produce more lattice stress as compared to B and C, and thus influence on the probability 

of generation of complex defect such as Gai

-Oi and the activation energy for generating Gai-Oi must 

be larger than those of Bi

-Oi and Ci-Oi. 

However, further study is necessary in o

and annihilation of radiation-induced defects. 

NASA/CP—2005-213431 215
5. Summary 

The present study fully explored the better radiation tolerance of gallium doped silicon solar cells 

as 

 drawn a quite complete picture towards the identification of the dominant 

rad

. Reference 

1 M. Yamaguchi, A. Khan, S. J. Taylor, M. Imaizumi T. Hisamatsu and S. Matsuda, IEEE 

2 Matsuda, and T. Aburaya, J. 

3 samatsu and S. Matsuda, J. Appl. Phys. 87 2162 (2000). 

 3836 

5 mbetta. and G. D. Watkins, Appl. Phys. Lett. 51, 1103 (1987). 

st PVSEC-11, 549 

7 erling, IEEE Trans. Nucl. Sci. Ns-23 1497 (1976). 

87). 

compared to conventional boron doped silicon solar cells after heavy fluence of 1 MeV electron 

and 10 MeV proton irradiation. Analysis of the carrier removal rate RC in boron, and gallium doped 

Si solar cells showed that carrier removal effects can be partially offset by using gallium as dopant 

instead of boron, which demonstrates the better radiation tolerance of gallium doped Si solar cells as 

compared to boron doped. 

Present study have also

iation-induced defects in Si. The radiation-induced defect EV+0.36 eV has been identified as Ci-Oi

complexes. The absence of Ec-0.18 eV complex center in gallium doped samples and linear 

dependence of its introduction rates on both boron and oxygen content fixed its identification as the 

Bi

-Oi complex in boron doped Si. One of the technologically important results of present study is that 

the gallium appears to strongly suppress the radiation induced defects, especially hole level EV+0.36 

eV (Ci

-Oi), which is thought to act as a recombination center as well as the dominant compensating 

center at EC-0.18 eV (Bi-Oi). As a result the effects of lifetime degradation and carrier removal could 

be partially offset to higher radiation fluences by using Ga as dopant instead of boron in Si space 

solar cells. These results reveal that gallium, as dopant in Si, is a promising material for next 

generation radiation hard space solar cells. 

6

Transactions on Electron Devices Vol. 46, No. 10, 2133 (1999). 

M. Yamaguchi, A. Khan, S. J. Taylor, K. Ando, T. Yamaguchi, S.

Appl. Phys. 86, 217 (1999). 

A. Khan, M. Yamaguchi, T. Hi

4 P. M. Mooney, L .J. Cheng, M. Sull, J. D. Gerson, and J. W. Corbett, Phy. Rev. B 15,

(1977). 

J. M. Tro

6 S. W. Glunz, S. Rein, W. Warta, J. Knobloch and W. Wettling, Tech. Dige

(1999). 

L.C. Kim

8 K. Graff and H. Pieper , J. Electrochem. Soc. 128, 660 (1981). 

9 J.M. Trombetta and G.D Watkins, Appl. Phys. Lett. 51, 1103 (19

10 G.D. Watkins, Phys. Rev. B12, 5824 (1975). 

NASA/CP—2005-213431 216
USING SRIM TO CALCULATE THE RELATIVE DAMAGE COEFFICIENTS FOR SOLAR CELLS 

Scott R. Messenger 

SFA, Inc., Largo, MD 20774 

Edward A. Burke 

Consultant, SFA, Inc., Largo, MD 20774 

Robert J. Walters and Jeffrey H. Warner 

Code 6818, Naval Research Laboratory, Washington, DC 20375 

Geofffey P. Summers 

Code 6818, Naval Research Laboratory, Washington, DC 20375 

Department of Physics, UMBC, Baltimore, MD 21250 

ABSTRACT 

It is shown how the widely available computer code SRIM can be used to calculate proton relative damage 

coefficients (RDCs) for solar cells. This approach is essential in the case of crystalline Si cells where the incident 

proton energy is reduced significantly in traversing the active volume of the cell. However, the same approach can 

be used in other cases such as GaAs-based multi-junction cells where the active region consists of several 

different material layers. In the case of Si cells analytic calculations of the non-ionizing energy loss (NIEL) are not 

satisfactory for calculating the energy dependence of the RDCs in contrast to technologies such as single junction 

GaAs and thin films such as CIGS, for which the NIEL approach has been extremely successful. In the SRIM 

approach it is assumed that the RDCs are proportional to the number of total vacancies formed. The calculation 

involves integrating the vacancy.txt files produced by SRIM over the active depth of the cells and normalizing the 

results to the number of vacancies produced by 10 MeV protons. The RDCs calculated from SRIM are found to 

agree well with experimentally measured values for both Si and MJ cells. In particular the “double hump” 

structure in the measured RDCs for several kinds of MJ technologies can be reproduced. 

INTRODUCTION 

Relative damage Coefficients (RDCs) are needed to estimate solar cell degradation in the complex proton and 

electron environment found in space. The most straightforward way to determine the RDCs is to measure them 

directly. In a typical case such an undertaking involves measuring cell degradation curves for various 

photovoltaic parameters for a range of incident proton and electron energies. Anspaugh [1] produced an 

excellent data set of this kind for single junction GaAs cells. The RDCs for normal incidence monoenergetic 

particles can be converted to the corresponding RDCs for an isotropic environment, from which the RDCs for 

shielded cells can be derived using the widely used approach developed at JPL and described in the Solar Cell 

Radiation Handbooks and elsewhere [2,3]. The problem with this approach is that acquiring the necessary data 

is extremely labor intensive and consequently expensive. 

In a series of papers over the last decade NRL showed that in many cases the energy dependence of the RDCs 

could be determined analytically using calculations of the nonionizing energy loss for particular particles and cell 

materials [4 and references therein]. These calculations have become progressively more sophisticated for 

electrons and protons, and have been recently extended to incident heavy ions [5]. When the NIEL is multiplied 

NASA/CP—2005-213431 217
by a particle fluence a quantity called displacement damage dose is obtained, which plays the same role for 

displacement damage that ionizing dose plays for ionization effects. When Anspaugh’s data for GaAs was 

converted to displacement damage dose it was found that the curves for E >10keV collapsed to a single 

characteristic curve. Indeed it was primarily the availability of Anspaugh’s excellent data set that enabled the 

displacement damage dose approach to cell degradation to be developed. 

The NIEL approach works well when the particle energy can be assumed to be constant across the active region 

of the device. This is the case for GaAs-based cells, where the active region can be considered thin for proton 

energies >~100 keV. In these cases the energy dependence of the NIEL and the relative damage coefficients 

(RDCs) are similar [4]. However, this is not true for very low energy protons incident on uncovered GaAs cells, 

and deviations between the energy dependence of the NIEL and the RDCs are observed at proton energies < 100 

keV [1]. The reason for the deviation is that protons with these energies slow down significantly or stop in the 

active region of the cell. 

For silicon cells, where the active region in typical cells is 100 times greater than is the case for GaAs, the 

problem is exacerbated. In GaAs cells, the active region is typically on the order of a few micrometers in depth 

from the front surface of the cell, as opposed to ~100 μm for crystalline silicon cells. This means that protons with 

relatively high energy will slow down significantly or reach the end of their tracks in the critical silicon cell regions 

more so than is the case in GaAs cells. The slowing-down effects complicate the analysis. What is more, 

experiments show that, for normally incident protons of a given energy, the damage per incident proton is a 

maximum when the particle track terminates in the active device region. Whereas for GaAs [1,6] the energies 

where this is a problem exist only for protons below a few hundred keV, it extends to several MeV in the case of 

silicon [2,3]. 

RDCs FOR CRYSTALLINE Si CELLS 

The observations made above can be quantified by analyzing actual experimental results. Figure 1 shows data 

for the degradation of Pmax for Si cells taken from the Solar Cell Radiation Handbook [2,3]. Results are shown 

for incident 100,and 300 keV, and 1 and 9.5 MeV protons. The two most striking features of these data are firstly 

that the curves are not similar in shape and secondly the curves do not come in any obvious order with increasing 

109 1010 1011 1012 1013

Proton Fluence (p+/cm2)

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

Normalized Maximum Power

100 keV

300 keV

1 MeV

10 MeV

Si

Proton Energy

Figure 1. The degradation of the maximum power point of Si cells as a result of proton 

irradiation [2,3]. The points of the curves are for identification only and are not measured data. 

NASA/CP—2005-213431 218
proton energy. The data for Si cells is in contrast to similar data for the degradation of Pmax in GaAs cells taken 

from Anspaugh and shown in Fig. 2. In this case the curves are similar in shape except for the two lowest 

energies (50 and 100 keV) and there is a clear trend with energy, with the lowest energy protons causing the 

most damage as would be expected from the interaction cross sections. It is the behavior of the 50 and 100 keV 

curves in Fig. 2 that suggests the cause of the results shown in Fig. 1. As noted above 50 and 100 keV protons 

do not completely traverse the active region of single junction GaAs cells so that the damage is done in these 

cases differently than for higher energies. Once the proton energy is > ~200 keV the damage is done uniformly 

and the degradation curves become similar in shape. For the data shown in Fig. 1 only in the case of the 9.5 

MeV curve do the incident protons traverse the active region of the Si cell. For the other energies the protons 

stop in different parts of the active region. 

109 1010 1011 1012 1013

Proton Fluence (p+/cm2)

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1.0

Normalized Maximum Power

50 keV

100 keV

200 keV

300 keV

500 keV

1 MeV

3 MeV

9.5 MeV

GaAs

Proton Energy

Figure 2. The degradation of the maximum power point of GaAs cells due to proton 

irradiation [1]. The points on the curves are for identification only and are not measured data. 

There is an added complication that arises when the degradation curves are not similar in shape and that is in the 

determination of the RDCs. In the JPL approach, the RDCs are determined by comparing the fluences of 

different energy protons that cause the same relative damage. Typically the 75% degradation point is used. If 

the curves are not parallel obviously the RDCs will vary if a different degradation point is chosen. 

Figure 3 shows the RDCs for Si cells determined from the kind of data shown in Fig .1. For comparison the 

energy dependence of the calculated NIEL for protons in Si is also plotted on the figure [7], with the two curves 

normalized at 10 MeV. It is clear that the NIEL only tracks the measured RDCs for energies where the slowed 

down energy in the active region Eso is comparable to the incident energy (Einc). 

It is reasonable to assume that the active region of the cells that produced the data shown in Fig. 1 ranged from 

~0.8 – 80 μm from the front of the cell. SRIM [8] runs then put the discussion above on a more quantitative basis 

as can be seen in Fig. 4. Figure 4 was obtained from the vacancy.txt output of SRIM, which has been converted 

to NIEL using the Kinchin-Pease approximation. The SRIM studies show that at energies below about 3 MeV the 

proton stops within the active region. 

NASA/CP—2005-213431 219
10-1 100 101 102

Proton Energy (MeV)

10-1

100

101

102

Relative Damage Coefficients

Proton NIEL

JPL Damage Coefficients

*All data normalized to 10 MeV protons

ESD<EINC ESD≈EINC

Si

Figure 3. Comparison of the calculated NIEL for protons in Si [7] with the measured RDCs for maximum 

power degradation [2,3 ]. In the figure, Einc refers to the incident proton energy and Eso refers to the 

slowed down energy. 

10-1 100 101 102 103

Depth (μm)

10-3

10-2

10-1

100

NIEL(keV/μm)

400 keV

1 MeV

4 MeV

10 MeV

Si

Proton Energy

Figure 4. SRIM derived NIEL as a function of the depth for an incident proton in Si [8] . 

The simplest approach to quantitatively account for the spatial nature of the damage is to assume that the RDC 

changes in direct proportion to the total damage produced within a minority carrier diffusion length of the junction. 

The energy deposited to displacements at low energies can be estimated by numerically integrating the SRIMderived NIEL over the particle track. The adjusted NIEL is then calculated by dividing the total deposited energy 

NASA/CP—2005-213431 220
by the width of the active region (80 μm). At high energies, where the protons traverse the cell with little change in 

energy, the damage factor is assumed to be proportional to the displacement damage energy deposited. This is 

given by the product of NIEL and the width (approximately 80 μm). The NIEL for 10 MeV protons in silicon is 

7.885x10-3 MeVcm2

/g [5]. This can be expressed as 1.829x10-3 keV/mm or 3.33x10-2 vacancies/mm [5]. The latter 

units are convenient for direct comparison to SRIM results, which are usually expressed as vacancies/mm The 

low-depth tail of the 10 MeV SRIM-derived NIEL indeed agrees with the analytical value obtained in [7]. 

The results obtained by applying this procedure (and normalizing at 10 MeV) are given in Fig. 5 and are 

compared with the data given in [3]. Inspection of Fig. 5 shows reasonable agreement between the calculated and 

the experimentally determined RDCs. The general shape of the calculated curve is not very sensitive to the exact 

value used for the diffusion length. As shown in the Fig. 5, an active depth value of 100 μm does not alter the 

shape or the curve significantly. 

10-1 100 101

Proton Energy (MeV)

100

101

102

Relative Damage Coefficients

Proton NIEL

JPL Damage Coefficients

SRIM Results (80 μm)

SRIM Results (100 μm)

*All data normalized to 10 MeV protons

Si

Figure 5. Comparison of the SRIM-derived RDCs for Si cells and the experimentally determined values, 

(see Fig. 3). 100 μm and 80 μm refer to the active depth assumed in the calculation. 

. 

RDCs FOR MULTI-JUNCTION CELLS

A similar analysis can be performed for the case of multi-junction cells. In this case the added complexity comes 

from the fact that the active region of the cells extends over several different materials; InGaP, GaAs and Ge. 

Figure 6 shows the degradation curves for Pmax for a three junction InGaP/GaAs/Ge cell technology [ ]. It can be 

seen that for E >400 keV the curves have similar shapes and follow the expected variation with energy. For lower 

energies the curves show the familiar variation that occurs when the incident protons slow down or stop in there 

active volume of the cells. Figure 7 shows experimentally determined [10,11] RDCs for Pmax degradation for 

several kinds of MJ cells compared to calculations of the NIEL for protons in GaAs. It should be noted that the 

energy dependence of the NIEL for GaAs is almost identical to that for Ge, and very close to that for InGaP. An 

obvious feature of the experimental data is the double hump structure seen ~0.1 MeV where the data starts to fall 

below the NIEL. At higher energies there is good agreement between the experimental data and the calculated 

NIEL. 

NASA/CP—2005-213431 221
108 109 1010 1011 1012 1013

Proton Fluence (p+/cm2)

0.5

0.6

0.7

0.8

0.9

1.0

Normalized Maximum Power

50 keV

100 keV

200 keV

400 keV

1 MeV

5 MeV

10 MeV

InGaP/GaAs/Ge 3J Solar Cell

Proton Energy

Figure 6. The degradation of the maximum power point of InGaP/GaAs/Ge triple junction 

cells due to proton irradiation [ ]. 

10-1 100 101 102

Proton Energy (MeV)

10-1

100

101

102

Pmax Relative Damage Coefficients

NIEL

SL 2J

SL 2J EOL

TEC 3J

SL3J

SL 3J EOL

Emcore 3J

MJ

*Data from Marvin 2000

Figure 7. Comparison of the calculated NIEL for protons in GaAs with the experimentally 

determined RDCs for several multijunction InGaP/GaAs/(Ge) cell technologies [9,10,11]. 

It is reasonable to assume in light of the discussion above for Si cells, that the deviation of the experimental data 

from the NIEL in the double hump region in Fig. 7 is due to slowing down of the incident protons in the active 

volume of the cell. This is confirmed by SRIM runs as shown in Fig. 8. The analysis that follows is aimed at using 

these SRIM runs to try to reproduce the double hump structure in the energy dependence of the RDCs as shown 

in Fig.7. 

NASA/CP—2005-213431 222
10-2 10-1 100 101 102 103 10-2

10-1

100

101

102

103

Penetration Depth (μm)

15.8 keV

25.1 keV

39.8 keV

63.1 keV

126 keV

199 keV

316 keV

501 keV

794 keV

1.26 MeV

1.99 MeV

3.16 MeV

5.01 MeV

7.94 MeV

Vacancy Production Rate (#/

12.6 MeV

μm/p

+)

InGaP GaAs Ge

Proton Energy

Figure 8. SRIM-derived vacancy introduction rates for protons on an InGaP/GaAs/Ge 

triple junction cell. 

In the case of a cell technology consisting of one material such as Si, it was shown above that a good 

representation of the energy dependence of the RDCs could be found by determining the total number of 

vacancies produced over the active volume. When the cell consists of several different materials it is necessary 

to take into account the fact that incident protons will usually affect the electrical properties of the junctions 

differently. For example InGaP cells are more radiation resistant than GaAs cells. The different radiation 

resistance could be due to a greater level of defect annealing in InGaP than in GaAs, for example, or it might be 

due to the different electrical properties of the defects produced. The result in InGaP/GaAs cells is that the 

degradation of the overall cell is largely determined by the softer GaAs subcell. This result is confirmed by 

measurements of the quantum efficiency (QE) of each of the junctions following irradiation with the same fluence 

of protons of different energies. A fluence of 1 x 1012/cm2

 of 50 keV protons, which stop in the top surface of the 

InGaP region, decrease only the blue part of the QE of the InGaP junction, whereas 100 keV protons, which stop 

in the back region of the InGaP junction decrease the QE of the red part of the InGaP junction. In both cases the 

change in the QE measured was quite small. However, the same fluence of 400 keV protons, which stop in the 

GaAs junction, produces a large drop in the GaAs QE even though 400 keV protons are less damaging than 100 

keV protons 

In light of these results the following approach was taken. Using the SRIM results in Fig. 8, the vacancy.txt files 

for energies up to 126 keV were used to calculate the total number of defects produced in InGaP over the whole 

range of the slowed down protons. These results were normalized to the total number of vacancies produced by 

10 MeV protons. For energies >126 keV only the defects produced in the GaAs layer were counted since these 

appear to be much more effective in degrading the operation of the cell than those in either the InGaP or Ge 

layers. The results of this analysis can be seen in Fig. 9 where the SRIM-derived RDCs are compared to the 

experimentally determined values. It can be seen in Fig. 9 that the SRIM analysis produces the double hump 

structure seen in the experimental data, although the calculated RDCs in this energy region are slightly higher 

than the measured values. It should be noted that SRIM does not include contributions from inelastic proton 

interactions, which are responsible for the flattening out of the total NIEL for protons in GaAs at E > 10 MeV. The 

SRIM-derived RDCs are therefore expected to fall below the experimental determined values at these higher 

energies. 

NASA/CP—2005-213431 223
10-1 100 101 102

Proton Energy (MeV)

10-1

100

101

102

Pmax Relative Damage Coefficients

NIEL

JPL RDCs (SP 2J)

SRIM RDCs

MJ

 Figure 9. Comparison of SRIM-derived RDCs, experimentally determined values [10], and 

 the calculated NIEL for protons [7] in GaAs, normalized at 10 MeV. 

CONCLUSIONS 

It has been shown that calculations of the total number of vacancies formed in active regions of solar cells using 

the Monte Carlo code SRIM can be used to calculate proton RDCs, which agree quite well with experimentally 

determined values. The approach is particularly useful in two cases. Firstly, in crystalline Si cells where because 

of the large active region of the cell protons in the MeV range slow down significantly or stop. Secondly, in 

multijunction cells such as InGaP/GaAs/Ge where the active volume consists of several monolithic layers of 

different materials. The ability to reproduce RDCs in this way is very useful because. from this basic curve for 

unidirectional, monoenergetic incident proton beams a family of RDC curves can be generated for isotropic 

protons traversing different amounts of shielding using the approach described in detail in references [3]-[6]. The 

performance of Si or multijunction solar cells can then be predicted for any defined space radiation environment. 

Comparison of predictions made for specific space missions using SRIM-derived RDCs and those experimentally 

determined are found to agree closely. This gives confidence in the reliability of SRIM-derived RDCs for 

developmental cells where experimental values are not yet available. 

REFERENCES 

[1] B.E. Anspaugh, Proc. 22nd IEEE Photovoltaic Specialists Conf., Las Vegas, NE, 1593, (1991). 

[2] Y. Tada, J.R. Carter, Jr., B.E. Anspaugh and R.G. Downing, Solar Cell Radiation Handbook, 3rd Edition 

JPL Publication 82-69, 1982. 

[3] B.E. Anspaugh and R.G. Downing, Radiation Effects in Si and GaAs Solar Cells Using Isotropic and 

Normally Incident Radiation, JPL Publication 84-61, 1984. 

[4] S.R. Messenger, G.P. Summers, E.A. Burke, R.J. Walters, and M.A. Xapsos, Prog. Photovolt.: Res. 

Appl., 9, 103-121 (2001). 

NASA/CP—2005-213431 224
[5] S.R. Messenger, G.P. Summers, E.A. Burke, M.A. Xapsos, R.J. Walters, E.M. Jackson, and B.D. 

Weaver, IEEE Trans. Nucl. Sci., 46-6, 1595-1602 (1999). 

[6] B.E. Anspaugh, GaAs Solar Cell Radiation Handbook, JPL Publication 96-9, 1996. 

[7] G.P. Summers, E.A. Burke, P. Shapiro, S.R. Messenger and R.J. Walters, IEEE Trans. Nucl. Sci. 40, 1372 

(1993). 

[8] J.F. Ziegler, J.P. Biersack, and U. Littmark, The Stopping and Range of Ions in Solids, Volume I, New 

York: Pergammon Press, 1985. (SRIM is freely available at http://www.srim.org.) 

[9] D.C. Marvin and J.C. Nocerino, Proc. 26th IEEE Photovoltaic Specialists Conf., Anchorage, AK, 1102 - 

1105 (2000). 

[10] D.C. Marvin, Aerospace Report TOR-2000(1210)-1. 

[11] P.R. Sharps, D.J. Aiken, C.H. Thang, and N.S. Fatemi, Proc. 17th EUPVSEC, Munich, Germany, (2001). 

NASA/CP—2005-213431 225
Physics of Quantum Structures in Photovoltaic Devices 

Ryne P. Raffaelle and John D. Andersen 

Physics Department, Rochester Institute of Technology 

Rochester, NY 14623 

 

Abstract 

There has been considerable activity recently regarding the possibilities of using various 

nanostructures and nanomaterials to improve photovoltaic conversion of solar energy. Recent 

theoretical results indicate that dramatic improvements in device efficiency may be attainable 

through the use of three-dimensional arrays of zero-dimensional conductors (i.e., quantum dots) 

in an ordinary p-i-n solar cell structure. Quantum dots and other nanostructured materials may 

also prove to have some benefits in terms of temperature coefficients and radiation degradation 

associated with space solar cells. Two-dimensional semiconductor superlattices have already 

demonstrated some advantages in this regard. It has also recently been demonstrated that 

semiconducting quantum dots can also be used to improve conversion efficiencies in polymeric 

thin film solar cells. Improvement in thin film cells utilizing conjugated polymers has also be 

achieved through the use of one-dimensional quantum structures such as carbon nanotubes. It is 

believed that carbon nanotubes may contribute to both the disassociation as well as the carrier 

transport in the conjugated polymers used in certain thin film photovoltaic cells. In this paper we 

will review the underlying physics governing some of the new photovoltaic nanostructures being 

pursued, as well as the the current methods being employed to produce III-V, II-VI, and even 

chalcopyrite-based nanomaterials and nanostructures for solar cells. 

Introduction 

The underlying physics providing the impetus for the use of nanostructures in solar cells 

can be understood through elementary quantum mechanics. One of the first lessons learned in 

an introduction to the field is that electron motion is governed by Schroedinger’s Equation and the 

potential V(r) 

() () () () r V r r E r

m

∇ Ψ + Ψ = Ψ − 2

2

2

h

 (1) 

If V(r) = 0, the solution is a simple plane-wave. In regions where V is non-zero, and the electron 

has enough energy to pass over of tunnel through, we will have a modified plane-wave solution 

(scattering states). However, where V(r) is negative we will have bound (negative energy) states. 

Although scattering and bound states differ in many ways, of particular interest is the fact that 

scattering states are continuous function of the electrons energy, whereas bound states are 

possible only for particular values of the energy. That is, bound states only occur for discrete 

values of energy (i.e., they are "quantized"). If two identical atoms are brought together, with an 

electron is the same quantum state, the energy levels must split due to the Pauli Exclusion 

Principle. 

Solutions to Schroedinger’s equation for electrons in a crystalline solid are called Bloch 

wavefuntions. Fourier’s theorem shows that the functions are simply a plane-wave solution 

multiplied by a periodic function whose periodicity matches that of the potential in the crystal. 

( ) ( ) ik r r u r e ⋅ Ψ = (2) 

NASA/CP—2005-213431 226
Confining Bloch electrons to a thin plane by providing a rectangular potential well perpendicular to 

the plane will also produce a discrete bound state spectrum. Such a system is an example of a 

Quantum Well (QW). If the electron is completely confined to the plane, due to an infinitely deep 

potential well of width az, the bound state energies are 

( )

2 2

2 2

2

2 2 































+ + 













 =

z

z

x y a

n

m

k k

m

E h h π (3)

Similarily, if constrained in two dimensions we have a quantum wire with energy levels given by 

































+ 































+ 













 =

2 2 2

2

2

2 2 z

z

y

y

x a

n

a

n

m

k

m

E h h π π (4)

and if constrained in three dimension we have a quantum dot with energy levels given by 

































+ 















+ 





























 =

2 2 2 2

2 z

z

y

y

x

x

a

n

a

n

a

n

m

E h π π π (5)

In the same way that bringing a number of atoms together in a solid results in “energy bands,” 

bringing a number of quantum structures together (i.e., multiple quantum wells, ordered arrays of 

quantum wires or quantum dots) will also result in bands. If these collections of quantum 

structures reside within a host semiconductor we will have what are called “minibands” (see 

Figure 1) Electronic States in “wells” are “quantized” and therefore have discrete-like energy 

levels. The mini-band levels depend on the width of the wells. These size dependent minibands levels afford photovoltaic designers the flexibility to tune the absorption spectrum of 

devices beyond what is available through the normal array of bulk semiconductor bandgaps. In 

addition to adjusting the absorption specturm of the device, these minibands also provide the 

possibility to take advantage of such things as two-photon processes and slow-cooling effects. 

EC

EV

Mini-bands

Figure 1. Idealized energy band diagram of a multiple quantum well structure showing electronic 

"mini-bands". 

To a first order approximion the effective bandgap of a semiconducting quantum dot can be 

calculated using 

2

* 2

2

2

ω π

m r

E

e

g

h h = + (6)

NASA/CP—2005-213431 227
Thus, for a real material the appropriate size to achieve mid-gap state in a suitable device 

structure can be determined. Figure 2. shows a atomic force micscopy image of InAs quantum 

dots grown by metal organic chemical vapor deposition on InGaAs at Essential Research 

Incorporated and imaged in the NanoPower Research labs at Rochester Institute of technology. 

Figure 2. Atomic force micrograph of InAs quantum dots grown on InGaAs. 

Historical Developments 

An antecedent to the use of nanostructures for photovoltaics lies in what has been 

referred to as the "impurity photovoltaic effect (IPV)"[1]. M. Wolf suggested in 1960 that 

impurities could be added in such a way to a conventional p-n junction solar cell that they would 

create available electronic states within the bandgap (see Figure 3). These states could then 

participate in sub-bandgap absorption and act to raise the overall efficiency of a solar cell. 

However Schockley and Queisser, early modelers of single junction photovoltaic, argued against 

the IPV by cited that the recombination losses, which would result from the introduction of the 

impurities would make IPV devices impractical [2]. 

hν1

hν3

Conduction Band

Valence Band 

hν2

Eg

Impurity 

State 

Figure 2. Idealized energy band diagram outlining the impurity photovoltaic effect. 

Photovoltaics designers were able to finally overcome the single junction Shockley -

Quiesser model limit through the use of tandem cells. Even multijunction solar cells have their 

theoretical limitations imposed by such things as current matching of the junctions and the actual 

materials issues imposed by non-idealized or real materials. Theoretically, at least, designers 

could envision solar cell efficiencies that dwarfed the single-junction limit by the use of an everincreasing number of junctions. 

NASA/CP—2005-213431 228
In 1990, a new approach to the possible design of solar cells was put forth in a seminal 

paper by Barnham and Duggen (see Figure 4) [3]. They proposed that the limits imposed on a 

tandem cell could be eliminated through the use of multiple quantum wells (MQW). This work 

was followed by suggestions that the use of these and similar nanostructured approaches could 

yield other benefits such as inverse Auger mechanisms with quantum efficiencies greater than 

one [4-5]. Luque and Marti showed that a quantum dot array incorporated in a p-i-n junction solar 

cell could dramatically exceed the Shockley - Quissier limit and even that of an ideal tandem cell 

(see Figure 5) [6]. The proposed use of quantum dots for spectral shifting and concentrator 

applications has also generated interest [7]. However, the largest impact in the development of 

photovoltaics from quantum dots has been their use as additives in thin film polymeric solar cell 

development [8]. It has been demonstrated that these quantized nanostructured materials will act 

as disassociation centers for the excitons created in the conjuated poymeric absorbers and can 

help facilitate carrier transport through these devices. 

 

Figure 4. Idealized band diagram of a multiple quantum well (MQW) solar cell. Electronic States 

in “wells” are “quantized” and therefore have discrete-like energy levels. The mini-band levels 

depend on the width of the well. K.W.J. Barnham and G. Duggan, J. Appl. Phys., 67, 7, 1990. 

Intermediate 

Band

Eg

hν1

hν2

hν3

Ec

Ev

Figure 5. Idealized energy band diagram of an intermediate band quantum dot solar cell (A. 

Luque and A. Marti, Phys. Rev Lett. 78, 5014 (1997). 

Solar Cell Enhancement with Nanostructures 

The question which normally arises from a discussion of quantum structures and solar 

cells is "can these quantum structures truly be used to enhance solar cell performance? " In 

order to address such a question we must first define what is meant by enhancement. In terms of 

the theoretical analysis of photovoltaic structures, one normally speaks of either a global or 

relative efficiency enhancement. For a global enhancement the question that must be answered 

is "can the efficiency of a nanostructured solar cell under solar illumination be greater than the 

optimum single-gap cell?" These comparisons normally involve many theoretical idealizations 

used in calculating limiting efficiencies. In contrast a relative efficiency enhancement refers to 

NASA/CP—2005-213431 229
whether in principle and practice a nanostructured solar cell can be more efficient under solar 

illumination than corresponding single-gap baseline cells for at least some range of baseline cell 

bandgaps. Finally, there can also be ancillary enhancements, other than just overall devices 

efficiency, whereby in practive a nanostructured solar cell can offer various advantages over 

single-gap cells under useful illumination and operating conditions. Several of these advantages 

that could be especially relevant in terms of space solar power development. For example, it has 

been demonstrated that a multiple quantum well structured can be more radiation resistant than a 

non-nanostructured counterpart [9]. Also, these structures can be used to tailor the thermal 

characteristics of cell to improve cell performance in thermophotovoltaic applications [10]. 

The question of relative efficiency enhancement was recently addressed when it was 

demonstrated that GaAs wells in AlGaAs increased efficiency and photocurrent under broadband 

illumination [11]. However, the observed open circuit voltage drop was inconsistent with original 

hypothesis and this coupled with larger than optimum Eg of AlGaAs left open the possibility that 

the wells merely provided bandgap tuning. The case of global efficiency enhancement is even 

much less clear than that of relative efficiency enhancement and has generated a considerable 

amount of debate. It was argued that the use of a MQW could be used to extend the spectral 

response of a solar cell without sacrificing photovoltage [12]. However, shortly thereafter a 

detailed balance equation calculation was published which showed that the efficiency of a MQW 

cell was the same as that of an ideal single-gap cell [13]. The assumptions used in this 

calculation were immediately challenged and a similar calculation with numerical estimates based 

on real wells actually showed a small global enhancement [14]. However, this result has also 

been criticized because the absorptivity and emissivity of the structure was treated differently 

[15]. There have been other models which use more realistic features and may be better suited 

to real MQW solar cells [16]. 

 The differences in the results of these models could be rectified if it was conclusively 

know whether or not one should assume uniform quasi-Fermi levels (QFLs) across a MQW solar 

cell. If not, then then the criticisms of the detailed balance analyses would no longer be valid. 

Several groups have explored the consequences of non-uniform QFLs, most notably in 

incomplete thermalization [17]. There has been considerable Experimental support for this 

hypothesis [18]. 

 Detailed balance efficiency limits in QWSCs with non-uniform QFLs, which include twophoton processes, has been calculated to be 63% [19]. (Ironically this is the same exact value 

that is obtained by Luque and Marti for their intermediate band solar cell (IBSC) utilized quantum 

dots). In addition, if one allows for reduced carrier transport through hot electron transport, limiting 

efficiencies could exceed even the 63% mark. Luque, Marti, and Quadra have studied various 

multiband models thermodynamically and have concluded that if photon-assisted escape is 

possible then global efficiency improvements are possible [20]. Also, if one neglects hot-electron 

transport, then the thermodynamic model also predicts 63%. However, Luque, Marti, and Quadra 

actually argue against non-uniform QFLs citing it would violate the second law of 

thermodynamics. 

 It is well accepted that the maximum thermodynamic limit for a single junction conversion 

of solar irradiance into electrical free energy in the radiative limit using detailed balance is 31% 

[21]. Therefore, one may view the use of a multi-junction cell approach is actually a means of 

reducing thermalization losses. (The limit of an infinite number of junctions perfectly matched to a 

one-sun solar spectrum being 66%). It can thus be argued that quantum structures can exhibit 

global efficiency improvements by reducing thermal losses. Several approaches to 

accomplishing this in practice have been outlined [22-24]. However, it must be noted in order to 

accomplish this the rates of photogeneration of carriers, transport, and interfacial transfer to the 

contacts must be comparable to the rate of carrier cooling. It has been predicted and 

experimentally shown that relaxation dynamics can be dramatically altered by quantization 

(nanostructures, quantum wells, superlattices, quantum wires, and dots). When carriers are 

confined in potential regions whose dimension is comparable to their deBroglie wavelength or 

Bohr exciton radius in the semiconductor bulk, their relaxation dynamics are dramatically altered 

(i.e., hot carrier extraction rates can be comparable to the rate of carrier cooling). 

NASA/CP—2005-213431 230
Nanostructured Solar Cell Fabrication 

The vast majority of nanostructured photovoltaics have been produced using epitaxial 

growth involving either molecular beam epitaxy (MBE) or metal organic chemical vapor deposition 

(MOCVD), which is sometimes referred to as metal organic vapor phase epitaxy (MOVPE). 

Groups such as the K. Barnham group at Imperial College in London and the A. Zunger group in 

the Space Vacuum Epitaxy Center or the University of Houston have been leaders in the use of 

nanostructure growth for photovoltaics. Recently there has been considerable effort in using 

these similar synthesis techniques to produce photovoltaic structure that incorporates quantum 

dots as opposed to the quantum wells. These attempts use what is referred to as StranskiKrastanow growth. This growth technique utilizes lattice strain to produce “islands” of material on 

a semiconductor surface, and this process can be repeated with alternating complete layers of 

quantum dots to produce self-organized columnar arrangement of dots (see Figure 6). 

Figure 6. Transmission electron micrograph of self-organized quantum dots (Talapin et. al. 

Nanoletters 2001, 1, 207 - 211). 

The most widely investigated nanomaterials for polymeric solar cells have been 

semiconducting nanocrystals, fullerenes, and single wall carbon nanotubes (SWNTs) [28-33]. In 

particular, CdSe quantum dots (QDs), nanorods, and tetrapods have all shown to improve the 

performance of solar cells utilizing conjugated polmers. This work has attracted the attention of 

the space power community for the simple reason that these cells could someday provide specific 

powers that would be unimaginable for their crystalline counterparts. To date the efficiencies are 

still in the single digits and the question of stability in a space environment has yet to be 

addressed. 

Conclusions 

 Ancillary improvements via quantum structures are probably more important to space 

photovoltaics than any other solar cell application. This is due to the potential they hold for better 

temperature coefficients, radiation resistance, and spectral shifting. The bandgap tuning 

available through the use of nanostructures could potentially be integrated into a multi-junction 

solar cell to add another means to improve current matching. Spectrum shifting through up- and 

down-conversion could be combined with almost any cell to improve performance. Quantum dots 

appear to be ideally suited for this application. These nanomaterials and other nanostructures 

should prove quite useful in new concentrator designs. Nanostructured materials may also hold 

the key by which polymeric-based cells reach viable efficiencies and become a contender for 

space solar power applications. 

NASA/CP—2005-213431 231
Acknowledgements 

The authors wish to thank the NASA Glenn Research Center Strategic Research Fund, NASA 

Cooperative Agreements NCC3-710, NCC3-563, National Science Foundation, and First-in-Class 

program of Rochester Institute of Technology. 

References 

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5. J. Werner, et. Al. Phys. Rev. Lett. 72, 3851 (1994). 

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7. K. Barnham, et. al., Appl. Phys. Lett. 76, 1197 (2000). 

8. W.U. Huynh, X. Peng, and A.P. Alivisatos, Adv. Mat. 11, 923 (1999). 

9. R.J. Walters, G.P. Summers, S.R. Messenger, A. Freundlich, C. Monier, F. Newman, 

Prog. In Photovolt.: Res. Appl. 8 (2000) 349; N.A. Sobolev et al. Phys. Stat. Sol B 224 

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10. P. Griffin, I. Ballard, K. Barnham, J. Nelson, A. Zachariou, J. Epler, G. Hill, C. Button, M. 

Pate, Sol. Energy Mater. Sol. Cells, 50 (1998) 213. 

11. K.W.J. Barnham, B. Braun, J. Nelson, M. Paxman, Appl. Phys. Lett. 59 (1991) 135. 

12. K.W.J. Barnham, G. Duggan, J. Appl. Phys. 67 (1990) 3490. 

13. G.L. Araujo and A. Marti, Sol. Energy Mater. Sol. Cells 33 (1994) 213. 

14. R. Corkish and M. Green, Proceedings of the 23rd IEEE Photovolt. Specialists 

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16. N.G. Anderson and S.J. Wojtczuk, J. Appl. Phys. 79 (1996) 1973. 

17. S. Kettemann, J.F. Guillemoles, Proc. of the 13th Europeand Photovolt. Spec. Conf.

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Ekins-Daukes, et. al. Physica E, (1998) 171; K.W.J. Barnham, et. al., J. Mater. Sci.: 

Mater. Electron. 11 (2000) 531; B. Kluftinger, et. al. Sol. Energy Mater. Sol. Cells, 66 

(2001) 501; K Barnham, et. al., Physica E 14 (2002) 27. 

19. S.P. Bremmer, R. Corkish, and C.B. Honsberg, IEEE Trans. Electron. Devices 46 (1999) 

1932. 

20. A. Luque, A. Marti, L. Cuadra, IEEE Trans. Electron. Devices 48 (2001) 2118; A. Luque, 

Physica E, 14 (2002) 107. 

21. W. Shockley and H.J. Queisser, J. Appl. Phys. 32 (1961) 510. 

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Rev. Chem. Phys. 52 (2001) 193. 

23. P.T. Landsberg, H. Nussbaumer, G. Willeke, J. Appl. Phys. 74 (1993) 1451; S. 

Kolodinski, J.H. Werner, T. Wittchen, H.J. Quiesser, Appl. Phys. Lett. 63 (1993) 2405. 

24. A. Luque and A. Marti, Phys. Rev. Lett. 78 (1997) 5014. 

25. F.E. Williams, A.J. Nozik, Nature 311 (1984) 21. 

26. F. Williams, A.J. Nozik Nature 271 (1978) 137. 

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28. V. Dyakonov, Thin Solid Films 451-452 (2004) 493-497. 

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NASA/CP—2005-213431 232
SINGLE WALL CARBON NANOTUBE-POLYMER SOLAR CELLS 

Sheila G. Bailey 

NASA Glenn Research Center, Cleveland, OH 44135, USA 

Stephanie L. Castro 

Ohio Aerospace Institute, Brookpark, OH 44142, USA 

Brian J. Landi, Thomas Gennett, and Ryne P. Raffaelle 

NanoPower Research Laboratories, Rochester Institute of Technology, Rochester, NY 14623, USA 

Investigation of single wall carbon nanotube (SWNT)-polymer solar cells has been conducted 

towards developing alternative lightweight, flexible devices for space power applications. Photovoltaic 

devices were constructed with regioregular poly(3-octylthiophene)-(P3OT) and purified, >95% w/w, 

laser-generated SWNTs. The P3OT composites were deposited on ITO-coated polyethylene 

terapthalate (PET) and I-V characterization was performed under simulated AM0 illumination. 

Fabricated devices for the 1.0% w/w SWNT-P3OT composites showed a photoresponse with an opencircuit voltage (Voc) of 0.98 V and a short-circuit current density (Isc) of 0.12 mA/cm2

. Optimization of 

carrier transport within these novel photovoltaic systems is proposed, specifically development of 

nanostructure-SWNT complexes to enhance exciton dissociation. 

INTRODUCTION

 The lure of large area, inexpensive, and environmentally benign solar cells has attracted many 

researchers over the past four decades or so. Since the early 1990’s, considerable attention has turned 

to lightweight, flexible organic thin-film photovoltaics based on soluble conducting polymers.1

 This new 

class of devices relies upon the interaction between a nanomaterial and a conjugated polymer.2

 In 

addition to enhancing photovoltaic conversion efficiency, the incorporation of nanomaterials can 

potentially improve photochemical, mechanical, and environmental stability. 

 Photon absorption in the organic based composites produces bound-state excitons. Dissociation of 

these charge pairs can be accomplished by the potential difference across a polymer-metal junction 

provided the excitons are near the interface. However, the dissociation can also be accomplished via 

electron accepting impurities.2

 Thus, under illumination a preferential transfer of electrons to the 

acceptors will leave holes to be preferentially transported through the conjugated polymer. This process 

is known as photoinduced charge transfer. Since the discovery of photoinduced charge transfer, a 

variety of acceptor materials have been introduced into donating conjugated polymers to produce 

photovoltaic devices (i.e., Buckminster fullerenes,3,4 CdSe quantum dots and nanorods,5

 and single wall 

carbon nanotubes6

). Typical devices are produced by placing the doped polymeric films between a 

transparent conductive oxide (TCO) top contact and a metallic back contact. 

 The use of SWNTS in this type of device is a particularly attractive approach for several reasons. 

The extremely high surface area, ~1600 m2

/g reported for purified SWNTs,7 offers a tremendous 

opportunity for exciton dissociation. Since SWNTs have diameters of ~1 nm and lengths on the order of 

microns, these materials exhibit significantly large aspect ratios (>103

). At low doping levels, percolation 

pathways are established providing the means for a high carrier mobility and efficient charge transfer. 

This has been a problem in the majority of polymer solar cells developed to date, even with the advent 

of semiconductor nanorods. Since the diffusion distances for excitons in conducting polymers, 

poly(phenylenevinylene) for example, have been reported at <10 nm,4

 the requirement for a sufficient 

percolation network of electron-accepting dopants in the polymer composite is substantiated. Electrical 

conductivity data has validated that SWNT-doped polymer composites demonstrate this extremely low 

NASA/CP—2005-213431 233
percolation threshold. For SWNT-epoxy composites, the electrical conductivity has been claimed to rise 

by nearly 105 S/cm between 0.1 and 0.2% w/w loading.8

 Other beneficial properties of SWNTs relevant to polymeric photovoltaic development include 

composite reinforcement and thermal management. Single wall carbon nanotubes have shown promise 

in the development of polymer composites with enhanced mechanical strength by load transfer from the 

polymer matrix to the dopant.9

 Tensile strengths for SWNTs have been estimated to equal ~20 GPa,10

while the Young’s modulus measured by atomic force microscopy is ~1 TPa.11 This high Young’s 

modulus and strength-to-weight ratio could help provide much needed mechanical stability to large area 

thin-film arrays. Single wall carbon nanotubes may also provide assistance in thermal management for 

such arrays. The thermal conductivity of an isolated (10, 10) SWNT has been theoretically predicted to 

be as high as 6600 W/mK.12 Polymer composites doped with as little as 1% w/w SWNTs have shown a 

70% increase in the thermal conductivity at 40 K.8

 

 The viability of incorporating SWNTs into a conducting polymer for photovoltaic devices was 

established in 2002, utilizing arc-generated SWNT- poly(3-octylthiophene)-(P3OT) composites.6

 Their 

results showed a diode response for devices constructed in the sandwich formation, containing the 

composite film between an indium-tin-oxide (ITO) front contact and aluminum back contact. There was 

a photoresponse under AM1.5 illumination for both the pristine P3OT device and a 1% SWNT-P3OT 

composite blend, albeit the composite exhibits current densities several orders of magnitude higher. 

The photoresponse for a 1% w/w doped SWNT-P3OT composite was reported to have an open-circuit 

voltage (Voc) of 0.75 V and a short-circuit current density (Isc) of 0.12 mA/cm2

, in comparison to a Voc= 

0.35 V and Isc= 0.7 μA/cm2

, for the pristine P3OT device.6 

 Originally, the SWNT-P3OT cells were described using the metal-insulator-metal (MIM) model for 

determining the Voc based on a work function differential between the two metal electrodes. In the 

reported case, the maximum Voc would be ~0.4 V based on values of 4.3 eV for Al and 4.7 eV for ITO,6

although work has been done to show some variability in the work function of ITO dependent on 

processing and oxygen content.13 Recent work has shown that investigation of differing work function 

metal electrodes that the MIM model is not applicable for SWNT-P3OT systems, due to a consistent Voc

of 0.75V. Rather, they propose that internal SWNT/polymer junctions based on the highest occupied 

molecular orbital (HOMO)-lowest unoccupied molecular orbital (LUMO) of the respective materials leads 

to enhanced charge separation and collection, providing the higher observed Voc value.14 In principle, 

control over these junctions may be garnered by reducing defect sites, altering the diameter 

distributions (which is inversely related to the bandgap in semiconducting types), or varying the 

concentration ratio of metallic to semiconducting for the SWNTs prior to dispersion. Such processing 

control offered by alternative synthesis conditions and purification strategies may further improve the Voc

and in turn produce more efficient SWNT-polymer solar cells. 

EXPERIMENTAL

 Single wall carbon nanotubes used for this study were synthesized using the pulse laser 

vaporization technique, employing an Alexandrite laser (755 nm) which pulsed the surface of a 0.6 at.% 

Ni/Co doped graphite target at a power density of 20 - 150 W/cm2

. The reaction furnace temperature 

was held at 1150°C, while the chamber was at a pressure of 400 torr with 100 sccm flowing Ar(g).

15 The 

as-produced SWNTs were collected from the condensed region on the quartz tube outside the furnace 

and purified by modification of the previously reported procedure.16 In short, ~50 mg of as-produced 

SWNTs were refluxed in 3M nitric acid for 16 hours, and then filtered over a 0.2 μm PTFE membrane 

filter with copious amounts of de-ionized water. After drying the membrane filter at 70°C in vacuo to 

remove the SWNT paper, thermal oxidation proceeded in air at 500°C for one hour in a muffle furnace. 

Use of a 6M Hydrochloric acid wash for one hour in a water bath sonicator was at times necessary to 

remove any remaining metal catalyst impurities, with similar filtering steps. Finally, thermal oxidation at 

550°C in air for one hour completed the purification. Raman spectroscopy, surface area analysis, 

scanning electron microscopy (SEM), optical absorption spectroscopy, and thermal gravimetric analysis 

(TGA) were utilized to fully characterize all raw soot and purified materials. 

 Preparation of the SWNT-P3OT composite solutions was performed using a series of mixing and 

sonication steps as previously demonstrated for other SWNT-polymer systems.17 The necessary 

amount of regioregular P3OT, supplied by Aldrich, was dissolved in chloroform using water bath 

sonication to achieve the pristine 15 mg/mL solution. Composite dispersion was performed by 

NASA/CP—2005-213431 234
combining the appropriate mass of purified SWNTs to the pristine solution at desired doping levels. The 

composite solution was then placed in a water bath sonicator for five minutes prior to high speed stirring 

at room temperature for 72 hours. Aliquots of this dispersed solution were cast on Teflon substrates at 

room temperature to produce desired thin films required for optical absorption spectroscopy. 

 Device fabrication involved the use of commercially-obtained, high quality (i.e., < 10 Ω/sq.) ITOcoated polyethylene terapthalate (PET) substrates. The required cleaning of the ITO substrate prior to 

deposition was performed by sonication and rinsing of cut samples (1 inch x 1.5 inches) in de-ionized 

water, acetone, and methanol. Initially, an intrinsic layer of pristine P3OT is spray deposited (~1-2 mL 

of the 15 mg/mL solution) onto the masked, 1 in.2

 active area of the substrate. This is followed with 

spray deposition of the SWNT-P3OT composite solutions at similar volumes. As seen in Figure 1, 

completion of the solar cell occurs when aluminum contacts (typical thicknesses of 1000Å) are applied 

to the ITO and SWNT-P3OT composite film layers. This is accomplished using thermal evaporation 

under vacuum at pressures <10-6 mbar with a standard shadow mask. 

 Solar cell testing was performed in an isolated black box configuration to assure standard 

calibration and reproducibility of results. Thin film SWNT-P3OT cells were placed on a temperaturecontrolled stage, held at 20oC, during analysis. The I-V characteristics were monitored using a Keithley 

237 source measurement unit. Simulated AM0 illumination was achieved by calibration of ELH lamps 

with a NASA certified Si standard cell. 

Al

Polyethylene Terapthalate(PET)

Indium Tin Oxide (ITO)

Al Poly-3-Octyl-thiophene (P3OT)

SWNT-doped P3OT

Al Al Grid Contact

(a) (b)

Al

Polyethylene Terapthalate(PET)

Indium Tin Oxide (ITO)

Al Poly-3-Octyl-thiophene (P3OT)

SWNT-doped P3OT

Al Al Grid Contact

(a) (b)

Figure 1. (a) Schematic representation and (b) image depicting the composition of fabricated SWNTP3OT flexible solar cells 

RESULTS

 The ability to elicit control over as-produced SWNTs is a major advantage for using pulse laser 

vaporization synthesis techniques. Such parameters as diameter distribution, defect density, and 

metallic to semiconducting ratios can be altered based on the synthesis conditions.15,16,18 The asproduced SWNTs used in this study were synthesized and purified under typical conditions, where 

substantial yields and minimal defects have been observed. Shown in Figure 2 is a representative (a) 

SEM image of purified, >95% w/w SWNTs, and (b) an overlay of the Raman spectra for the radial 

breathing mode (RBM) of this sample for incident laser energies of 1.96 and 2.55 eV. The SEM image 

shows an entangled “mat” of SWNTs with no significant metal catalyst impurities or amorphous carbon 

coatings. The high purity SWNTs will “bundle” due to the Van der Waals interaction between the 

individual tubes. Control over the influence of bundling can potentially be addressed through polymeric 

interaction. Previous work has suggested a “debundling” effect with SWNT-Nafion composites, 

observed by SEM, that was proposed to reduce the necessary doping level for a percolation threshold.17 

Dispersions of SWNTs in P3OT may follow similar results, although full evaluation of this has not yet 

been performed. In addition, Raman spectroscopy can be used to determine the diameter distribution, 

chirality, and defect density of the as-produced and purified SWNTs. The RBM is routinely used to 

NASA/CP—2005-213431 235
calculate the diameter distribution.19 Using two incident laser energies during analysis allows for the 

determination of SWNT type, i.e. semiconducting vs. metallic.20,21 For this diameter range, the 1.96 eV 

laser resonantly enhances predominantly metallic SWNTs which correspond to ~1.2-1.4 nm in 

diameter.19 Similarly, the 2.55 eV laser probes the semiconducting type, also shown to contain 

equivalent diameters for the purified sample. 

The homogeneous distribution of SWNTs in a polymer matrix is dependent upon the ability of the 

polymer chain to associate with the SWNT superstructure. Stable composite dispersions of 0.1% and 

1.0% w/w SWNTs in P3OT were produced and analyzed in this study. Optical spectroscopy was 

conducted to observe the SWNT doping level effects on the absorption properties and infer potential 

electronic interactions between dopant and polymer. 

100 150 200 250 300

Raman Intensity

Raman Shift (cm-1)

1.96 eV

2.55 eV 161

181

164

179

(a) (b)

100 150 200 250 300

Raman Intensity

Raman Shift (cm-1)

1.96 eV

2.55 eV 161

181

164

179

(a) (b)

Figure 2. (a) SEM image of purified, >95% w/w SWNTs, and (b) overlay of Raman spectrum for the 

RBM of (a) at incident laser energies of 1.96 and 2.55 eV. The labeled peaks correspond to a diameter 

distribution between 1.2-1.4 nm for both metallic and semiconducting SWNTs. 

Shown in Figure 3 is an overlay of these spectra where the pristine P3OT shows strong optical 

absorption at energies >2 eV. As expected, variation in absorption properties for the SWNT-P3OT 

composites is observed as the doping level increases. Through modification of the SWNT doping level 

in the polymer, it is possible to alter the absorption pattern of these composite materials. In fact, at 

these relatively low doping levels, the P3OT shows a significant enhancement in absorption through the 

near-IR and visible regions. The gray curve for purified SWNTs is offset from the other three for clarity, 

but indicates the typical complex pattern observed for this spectral range. The two lowest energy peaks 

for SWNTs (~0.7 eV and ~1.3 eV) are representative of the energy transitions corresponding to the 1st

and 2nd Van Hove (VH) singularities for semiconducting types, respectively.22 The SWNT-P3OT 

composites display a relative quenching of the 1st VH compared to the 2nd VH, which is typical of doping 

effects on the SWNTs.23,24 Therefore, an apparent doping of the semiconducting SWNTs by the 

polymer is present. However, due to the higher intensity at larger energies, typical of absorption for the 

metallic SWNTs (~1.8 eV), any interaction between these SWNT types and the polymer is presently 

unavailable. 

Application of SWNT-P3OT composite solutions to the constructed devices was performed using a 

solution-spray technique onto the ITO/PET substrates. The unique approach to deposit these 

photoactive composites on polymer substrates demonstrates the inherent advantages of thin film 

polymeric solar cells, namely low structural weight and flexibility. These novel polymeric solar cells 

were constructed for pristine P3OT, 0.1% w/w, and 1.0% w/w SWNTs in P3OT, and tested under 

simulated AM0 illumination to determine the respective I-V characteristics. The typical photoresponse 

observed for pristine P3OT and 1% w/w SWNT-P3OT cells is shown in Figure 4. Although each overlay 

NASA/CP—2005-213431 236
shows significant enhancement in both Isc and Voc upon illumination, there is an absense of the typical 

diode “knee” for both cases. The cause of the larger than expected reverse bias currents is presently 

not understood. This effect may be attributed to long-lived energy states for the SWNTs and P3OT or 

the presense of carrier traps which reduces the Isc. 

0

0.1

0.2

0.3

0.4

0.5

0.8 1.2 1.6 2

A

b

s

o

r

b

a

n

c

e

a(

u.

.

)

Wavelength (eV)

Purified SWNTs

1.0% w/w SWNT-P3OT

0.1% w/w SWNT-P3OT

Pristine P3OT

Figure 3. Optical absorption overlay for thin films of pristine P3OT (black), 0.1% w/w SWNT-P3OT 

composite (red), 1.0% w/w SWNT-P3OT composite (blue), and purified SWNTs (gray). Enhancement in 

absorption for the composites relative to the pristine P3OT is clearly observed 

-0.03

-0.02

-0.01

0

0.01

0.02

-0.8 -0.4 0 0.4 0.8

Current Density (μA/cm2

)

Voltage (V)

Dark

AM0

-0.3

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

-0.8 -0.4 0 0.4 0.8

Current Density (mA/cm2

)

Voltage (V)

Dark

AM0

(a) (b)

-0.03

-0.02

-0.01

0

0.01

0.02

-0.8 -0.4 0 0.4 0.8

Current Density (μA/cm2

)

Voltage (V)

Dark

AM0

-0.3

-0.25

-0.2

-0.15

-0.1

-0.05

0

0.05

0.1

-0.8 -0.4 0 0.4 0.8

Current Density (mA/cm2

)

Voltage (V)

Dark

AM0

(a) (b)

Figure 4. Characteristic I-V plots in the dark (gray) and under simulated AM0 illumination (black), 

displaying the photoresponse for (a) pristine P3OT and (b) 1% w/w SWNT-P3OT composite solar cells 

Further characterization of the composite cells showed a corresponding increase in the Isc as the 

doping level of SWNTs increased in the P3OT. Figure 5 shows an I-V overlay under the simulated AM0 

illumination for the pristine P3OT, 0.1%, and 1.0 % w/w SWNT-P3OT cells in the region of forward bias. 

Clearly evident is an order of magnitude increase in the Isc from the pristine P3OT to the 0.1% w/w 

SWNT-P3OT composite cell. An additional increase of ~50% is observed when the doping level 

reaches 1.0% w/w, with an Isc equal to 0.12 mA/cm2

 for the cell. This result indicates that there is an 

apparent conductivity effect on the Isc in the composite films which can be controlled by SWNT doping. 

Optimization of the doping level for maximum current density is a necessary step at improving the 

overall efficiency of these cells. Similarly, the Voc more than doubled for both SWNT-doped cells, 

NASA/CP—2005-213431 237
although the 0.1% w/w composite has a slightly higher value than the 1.0 % w/w composite. 

Interestingly, the measured Voc of 0.98 V is signficantly higher than the recent report, where they 

postulated that the energy difference in the HOMO-LUMO levels of the SWNT/polymer junction is 

responsible for the open-circuit-voltage.14 An explanation of this current result may reside in the purity 

and defect density of the SWNTs. This is an important variable to consider since the interaction 

between SWNTs and the polymer can directly affect the junction by which exciton dissociation occurs. 

Other variables to evaluate for SWNT-polymer solar cells are the diameter distributions and metallic to 

semiconducting ratios of the dopant. If the SWNTs are primarily acting as a conductive network, then 

the presence of metallic types would be the preferential dopant. However, the possibility exists that 

semiconducting SWNTs are participating in the photoconversion process since a recent report has 

shown certain photoconductivity effects exist for SWNTs.25 In either case, selective dispersion of one 

type could lead to higher SWNT-P3OT solar cell efficiencies. Above all though, the ability to dissociate 

and extract the generated carriers upon illumination in these conducting polymer solar cells is the 

foremost concern to augment the overall efficiency. 

-0.12

-0.1

-0.08

-0.06

-0.04

-0.02

0

0.02

0 0.2 0.4 0.6 0.8 1 1.2

Current Density (mA/cm2

)

Voltage (V)

P3OT Cell

0.1% w/w SWNT-P3OT Cell

1.0% w/w SWNT-P3OT Cell

Figure 5. Overlay of the I-V relationship under simulated AM0 illumination for pristine P3OT (gray), 

0.1% w/w SWNT-P3OT composite (black dashed), and 1.0% w/w SWNT-P3OT composite (black line) 

solar cells. 

 A next-generation approach to maximizing the charge separation and carrier transport in SWNTdoped P3OT cells may be through attachment of nanostructured components to the SWNTs (see Figure 

6a). Such versatility enables tuning of the optoelectronic properties of the dopant in the polymer cell to 

match the energy bands between components and conducting polymer HOMO/LUMO levels. Our 

approach to improve device efficiency is through covalent attachment of semiconducting quantum dots 

to SWNTs. It is expected that conjugation of SWNTs and quantum dots prior to dispersion will facilitate 

charge dissociation and transport in the polymer solar cells. Our initial attempt at establishing the 

necessary chemistry has been successful with attachment of CdSe quantum dots to functionalized 

SWNTs through an ethylenediamine (en) linker, similar to the reported procedures.26,27 Shown in Figure 

6b is a tapping-mode atomic force micrograph (AFM) of our resulting CdSe-en-SWNT complexed 

sample. Incorporation of this material into similar P3OT composite solar cells as described here, is 

currently under investigation. 

NASA/CP—2005-213431 238
Conjugated

Polymer 

Functionalized 

quantum dot

SWNT 

backbone

Conjugated

Polymer 

Functionalized 

quantum dot

SWNT 

backbone

(a) (b)

Conjugated

Polymer 

Functionalized 

quantum dot

SWNT 

backbone

Conjugated

Polymer 

Functionalized 

quantum dot

SWNT 

backbone

Conjugated

Polymer 

Functionalized 

quantum dot

SWNT 

backbone

Conjugated

Polymer 

Functionalized 

quantum dot

SWNT 

backbone

(a) (b)

Figure 6. (a) Schematic representation of the approach to conjugate quantum dots with SWNTs prior to 

dispersion in the conducting polymer; (b) AFM image of CdSe-en-SWNTs to be used as dopant in future 

SWNT-P3OT solar cells. 

CONCLUSIONS

 The ability to construct flexible solar cells containing SWNT doped polymer composites has been 

demonstrated. Devices containing 1% w/w SWNTs in P3OT have shown a photoresponse with low 

current densities but relatively high open-circuit voltages (~1 V). Attempts to improve the exciton 

dissociation and carrier transport in these systems may rely upon controlled processing of SWNTs for 

appropriate material properties. Ultimately, the ability to covalently attach other nanomaterials to the 

SWNTs could result in higher polymer solar cell efficiencies. 

ACKNOWLEDGEMENTS

This work was supported by NASA Glenn Research Center through grant NAG3-2595 and the National 

Science Foundation through grant ECS 0233776. The authors wish to thank Stephen Fahey for 

obtaining the AFM images. 

REFERENCES

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(2) Granstrom, M.; Petritsch, K.; Arias, A. C.; Lux, A.; Andersson, M. R.; Friend, R. H. Nature 1998, 

395, 257-360. 

(3) Camaioni, N.; Ridolfi, G.; Casalbore-Miceli, G.; Possamai, G.; Garlaschelli, L.; Maggini, M. Solar 

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3120-3122. 

(5) Huynh, W. U.; Dittmer, J. J.; Alivisatos, A. P. Science 2002, 295, 2425-2427. 

(6) Kymakis, E.; Amaratunga, G. A. J. Appl. Phys. Lett. 2002, 80, 112-114. 

(7) Cinke, M.; Li, J.; Chen, B.; Cassell, A.; Delzeit, L.; Han, J.; Meyyappan, M. Chem. Phys. Lett.

2002, 365, 69. 

(8) Biercuk, M. J., Llaguno, M.C., Radosavljevic, M., Hyun, J.K., Johnson, A.T., Fischer, J.E. Appl. 

Phys. Lett. 2002, 80, 2767-2769. 

(9) Schadler, L. S.; Giannaris, S. C.; Ajayan, P. M. Appl. Phys. Lett. 1998, 73, 3842. 

(10) Li, F., Cheng, H.M.; Bai, S.; Su, G.; Dresselhaus, M.S. Appl. Phys. Lett. 2000, 77, 3161-3163. 

(11) Salvetat, J.; Briggs, G. A. D.; Bonard, J.; Bacsa, R. R.; Kulik, A. J.; Stockli, T.; Burnham, N. A.; 

Forro, L. Phys. Rev. Lett. 1999, 82, 944. 

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(12) Berber, S.; Kwon, Y.; Tomanek, D. Phys. Rev. Lett. 2000, 84, 4613. 

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(16) Dillon, A. C.; Gennett, T.; Jones, K. M.; Alleman, J. L.; Parilla, P. A.; Heben, M. J. Adv. Mater.

1999, 11, 1354. 

(17) Landi, B. J.; Raffaelle, R. P.; Heben, M. J.; Alleman, J. L.; VanDerveer, W.; Gennett, T. Nano 

Lett. 2002, 2, 1329. 

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(19) Rao, A. M.; Chen, J.; Richter, E.; Schlecht, U.; Eklund, P. C.; Haddon, R. C.; Venkateswaran, U. 

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NASA/CP—2005-213431 240
High-Temperature Solar Cell Development

Geoffrey A. Landis,

NASA John Glenn Research Center

21000 Brookpark Road, Cleveland, OH 44135

 Danielle Merritt and Ryne P. Raffaelle

Rochester Institute of Technology

Rochester, NY

David Scheiman

Ohio Aerospace Institute

Brook Park, OH 44142

Missions for High Temperature Solar Cells

The vast majority of space probes to date have relied upon photovoltaic power

generation. If future missions designed to probe environments close to the sun (Figure 1)

will be able to use such power generation, solar cells that can function at high temperatures,

under high light intensity, and high radiation conditions must be developed. The

significant problem is that solar cells lose performance at high temperatures.

Approaches to solar arrays for near-sun missions include:

• High epsilon/low alpha coatings

• Array off-pointing: array points at angle to sun

• Partially-populated array (with missing cells replaced with mirrors)

• Reflective coatings

• Solar cells designed to operate at high temperature

Active cooling techniques, such as use of solid-state refrigerators, in general require

more power to operate than the resultant gain in efficiency.

 For example, current plans for Mercury missions reflect away most of the incident

solar energy to limit the operating temperature and avoid destroying the array (figure 2).

While this is one solution to the problem, this does not optimally use the solar energy, and

it would be desirable to develop solar cells that can perform well at high temperatures,

rather than developing techniques to reduce the temperature at the price of reduced

performance.

NASA/CP—2005-213431 241
Figure 1: Artist's concept of the Solar Probe mission, proposed to reach a distance of 4

solar radii from the sun.

Figure 2: Solar array for the MESSENGER mission to Mercury (Johns Hopkins

Applied Physics Laboratories). The majority of the sun-facing surface consists of mirrors

to reflect the incident solar radiation and limit temperature.

NASA/CP—2005-213431 242
NASA requirements for solar power systems for high temperature near-sun operation

has the goals [1]:

Improved efficiency at high temperature

Improved lifetime at high temperature

Solar cells are desirable for elevated operating temperatures ranging from the Mercury

orbiter, in which the solar array equilibrium temperature (assuming a sun-facing array with

no reflectors) would be 450 C, to the proposed Solar Probe mission [2], for which, at 4

solar-radii, a sun-facing array with no reflectors would reach an equilibrium temperature of

approximately 2300 C. Solar cells made from wide bandgap compound semiconductors

are an obvious choice for such an application, since the higher voltage of wide bandgap

solar cells results in less degradation [3,4]. For example, silicon solar cells (1.1 eV) lose

about 0.45% of their power per degree C increase in operating temperature. GaAs cells

(1.4 eV) lose about 0.21% per degree C [5].

The normalized temperature coefficient (1/ d/dT) 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:

1/ d/dT = 1/Voc dVoc/dT + 1/Jsc dJsc/dT + 1/FF dFF/dT (2)

The Voc variation contributes the majority of the change in efficiency. Fill factor

variation in general tracks the open circuit variation. 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 to longer wavelength portions

of the spectrum, and therefore the short circuit current actually increases with temperature.

Hence, the Jsc variation term is roughly proportional to the incident spectral intensity at

wavelengths near the band edge [5].

Since the Voc variation with temperature is roughly the same for cells of different

bandgap, while the actual Voc increases with bandgap, the normalized temperature

coefficient, 1/ d/dT increases directly with bandgap. However, since the photon flux

from the sun decreases at high photon energies, an optimum bandgap exists for each

temperature.

To verify the efficiency of wide bandgap solar cells at high temperatures, we measured

a GaInP solar cell (1.6) as a function of temperature from room temperature up to 400 C.

As shown in figure 3, open circuit voltage and fill factor decrease with temperature, while

the short circuit current shows a slight increase. Power loss [1/P dP/dT] is about 0.177%

per degree, with irreversible degradation due to shunting occurring slightly above 350C.

The theoretical performance of solar cells as a function of bandgap and temperature is

shown in figure 4. As can be seen, the optimum bandgap shifts from about 1.4 volts at

room temperature (27C) to about 2.3 volts at 900C.

NASA/CP—2005-213431 243
400

600

800

1000

1200

1400

0 50 100 150 200 250 300 350 400

InGaP High Temperature

Voc (mV)

Temp (°C)

50

55

60

65

70

75

80

0.0 50.0 100.0 150.0 200.0 250.0 300.0 350.0 400.0

InGaP High Temperature

Fill Factor

Temp (°C)

15

20

25

30

35

40

45

50

55

0.0 50.0 100.0 150.0 200.0 250.0 300.0 350.0 400.0

InGaP High Temperature

Pmax (mW)

Temp (°C)

Figure 3: Open circuit current, fill factor, and efficiency of a GaInP solar cell measured as a

function of temperature, from 0 to 400 C.

NASA/CP—2005-213431 244
AM0 

FF = 0.80, Pin = 0.13661 (W/cm^2)

0

5

10

15

20

25

0.5 1 1.5 2 2.5 3 3.5

Bandgap (eV)

Efficiency (%)

T = 300K

T = 573K

T = 673K

T = 773K

T = 873K

T = 973K

T = 1073K

T = 1173K

Figure 4: Theoretical efficiency of a solar cell as a function of bandgap, showing the shift

of optimum bandgap from about 1.4 volts at room temperature (27C) to about 2.3 volts at

900C.

In order develop solar cells for such applications, we have initiated a program to

manufacture and measure the photovoltaic performance of wide bandgap solar cells,

including cells from GaInP, GaP [6], GaN, and SiC [7]. Figure 5 shows a SiC solar cell

developed at NASA Glenn in collaboration with the Rochester Institute of Technology and

Cree Semiconductors [8].

Figure 5: SiC solar cell. Left: schematic cross section. Right: photo of cell.

NASA/CP—2005-213431 245
In addition to the reversible loss of theoretical performance with temperature, solar

array operation at high temperature needs to avoid irreversible degradation leading to

destruction of the arrays. Effects that produce irreversible performance loss include:

• Ohmic contact degradation [9, 10]

• Dopant diffusion

• Compound semiconductor degradation

• Interconnect-related failure

• Coverglass debonding [11]

• Array structural degradation.

Technologies to deal with these problems have been developed under other programs

[9, 10, 11].

Conclusions

High-temperature operation of solar cells is of interest to future NASA missions.

Technology solutions such as off-pointing can reduce operating temperature, but also

reduce power from the array. New solar cells that can operate at high temperature are

desirable; this requires development of high bandgap semiconductors. A program to

develop cells for high temperature operation, including GaInP, GaN, SiC and GaP cells, is

in progress.

Achieving satisfactory operating lifetime at high temperature is an issue that has not yet

been addressed in detail.

References

1. G. Landis and S. Bailey, "Photovoltaic Power for Future NASA Missions," paper

AIAA-2002-0718, AIAA 40th Aerospace Sciences Meeting, Reno NV, January 14-17

2002.

2. S. Kerridge, et al. "Cost Effective Mission Design for a Small Solar Probe," Acta

Astronautica, Vol. 35, Suppl., pp. 257-266 (1995).

3. D. Scheiman, G. Landis, and V. Weizer, "High-Bandgap Solar Cells for Near-Sun

Missions," presented at Space Technology & Applications International Forum

(STAIF-99), Conf. on Global Virtual Presence, Albuquerque NM, 30 Jan.- 4 Feb. 1999;

AIP Conference Proc. 458, 616-620.

4. J. Fan, "Theoretical Temperature Dependence of Solar Cell Parameters," Solar Cells

17, pp. 309-315 (1986).

5. G. Landis, "Review of Solar Cell Temperature Coefficients for Space," Proc. XIII

Space Photovoltaic Research and Technology Conference, NASA CP-3278, NASA

Lewis Research Center, June 1994, 385-400.

NASA/CP—2005-213431 246
6. O. Sulima et al., “High Temperature AlGaP/GaP Solar Cells for NASA Space

Missions,” presented at the 3rd World Conference on Photovoltaic Energy Conversion,

Osaka Japan, May 11-18 2003.

7. R. Raffaelle, S. Bailey, P. Neudeck, R. Okojie, C. Schanabel, M. Tabib-Azar, D.

Scheiman, P. Jenkins, and S. Hubbard, "Optical and Electrical Characterization of SiC

Devices, Proc. 28th IEEE Photovoltaic Specialists Conf., pp. 1257-1260 (2000).

8. S. Bailey and R. Rafaelle, "Silicon Carbide Solar Cells Investigated," Research &

Technology 2001, NASA Technical Memorandum (2002),

9. S. P. Tobin et al. “Advanced Metallization for Highly Efficient Solar Cells,” Proc. 19th

IEEE Photovoltaic Specialists Conf., pp. 70-75 (1987).

10. M. B Spitzer., et al. "Gallium Arsenide Concentrator Solar Cells with Highly Stable

Metallization,” Proc. 20th IEEE Photovoltaic Specialists Conference, pp. 930-933

(1988).

11. P. Younger, G. Landis and R.G. Tobin, "Integral Glass Covering of Spacecraft Solar

Cells by Electrostatic Bonding," Proceedings of the 14th IEEE Photovoltaic Specialists

Conference, San Diego, CA; 1240-1245 (1980).

NASA/CP—2005-213431 247
Integrated Phase Array Antenna/Solar Cell System For

Flexible Access Communications (IA/SAC)

E.B. Clark, R.Q. Lee, A.T. Pal, D.M. Wilt, and B.D. McElroy

NASA Glenn Research Center

Cleveland, OH 44135

C.H. Mueller

Analex Corporation

Cleveland, OH 44135

Abstract

This paper describes recent efforts to integrate advanced solar cells with printed planar

antennas. Several previous attempts have been reported in the literature, but this effort is

unique in several ways. It uses Gallium Arsenide (GaAs) multi-junction solar cell

technology. The solar cells and antennas will be integrated onto a common GaAs

substrate. When fully implemented, IA/SAC will be capable of dynamic beam steering.

In addition, this program targets the X-band (8 – 12 GHz) and higher frequencies, as

compared to the 2.2 – 2.9 GHz arrays targeted by other organizations. These higher

operating frequencies enable a greater bandwidth and thus higher data transfer rates. The

first phase of the effort involves the development of 2 x 2 cm GaAs Monolithically

Integrated Modules (MIM) with integrated patch antennas on the opposite side of the

substrate. Subsequent work will involve the design and development of devices having

the GaAs MIMs and the antennas on the same side of the substrate. Results from the

phase one efforts will be presented.

Introduction

Recently, several attempts to integrate

printed planar antennas onto solar cells

have been reported in the literature. The

first work reported in 1995 involved

attaching solar cells to the surface

of a single micro-strip antenna for a

micro-satellite application [1].

Measured data from this study showed

very good antenna performance with the

solar cells having very little impact on

the radiation pattern and return loss at

S-band. In 1999, JPL attempted to

combine antenna and solar cells for a

Mars rover application at 400 MHz [2].

These early works attempted to

demonstrate feasibility using

commercial solar cells and a single

antenna. Integrating multiple antenna

elements and solar cells on the same

surface, was first reported by a group in

Europe [3-4]. Two small antenna arrays

integrated with solar cells were

demonstrated. The first was an 8-

NASA/CP—2005-213431 248
element (4x2) array with 9 solar cells

strips connected in series and

superimposed on the antenna elements.

The second was a 4x4 array of circularly

polarized slot dipoles co-located on the

same surface of the solar cells array.

Amorphous silicon solar cells with about

5-12% efficiency were used in both

demonstrations.

Compared to prior work, the GRC

proposed integrated phased array/solar

array cell (IA/SAC) concept is

technologically more advanced, and is

unique in several ways. As stated

previously, the earlier work used silicon

cells with a maximum efficiency of

about 15%. This effort uses GaAs based

monolithically integrated module (MIM)

technology, which has a greater

efficiency and can be easily tailored to

produce a desired voltage. In addition,

IA/SAC modules will be integrated on a

common substrate as opposed to the

previous efforts that developed the

components on separate surfaces then

combined them post processing. Finally,

when fully implemented, IA/SAC arrays

will be capable of dynamic beam

steering in the X-band (8-12 GHz) or

higher frequencies. To the best of our

knowledge, in all prior developmental

efforts, the antennas are passive arrays

capable of only fixed point-to-point

communications operating in the 2.2 –

2.9 GHz range.

Experimental Procedure

In the first phase of the project, we

designed and developed single junction

2x2cm GaAs MIM devices. MIM

devices are grown on semi-insulating

substrates and have both the positive and

negative contacts on the top-side of the

device (figure 1). Elimination of the

contacts on the back surface allows the

semi-insulating substrate to act as the

dielectric layer between the ground

plane and radiating elements of the

antenna.

Figure 1. Diagram of a GaAs MIM

device showing the device layers and

interconnect.

The use of MIM technology was a

crucial step towards the ability to

integrate the solar cell and antenna on a

common substrate. The Photovoltaic and

Space Environments (PVSE) Branch of

NASA Glenn originally developed the

MIM concept for thermophotovoltaic

(TPV) energy conversion. This device

has been the focus of a multi-million

dollar, multi-year development by

Bechtel Bettis, Inc [5, 6, 7]. One of the

main goals in the first phase of this effort

was to increase the size of the GaAs

MIMs to 2cm x 2cm. This is

significantly larger than previous devices

and it is anticipated that these MIM

devices will have a conversion

efficiency comparable to GRC produced

GaAs solar cells (~20% efficiency). The

initial devices will have a structure

similar to figure 2 but will be optimized

for the IA/SAC project.

InGaP Window

GaAs Emitter

GaAs Base

InGaP Window

GaAs Lateral

Conductor

Semi-Insulating

GaAs

Interconnect

Dielectric

NASA/CP—2005-213431 249
Lateral Conduction Layer

0.05 μm p GaAs contact

0.1 μm p InGaP window

0.5 μm p GaAs emitter

2 μm n GaAs base

0.05 μm n InGaP window

2 μm n GaAs

 350 μm GaAs

Semi Insulating Substrate

Ground Plane / BSR

Figure 2. GaAs single junction IA/SAC

MIM device structure.

The GaAs MIMs were grown on 2 inch,

double side polished GaAs wafers by

Organometalic Vapor Phase Epitaxy

(OMVPE) at 6200

c and 190 torr. The

specular films were then processed into

1cm x 1cm MIMs. After AM0 currentvoltage (IV) measurements were taken,

patch antennas were deposited on the

opposite side of the wafer by E-beam

evaporation. After deposition of the

antenna, AMO IV measurements were

repeated to determine whether there was

any degradation to the solar cells.

Results and Analysis

Figure 3. Pre- and Post antenna

deposition AM0 IV measurements.

The IV measurements in figure 3 show

that there was no degradation in the

performance of the solar cells due to the

deposition of the antenna. In fact, an

increase in the cells performance is

noted. This is most likely due to

additional contact annealing that occurs

during the deposition of the antenna. At

the time of this writing, the results of the

testing performed in the Microwave

Antenna Range at NASA GRC were

unavailable.

Figure 4. IA/SAC module being

mounted for testing in antenna range.

Conclusion

We have demonstrated the feasibility of

combining advanced solar cells with RF

antenna technology on a common

substrate with no degradation to the solar

cells. Future phase one work will focus

on characterizing the antenna and

increasing the size of the MIM to 2cm x

2cm.

References

1. M. Tanaka, Y. Suzuki, K. Araki and

R. Suzuki, “Microstrip antennas with

solar cells for microsatellites,”

Electronics Letters, Vol. 31, 1996, pp.

263-266.

I-V Curve for NAS847-5 before and after

antenna deposition

0.001

0.0015

0.002

0.0025

0.003

0.0035

01234567

Voltage, V

Current, mA

Before

After

NASA/CP—2005-213431 250
2. J. Huang, “Mars rover antenna for

solar-array integration,’ JPL TMO

Progress Report 42-136, Feb. 15, 1999.

3. S. Vaccaro, P. Torres, J.R. Mosig, A.

Shah, J.F. Zurcher, A.K. Skrivervik, F.

Gardiol, P. de Maagt and L. Gerlach,

“Integrated solar panel antennas’”

Electronics Letters, vol. 36, No. 5, 2000.

4. S. Vaccaro, P. Torres, J.R. Mosig, A.

Shah, A.K. Skrivervil, J.F. Zurcher, P.

de Maagt, and L. Gerlach, “Combination

of antennas and solar cells for satellite

communications,” Microwave and

Optical Technology Letters, Vol. 29, No.

1, April 2001, pp. 11-16.

5. D.M. Wilt, et al., “Monolithic

Interconnected Modules (MIMs) for

Thermophotovoltaic Energy

Conversion,” Journal of Semiconductor

Science and Technology, accepted for

publication in 2003.

6. D.M. Wilt, D. L. Chubb,

“Thermophotovoltaic Energy

Conversion Technology Development at

NASA Glenn Research Center,” Proc.

16th European PV Conf, Glasgow,

Scotland, May 2000.

7. D.M. Wilt, et al., “0.55eV n/p/n MIM

TPV Cell Development,” Proc. 28th

PVSC, Anchorage, Alaska, 2000.

NASA/CP—2005-213431 251
 ELECTROSTATIC CHARGING OF MIRRORS IN SPACE: 

 A PLAUSIBLE CAUSE OF SOLAR PANEL ANOMALIES ON SATELLITES

Shu T. Lai 

Space Vehicles Directorate

Air Force Research Laboratory 

Hanscom AFB., MA 01731

ABSTRACT 

The entire fleet of Boeing Model 702 geosynchronous satellites has suffered from a similar fate: degradation 

of the solar cell panels. Mirrors flank both sides of the solar cell panels. Degradation, sometimes sudden 

and stepwise, shortens the lifetime of the solar cells. We suggest that space environment effects play an 

important role in damaging the solar cells. As a cornerstone in this idea, we expound a theorem that high 

reflectivity reduces photoemission. With little or no photoemission, mirrors often charge to minus kilovolts in 

eclipse as well as in sunlight, whenever the space plasma is hot enough. Since the rest of the solar panel 

does not have this mirror property, differential charging between the mirrors and the rest of the solar panel 

occurs during eclipse exits. We show the charging data obtained during an eclipse exit on LANL-97A 

satellite for supporting the idea of differential charging. Finally, we recommend this important mirror charging 

property to be taken in account in future solar panel designs and in commercial products of spacecraft 

charging computer codes. 

INTRODUCTION

This work is motivated by the news [1] that all Boeing “Model 702” 

geosynchronous satellites have suffered from fast degradation of their 

solar cells. In particular, PanAm’s PAS-7, with solar panels of the same 

design, suffered from a 25% degradation of its solar cell efficiency, 

when the satellite came out of its eclipse, Sept 6, 2001 [2]. These 

satellites feature long solar panels flanked by reflectors (mirrors) on 

both sides [Figure 1]. The purpose of using the mirrors is to enhance 

the efficiency of sunlight collection on the solar cells. Boeing has now 

decided to withdraw the mirrors from all future 702 satellites. Motivated 

by this news, we wish to offer a plausible cause. The cause is built on 

a cornerstone, which we call mirror charging. We will build this 

cornerstone in this paper. While the engineering details of the 702 

satellites are unknown, we do not attempt to justify our theory with the 

actual cases. The true cause may never be known. 

F

Figure 1. Boeing model 702 

satellite with solar panels flanked 

by mirrors [3] 

NASA/CP—2005-213431 252
PHOTOEMISSION FROM MIRRORS 

Photoelectrons from surfaces are generated by photons impacting the surface. According to Einstein, a 

photon can generate a photoelectron of energy E, only if the photon energy h< exceeds the work function W. 

h EW ν = + (1) 

With a high reflectance surface, the reflected photons are almost as intense as the incoming ones, implying 

little energy is lost in the reflection. As a result, little energy is imparted to the surface material for generating 

photoelectrons. The most important spectral line in sunlight is the Lyman ", which has h< =10.2 eV [4]. The 

work function of typical spacecraft surfaces is about W = 5 eV [5]. If, for example, only 48% of the sunlight 

photon energy h< is available, there would not be enough energy to generate photoelectrons. The 

photoelectron yield Y per unit incident photon is related to the photoelectron yield ( per unit absorbed photon 

by the relation [6,7]: 

Y R = − γ (1 ) (2) 

where R is the reflectance. The depth of reflectance is shallower than the depths of ionization, photoemission 

and attenuation [8]. 

 Modern mirrors in space are made of aluminum. The reflectance R of highly polished aluminum surfaces is 

about 90% in the Lyman " wavelength region [4]. With a commercial space mirror coating [9], the reflectance 

R is in the upper 80%. In view of such high reflectance R, we conjecture that only little photoemission is 

generated from high reflectance surfaces in space. 

MIRROR CHARGING 

Surfaces often charge to high voltages (up to hundreds or thousands of negative Volts) in hot plasmas at 

geosynchronous altitudes in eclipse. Charging occurs if the space plasma electron temperature exceeds a 

critical temperature [10]. The exact potential is controlled by a current balance equation [11]: 

e ( )1 ( ) ( ) ( ) ( ) [ ] s b i ph I II II φ φφ φ φ −− −= (3) 

where Ie is the incoming electron current, Is is the secondary electron current, Ib the backscattered electron 

current, Iph the outgoing photoelectron current, and N the surface potential. 

The photoemission current Iph from most surfaces usually exceeds the incoming plasma electron current Ie. 

Therefore, no charging to negative potentials is expected in sunlight for most surfaces. Charging of 

conducting surfaces in sunlight is usually to a few positive Volts only because photoelectrons have only a few 

eV in energy E [12]. 

However, there is only little photoemission from mirrors. 

A mirror behaves in sunlight as if it is almost in eclipse, 

as far as charging is concerned. Therefore, we 

conjecture that, mirrors charge to negative potentials no 

matter if they are in sunlight or in eclipse. 

As an example, suppose we consider two surfaces side 

by side on a geosynchronous spacecraft. Suppose they 

have about the same surface properties except one (A) is 

a mirror while the other (B) is not. During eclipse 

passage, there is no photoemission and therefore both 

behave similarly, viz., charging to negative kilovolts. Upon Figure 2. Development of differential charging. 

eclipse exit, sunlight shines, surface A continues to charge 

to high negative volts because A emits little or no photoemission, and surface B charges to a few positive 

Volts as a result of photoemission. The situation becomes one of differential charging [Figure 2]. 

NASA/CP—2005-213431 253
Furthermore, it is well known that hot electrons of the space plasma at the geosynchronous midnight region 

convect eastwards to the morning sector as a result of cross-geomagnetic tail electric field and curvature drift 

by the Earth’s magnetic field. A geosynchronous satellite orbits the Earth in the same direction of the Earth’s 

rotation, viz., eastwards. Therefore, upon eclipse exit, the satellite may encounter hot electrons. Spacecraft 

charging occurs if the plasma electrons are hotter than a critical temperature, typically 1 to 3 keV depending 

on the surface material [10]. Indeed, statistics show that spacecraft charging to high negative voltages is 

more likely in the morning sector. 

SPACE HAZARD 

Differential charging may cause adverse effects. Sudden development of differential charging is potentially 

hazardous. Suppose two neighboring surfaces (a mirror and a non-mirror) develop a potential difference of 

about a kilovolt. There can be two adverse effects; (1) a discharge between them may suddenly occur, and 

(2) sputtering may cause slow degradation of the mirror, reducing its efficiency. A discharge can cause a 

sudden and stepwise degradation to the solar panel system. Sputtering may pose a long term effect, viz., 

slow degradation of the system, resulting in shortening the expected lifetime of the solar panel system. 

The PAS-7 satellite did not have instruments 

onboard to measure charging or the hot 

electrons. The nearest satellite that had this 

capability was the Los Alamos National 

Laboratory satellite LANL-97A. Indeed, 

LANL-07A showed charging to N = –3 kV 

during the entire eclipse passage in early 

morning of Sep 6, 2001 [Figure 3]. The 

charging level N dropped to N = 0 abruptly 

upon eclipse exit. Although PAS-7 had no 

such measurement capabilities, we infer that 

PAS-7 might behave likewise, viz,., charging 

to about N = –3 kV during the eclipse 

passage. Upon the PAS-7 eclipse exit, N

might drop to 0 abruptly. However, the PAS-7 

mirrors probably continued to behave almost 

as if in eclipse, because they emitted little or 

no photoelectron current no matter if in 

eclipse or in sunlight. As a result, differential 

charging of the order of kV might emerged 

abruptly upon eclipse exit in the morning sector. Figure 3. Eclipse charging of LANL 97-A satellite to –3 kV. 

CONCLUSION 

All Boeing model 702 satellite solar panels flanked by mirrors have experienced gradual or sudden 

degradations. The true cause is unknown and Boeing has ceased to use this design. We have offered a new 

concept – mirror charging. We conjecture that mirrors do not generate much photoemission. Accordingly, 

they charge to high negative voltages no matter if in eclipse or in sunlight. When a mirror and a neighboring 

non-mirror surface come out from eclipse, differential charging up to kV may occur, depending on the 

temperature of the space plasma. The effects of differential charging are potential hazards of discharges 

between surfaces and gradual degradation of the solar panel system due to sputtering by keV ions. 

We suggest that the concept of mirror charging be studied in the laboratory and in space. In laboratory 

studies, care must be taken to separate the photoemissions from mirror samples and from the vacuum 

chamber walls. In space studies, measurements on differential charging should be made. Finally, we 

suggest that mirror charging be incorporated into commercial codes of spacecraft charging. 

NASA/CP—2005-213431 254
ACKNOWLEDGMENT 

The Los Alamos Magnetospheric Plasma Analyzer (MPA) measurements were obtained from the CDAWeb 

data service at NASA Goddard Space Flight Center. We thank Michelle Thomsen for permission to use the 

MPA data.

REFERENCES 

[1] Space News, p.20, Oct 1, 2001.

[2] http://sat-nd.com/failures/timeline.html

[3] http://www.hughespace.com/factsheets/702/ 

[4] Hinteregger, H.E., K.R. Damson and L.A. Hall, Analysis of photoelectrons from solar extreme ultraviolet, J. 

Geophys. Res., Vol.64, 961-964, 1959. 

[5] CDC Handbook of Physics and Chemistry, CDC Press, 2002.

[6] Highes, A.L. and L.A. Dubridge, Photoelectric Phenomena, McGraw-Hill, New York, 1932. 

[7] Samson, J.A.R., Techniques of Ultraviolet Spectroscopy, John Wiley, New York, 1967. 

[8] Spicer, W.E., Photoelectric emission, in Optical Properties of Solids, F. AbelJs (ed.), pp.755-858, NorthHolland Publishing Co., Amsterdam and London, 1972. 

[9] http://www.oriel.com/down/pdf/12010.pdf. 

[10] Lai, S.T. and D. Della-Rose, Spacecraft charging at geosynchronous altitudes: New evidence of the 

existence of critical temperature, J. Spacecraft & Rockets, vol.38, No.6, 922-928, 2001. 

[11] Hastings, D. and H. B. Garrett, Spacecraft Environment Interactions, Cambridge University Press, 

Cambridge, UK, 1996. 

[12] Lai, S.T., H.A. Cohen, T.L. Aggson, and W.J. McNeil, Boom potential on a rotating satellite in sunlight, J. 

Geophys. Res., vol.91, No.A11, pp.12137-12141, 1986. 

NASA/CP—2005-213431 255
EXPERIMENTAL STUDY OF ARCING ON HIGH-VOLTAGE SOLAR ARRAYS 

Boris Vayner, 

Ohio Aerospace Institute, Brook Park, Ohio 44142 

Joel Galofaro, and Dale Ferguson 

NASA Glenn Research Center, Cleveland, Ohio 44135 

1. Introduction 

 The main obstacle to the implementation of a high-voltage solar array in space is arcing on the 

conductor-dielectric junctions exposed to the surrounding plasma. One obvious solution to this 

problem would be the installation of fully encapsulated solar arrays which were not having 

exposed conductors at all. However, there are many technological difficulties that must be 

overcome before the employment of fully encapsulated arrays will turn into reality. An alternative 

solution to raise arc threshold by modifications of conventionally designed solar arrays looks 

more appealing, at least in the nearest future. A comprehensive study of arc inception 

mechanism [1-4] suggests that such modifications can be done in the following directions: i) to 

insulate conductor-dielectric junction from a plasma environment (wrapthrough interconnects); ii) 

to change a coverglass geometry (overhang); iii) to increase a coverglass thickness; iiii) to outgas 

areas of conductor-dielectric junctions. The operation of high-voltage array in LEO produces also 

the parasitic current power drain on the electrical system. Moreover, the current collected from 

space plasma by solar arrays determines the spacecraft floating potential that is very important 

for the design of spacecraft and its scientific apparatus. In order to verify the validity of suggested 

modifications and to measure current collection five different solar array samples have been 

tested in large vacuum chamber. Each sample (36 silicon based cells) consists of three strings 

containing 12 cells connected in series. Thus, arc rate and current collection can be measured on 

every string independently, or on a whole sample when strings are connected in parallel. The 

heater installed in the chamber provides the possibility to test samples under temperature as high 

as 80 C that simulates the LEO operational temperature. The experimental setup is described 

below. 

 2.Experimental setup 

 Low Earth Orbit (LEO) plasma environment was simulated in two different vacuum vessels: 1) 

horizontal vacuum chamber (1.8 m diameter and 2.5 m long) equipped with cryogenic pump; 2) 

large vacuum tank (1.8 m diameter and 3 m height) with four diffusion pumps. The vacuum 

equipment provided pressure as low as 0.5 μTorr. The essential difference between these two 

tanks is that the residual water vapor partial pressure in horizontal chamber is five times lower than 

in vertical tank. Each vessel has one Kaufman plasma source that generates xenon (or argon) 

plasma with electron density ne=(0.1-10)·105

 cm-3 , temperature Te=0.6-1.2 eV, and neutral gas 

pressure p=(0.7- 7)·10-5 Torr which can be kept steady during the experiment. To measure plasma 

parameters, Langmuir probes with diameter 2 cm were employed (two in each tank). To determine 

an ion distribution function and to improve measurements of electron temperature one retarding 

potential analyzer (RPA) was mounted on the bottom of vertical tank. It was found that the ion 

(xenon) thermal flux in the experiment is about three times lower than ram ion flux in LEO, and the 

electron temperature is 5-10 times higher than in ionosphere. However, the number densities are 

simulated with a quite high accuracy, and one can believe that the results of high-voltage 

experiments in vacuum chambers are fairly adequate to the outcomes of processes in LEO plasma. 

NASA/CP—2005-213431 256
To control plasma chemical composition (particularly, water vapor and oil partial pressures) a 

quadruple mass spectrometer was installed in each tank. 

 The sample (or set of samples) is vertically mounted in the middle of the chamber, and it is 

biased to a voltage power supply through a capacitor and a 10 kΩ resistor network back to ground. 

An additional power supply (Solar Array Simulator-SAS) is used to generate electrical field 

perpendicular to the dielectric side surface for investigating arc inception on semiconductordielectric junction and inception of sustained discharges between adjacent strings. Diagnostic 

equipment includes two current probes to measure discharge current and SAS current, and one 

voltage probe that allows us to register voltage pulse on the sample during the discharge (Fig.1). 

To measure optical spectra of arc plasma an intensified CCD (1024x512 pixel) camera with optical 

spectrometer is installed (Fig.2). The arc sites are determined by employing a video camera and 

VCR. Most experiments were performed at room temperature (15C), but some tests had been 

done at the temperature +80 C simulating the exposure of solar array to full sun in LEO. 

Fig.1. Circuitry diagram for arc inception study. 

3. Arc spectra 

 Previous experimental data and theoretical analysis have demonstrated that water molecules 

absorbed on the side surface of dielectric (coverglass+adhesive) can play a decisive role in the 

process of arc inception [5-7]. It is known that spectra of vacuum arcs consist of cathode metal 

lines only [8] but adding air in vacuum chamber (10-4 Torr) results in appearance of hydrogen and 

hydroxyl lines [9]. Partial pressure of water vapor and nitrogen is always below 10-5 Torr for the 

current experiment. Thus, the presence of hydrogen and hydroxyl in the arc plasma would be a 

good indicator of water ions dissociative recombination. The presence of other species in arc 

plasma may reveal other important processes in the discharge development. To elucidate all 

these problems the measurements of optical spectra have been performed for silicon solar array 

sample (Fig.3). All dielectric-conductor junctions besides one interconnect area were insulated by 

NASA/CP—2005-213431 257
tape to exclude spectra from arcs between cells. The spectral resolution was determined as 0.12 

nm/pixel by using standard calibration lamps. To increase the arc luminosity an additional

Fig.2. Setup of optical spectra measurements. 

capacitor (usually 1 μF) was installed between negatively biased electrode and ground. Gate pulse 

generator (Fig.2) provided varying both gate pulse width and time delay; thus, it was possible to 

measure the intensities of spectral lines on different stages of a discharge development. One 

example of the time sequence of discharge current pulse and gate pulse is shown in Fig.4. In 

addition to hydrogen (H ), hydroxyl (OH), and metal lines (Ag) arcing on the sample revealed also 

some molecular radicals identified according to Ref. 10 (Fig.5). 

 The results of spectral measurements, observation of decreasing arc rate with number of arcs 

(conditioning), and theoretical estimates [11] are very strong argument in favor of the idea that in 

order to raise arc threshold solar array surfaces must be thoroughly outgassed. The validity of this 

hypothesis has been confirmed by ground tests described in Sec. 5 below. 

NASA/CP—2005-213431 258
Fig.3. Silicon solar array sample used for measurements of arc spectra. 

Fig.4. One example of pulses time sequence applied for spectral measurements. 

NASA/CP—2005-213431 259
 

Ag I 

Ag III 

Arc spectrum (Ag lines)

Hα 

NASA/CP—2005-213431 260
Fig.5. One example of emission spectrum of arc plasma. 

 

OH band

NASA/CP—2005-213431 261
 

 4. Contamination of plasma environment 

 When spacecraft is coming out of eclipse, solar array temperature is rising due to exposure to 

Sun radiation. Operational temperature of an array in LEO conditions is approximately 80-100 C. It 

is believed that the adsorbed contaminants from the array surface are evaporating at a high rate 

due to heating. On another hand, when solar array sample is installed in a vacuum chamber its 

surface is contaminated not only by “natural” species (adsorbed water, atmospheric gases, and 

products of a technological process) but also other contaminants. The most abundant of these 

contaminants is vacuum pump oil. To measure chemical composition of background vacuum and 

to determine plasma contamination due to heating, the quadruple mass-spectrometer is installed in 

large chamber. Two solar array samples consisting of 36 cells (4x6 cm) each are mounted on an 

aluminum sheet with electrical heater placed on the back (Fig.6). This heater provides enough 

power to radiately heat sample from 15 C to 80 C for about 50 minutes (Fig.7). It is seen that the 

increase in water vapor partial pressure is considerably higher than the plain isochoric increase 

0 T0

T

p

p ∆ = ∆

 , and this observation confirms the presence of water absorbed on solar array surface. 

Fig.6. Solar array samples installed in vacuum chamber. 

NASA/CP—2005-213431 262
Fig.7. Increase of water vapor pressure is considerably higher than 

 plain isochoric increase. 

Five cycles of heating-cooling sample in vacuum chamber resulted in significant drop of residual 

water vapor pressure (Fig.8). 

 

Fig.8. Water vapor partial pressure in course of thermal cycling (rel. units). 

Initial Heating 

NASA/CP—2005-213431 263
 5.Arcing in plasma 

 Five types of tested solar arrays are shown in Table 1. Each string (12 cells in series) is tested 

separately to measure arc inception voltage and arc rate. Measurements reveal significant 

differences in these parameters even for strings belonging to one sample. There are two reasons 

explaining such observations: manufacturing process peculiarities and geometrical design of a 

sample. In fact, the middle string is separated from neighboring strings by narrow gaps (0.8 mm) 

covered with a thin RTV layer while two other strings have edges with underlying semiconductor 

and dielectric exposed to the plasma. Manufacturing peculiarities demonstrate themselves when 

one compares arc parameters for two outer strings and finds considerable differences. And arc 

sites are located mostly on interconnects for middle string while great part of arcs on outer strings 

has been observed on cell edges. To preserve the homogeneity of collected data one common 

experimental procedure is used for all measurements of arc inception voltages and arc rates: 1) 

string is initially biased to voltage well below an expected arcing threshold; 2) 15-30 minute time 

interval is allowed to register (or to not register) an arc; 3) voltage is increased on 10-20 V; 4) arc 

rate is defined as an average over a respective time span. 

Table1. Five types of solar array samples tested in the large chambers. 

Sample.: Coverglass : Overhang : Cell size : Interconnect 

 No : Thickness (μm): Material : (μm) : (cm) : 

________________________________________________________________________ 

1 300 UVR 0 4x6 exposed 

2 150 UVR 0 4x6 exposed 

3 150 CMX UVR 0 4x6 exposed 

4 150 UVR 250 4x6 exposed 

5 150 UVR 0 8x8 wrapthrough 

________________________________________________________________________ 

 On the first stage of the test, two samples (#1 and #2 in Table 1) are mounted on the heater 

plate and installed in chamber. The results of measurements for middle strings at the room 

temperature are shown in Fig.9. Obviously, arc inception voltage is lower for the panel with 

thinner coverglass, and arc rates differ significantly. Arc rates have been also determined at 

high temperature (Fig.10). Arc rates are widely scattered over a range of voltages 280-380 V. In 

general, the temperature rise to 80 C results in significant increase of arc inception voltage (40-60 

V). In particular, inner strings are not arcing below 300 V. It is worth noting that measurements 

shown above have been done at comparatively high water vapor partial pressures: 4 μTorr at 15 

and 15-30 μTorr at 80 C. These values are much higher than one can anticipate in LEO 

conditions. 

The decrease of an arc rate during the process of continuing arcing (conditioning) has been 

measured by biasing the whole sample #2 to -400 V and measuring average arc rate for every 

four minutes. Additional capacitance is increased from 0.22 μF to 1μF to accelerate conditioning. 

After about 70 arcs, arc rate drops from 3.25 arc/min to the magnitude of 1 arc/min and stays 

practically steady for the next 30 arcs. To verify the influence of plasma density on arc rate this 

parameter has been increased by factor 1.5, and arc rate was measured for the next 100 arcs. 

Finally, arc rate has decreased to 0.25 arc/min after about 200 arcs. Thus, the influence of 

conditioning on previous measurements of arc rates for separate strings belonging to different 

samples is insignificant, particularly because of low capacitance (0.22 μF) used in these tests. 

NASA/CP—2005-213431 264
Fig.9. Arc rate on the middle strings (samples #1 and #2). 

Fig.10. Arc rate on the middle string of sample #2. 

 To test the possibility of outgassing of the whole sample by heating it to 80 C and pumping 

out an excess of water vapor the sample #2 has been undergone to five thermal cycles (see 

Fig.8). Arc rates are measured for all three strings connected in parallel at room temperature 

before the first cycle and after the fifth cycle. The results are shown in Fig.11. 

NASA/CP—2005-213431 265
Fig.11. Arc rate on sample #2 (three strings in parallel). 

It is seen in Fig.11 that arc rate decreases about 30 times and arc threshold increases 

approximately 30 V due to outgassing. It should be noted that minimum partial pressure of water 

vapor in xenon plasma reached in these experiments was 1.4 μTorr that seems not low enough to 

outgas sample surfaces to the degree expected in LEO. 

 Arcing on the sample #3 (CMX UVR coverglass) does not show any measurable differences 

comparatively to arcing on sample #2. Arc inception voltage for the middle string is 300 V, and 

two other strings are arcing at lower voltages due to considerable percentage of arcs on the cell 

edges. On the contrary, the test results for the sample #4 (250 μm overhang) look much more 

prospective. Arc inception voltage is 80-100 V higher, and arc rate is lower for the middle string. 

Two other strings have also demonstrated the decrease of arc rate in spite of arcing on cell 

edges. The increase of arc inception voltage to 480 V for the hot sample is particularly important. 

It seems that the array with coverglass overhang and additional insulation of cell edges can 

operate at 400 V in LEO conditions. 

 According to existing model of arc inception [12,13] the most probable arc site on an array 

surface is a conductor-dielectric junction exposed to the plasma. Thus, if all interconnects are 

insulated from the surrounding plasma the probability of a discharge decreases significantly. One 

of the possible realizations of this idea is the array design with wrapthrough interconnects 

(sample #5, Fig.12). Such design cannot prevent arcing at very high negative potential because 

edge of semiconductor (silicon, germanium, or other) stays exposed to the plasma, and many 

tests (including ones described in this paper) have demonstrated intensive arcing on cell edges. 

However, considerable increase of arc inception voltage can be expected, particularly for the 

middle string. Test results confirm these expectations (Fig.13). The inception of arc is observed 

on the middle string at bias voltage 440 V that is 60 V higher than arc inception voltage for the 

NASA/CP—2005-213431 266
Fig.12. Solar array sample with wrapthrough interconnects. 

string with coverglass overhang. Arc sites are located between adjacent cells belonging as to 

middle string as to neighboring strings. Two other strings demonstrate much lower arc inception 

voltages due to arcing on edges. Unfortunately, experimental setup has not provided a possibility 

to heat this sample above room temperature but even the results obtained to date show that this 

kind of solar array can be used in LEO to generate power at voltage 450 V if array edges are 

electrically insulated. 

Fig.13. Arc rate on sample #5. 

 Short electrostatic discharges studied above are certainly undesirable events that must be 

prevented for reliable operation of the spacecraft. However, this kind of transients are not 

damaging solar array irreversibly, at least in cases of low additional capacitances used in current 

experiments. Sustained discharges initiated between adjacent cells with a few tens volt potential 

difference are much more dangerous [14,15] because they can destroy cells and underlying 

substrate that results in considerable loss of power. Samples #1, #2, and #4 have been tested 

against an inception of sustained arc between two strings. The circuitry diagram for the test is 

shown in Fig.1. Test starts with lower limits on SAS voltage and current. After the registration of 

5-10 arcs these parameters are gradually changed and more arcs are generated until initial sign 

NASA/CP—2005-213431 267
of sustained discharge is seen on the oscilloscope. This sign represents the SAS current pulse 

that continues much longer than original arc. The corresponding SAS voltage and current are 

considered as threshold parameters because even a small increase of them (10 V and 0.25 A) 

results in spectacular event shown in Fig. 14. In this case the sustained discharge has been 

quenched after 20 s by turning SAS off. Damaged part of the sample is shown in Fig.15. 

Threshold parameters depend on solar array design: they are 40 V and 1 A for sample #1, 60 V 

and 2 A for sample #2, and 80V and 1.6 A for sample #4. 

Fig.14. Sustained arc between adjacent strings on sample #2. 

5. Scaling of arc parameters 

 Even short transients are detrimental for spacecraft, and the degree of damage increases with 

the increase of arc current amplitude and pulse width. These two parameters depend on the 

amount of electrical charge leaking into surrounding plasma during the discharge time. There are 

currently two theoretical models that allow estimating lost electrical charge and its dependence on 

the array capacitance. First model7

 is based on the suggestion that the discharge generates an 

expanding plasma sheath neutralizing positive charge on top of coverglass. If plasma expends with 

a constant speed the discharge time is proportional to the array linear dimensions, or, in another 

terms, to the square root of an array capacitance. This dependence has been proved in many 

experiments [16,17]. However, the distance that plasma can expand on is limited to about 1 m in 

simulated LEO conditions [18]. Thus, according to the first model the upper limit for the effective 

capacitance is the capacitance of the part of solar array with area approximately 1 m2

. The second 

model also envisages that both arc current amplitude and pulse width are proportional to the square 

root of a capacitance but this prediction is based on the dynamics of ionization-recombination 

processes in the discharge plasma [6]. If the second model is correct the effective capacitance is 

only two-three times less than the capacitance of a whole solar array. A simple experiment has 

been performed to verify the validity of the second model. Two solar array 

NASA/CP—2005-213431 268
Fig.15. Damage induced by sustained arc. 

samples (sample #2) are mounted on aluminum panel with grounded aluminum plate installed 

between samples. The height of the plate is 7.5 cm bigger than the distance between aluminum 

panel and top of the sample. Such arrangement prevents the expansion of plasma sheath from 

one sample to another. The additional capacitor of 1000 pF is used in this particular experiment. 

 The capacitance of one cell can be calculated as 

1

2 1

1 2

1

0 1 1

1 1

−

⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛ = +

d

d

d

S C

ε

ε ε ε

 (1) 

where ε 1,2 are dielectric constants of coverglass and adhesive, d1,2 are their thicknesses, and S1 is 

a cell area. 

Dielectric constants are: for borosilicate coverglass ε 1=6.7, and for epoxy ε 2=3.6. Thus, a quite 

reliable estimate can be obtained: C1=590 pF/cell. Moreover, the scaling does not practically 

depend on exact numbers for largest capacitances. Ten measurements of arc current pulse 

widths for each configuration have been done by biasing one string, three strings, and six strings 

in parallel. The results are shown in Fig.16. The scaling is confirmed with a very high accuracy, 

which means that adequate ground simulations of arcing on spacecraft surfaces have to be 

performed with a very large additional capacitance (for instance, about 1000 μF for ISS). 

7.Current collection 

 One solar cell provides current of 1 A in order of magnitude while collected current is scaled in 

hundred microamps. Thus, the role of collected current in a parasitic power drain is certainly 

negligible. However, the floating potential of the spacecraft strongly depends on the current 

collected by the solar array [19]. There are three main factors that influence the magnitude of 

collected current: i) solar array design; ii) solar array temperature; iii) parameters of surrounding 

NASA/CP—2005-213431 269
plasma. Obviously, the design with coverglass overhang and with wrapthrough interconnects 

offers arrays with considerably decreased collected currents. Electron number density and 

electron temperature also influence on current collection. Ground tests that simulate an electron 

component of LEO plasma quite reasonably provide reliable data for current collection by cells 

with positive potentials with respect to surrounding plasma. Test data containing measurements 

of collected current for negatively biased cells are applicable to the analysis of spacecraft floating 

potential not better than in order of magnitude because the characteristics of ion component are 

different in ground tests and in LEO. 

Fig.16. Pulse width scaling measured by biasing to -340 V one, three, and six 

 strings of sample #2. Error bars (±1 ) are calculated from ten measurements 

 for each point. 

 A few examples of the dependence of electron current collection on bias voltage are shown in 

Fig.17. It is seen than the increase of electron number density results in the almost proportional 

increase in current collection. But the dominant factor in the current collection is an array 

temperature. The magnitude of collected current grows more than three times when array 

temperature reaches 79 C. This observation must be taken into account for the computations of 

spacecraft floating potentials. Ion currents are measured by biasing separate strings up to 100 V 

negative, and these currents do not exceed 1 μA for all situations studied even though the same 

effect of significant increase due to heating is also found. 

 Measurements of collected currents for the sample with coverglass overhang have 

demonstrated the decrease in magnitude close to the factor 2 comparatively to sample with a 

standard design. Cell with wrapthrough interconnects collects not much less current than cell with 

coverglass overhang but it generates three times higher power. It seems that tests in simulated 

plasma environment are suitable for creation a data base for further computations of the 

spacecraft floating potentials in LEO. 

8. Conclusions 

 Comprehensive tests of five different types of solar array samples in simulated LEO plasma 

environment have demonstrated that the highest arc threshold (440 V) can be achieved for an 

array with wrapthrough interconnects if edges of strings are not exposed to the plasma. This 

design is also effective in decreasing of an array current collection. The design with exposed 

interconnects but with coverglass overhang also provides significant improvement comparatively 

to the conventional design. Particularly, arcing on the sample cannot be initiated at potentials 

below 300 V even under room temperature, and arc threshold increases to 420 V under 

NASA/CP—2005-213431 270
temperature 72 C. The increase of coverglass thickness also results in some improving of array 

parameters. Thorough outgassing of solar array surfaces may result in significant decrease of 

arc rate for a conventionally designed solar array. 

a)

b) 

Fig.17. Electron current collection: a) 1-sample #2, str.2, ne=2*105

 cm-3,15 C; 2-sample #1, str.1, 

 ne=7*105

 cm-3,15 C; 3-sample #1, str.2, ne=2*106 cm-3,15 C; 4-sample #2, 

 str.1, ne=5*105

 cm-3, 79 C. b)1-sample #3, str.2, ne=1*106 cm-3; 2- sample #4, str.2, 

 ne=1*106

 cm-3; 3 and 4-sample #5, str.1 and 2 respectively, ne=3.5*105 cm-3, temp. 15 C. 

References 

1.Stevens, N.J. “Interactions Between Spacecraft and the Charge-Particle Environment”, Proc. 

 Spacecraft Charging Technology Conference, NASA CP-2071, 1978, p.268-294. 

2.Ferguson, D.C. “Solar Array Arcing in Plasmas”, NASA CP-3059, 1989, p. 509. 

3.Hastings, D.E. “A Review of Plasma Interactions With Spacecraft in Low Earth Orbit”, 

NASA/CP—2005-213431 271
 Journal of Geophys. Research, Vol.100, No.A8, 1995, pp.14,457-14,483. 

4.Parks,E.D.,Jongeward,G.,Katz,I., and Davis, V.A. “Threshold-Determining Mechanisms for 

 Discharges in High Voltage Solar Arrays”, Journal of Spacecraft and Rockets, Vol.24, 

 No.4, 1987, pp.367-371. 

5. Upschulte, B.L., Marinelli, W.J., Carleton, K.L., Weyl, G., Aifer, E., and Hastings, D.E. 

 “Arcing on Negatively Biased Solar Cells in a Plasma Environment”, Journal of 

 Spacecraft and Rockets, Vol.31, No.3, 1994, pp.493-501. 

6.Vayner, B., Galofaro, J., Ferguson, D., and Degroot, W. “Electrostatic Discharge 

 Inception on a High-Voltage Solar Array”, AIAA Paper 2002-0631, Jan.2002. 

7. Galofaro, J., Vayner, B., Ferguson, D., and Degroot, W. “A Desorbed Gas Molecular 

 Ionization Mechanism for Arcing Onset in Solar Arrays Immersed in a Low-Density 

 Plasma”, AIAA Paper 2002-2262, May 2002. 

8. Anders, A., and Anders, S. “Emission Spectroscopy of Low-Current Vacuum Arcs”, 

 Journal of Physics D:Appl.Phys., Vol.24, 1991, pp.1986-1992. 

9.Yotsombat, B., Davydov, S., Poolcharuansin, P., Vilaithong, T., and Brown, I.G. 

 “Optical Emission Spectra of a Copper Plasma Produced by a Metal Vapor Vacuum Arc 

 Plasma Source”, Journal of Physics D: Appl.Phys., Vol.34, 2001, pp.1928-1932. 

10. Pearse, R.W.B., and Gaydon, A.G. “The Identification of Molecular Spectra”, 3rd

 

 edition, Chapman&Hall LTD, London, UK, 1963, 347p. 

11 Vayner, B.V., Galofaro, J.T., and Ferguson, D.C. “The Neutral Gas Desorption and 

 Breakdown on a Metal-Dielectric Junction Immersed in a Plasma”, AIAA Paper 2002- 

 2244, May 2002 

12.Cho,M., and Hastings, D.E. “Computer Particle Simulation on High-Voltage Solar Array 

 Arcing Onset”, Journal of Spacecraft and Rockets, Vol.30, No.2, 1993, pp.189-205. 

13.Hastings,D.E., Weyl,G., and Kaufman, D. “Threshold Voltage for Arcing on Negatively 

 Biased Solar Arrays”, Journal of Spacecraft and Rockets, Vol.27, No.5, 1990, p.539-544. 

14.Hoeber, C.F., Robertson, E.A., Katz, I., Davis, V.A., and Snyder, D.B. “Solar Array 

 Augmented Electrostatic Discharge in GEO”, AIAA Paper 98-1401, Jan. 1998. 

15. Snyder, D.B., Ferguson, D.C., Vayner, B.V., and Galofaro, J.T. ”New Spacecraft- 

 Charging Solar Array Failure Mechanism” 6th Spacecraft Charging Technology 

 Conference, Nov. 2-6, 1998, Air Force Res. Lab., Hanscom AFB, MA, USA, pp.297- 

 301. 

16. Balmain, K. G., and Dubois, G. R. “Surface Discharges on Teflon, Mylar and Kapton”, 

 IEEE Transactions on Nuclear Science, Vol. 26, No.12, 1979, pp.5146-5151. 

17. Snyder, D.B. “Characteristics of Arc Currents on a Solar Cell Array in a Plasma”, 

 IEEE Transactions on Nuclear Science, Vol.31, No.4, 1984, pp.1584-1587. 

18.Cho, M., Ramasamy, R., Hikita, M., Tanaka, K., and Sasaki, S. “Plasma Response to 

 Arcing in Ionospheric Plasma Environment: Laboratory Experiments”, Journal of 

 Spacecraft and Rockets, Vol.39, No.3, 2002, pp.392-408. 

19. Ferguson, D.C., and Hillard, G.B. “In-Space Measurements of Electron Current 

 Collection by Space Station Solar Arrays”, AIAA Paper 95-0486, Jan. 1995. 

NASA/CP—2005-213431 272
LIQUID PHASE DEPOSITION OF SINGLE-PHASE ALPHA-COPPER-INDIUM-DISELENIDE 

J. Cowen 

Case Western Reserve University, Cleveland, Ohio 

L. Lucas 

Case Western Reserve University, Cleveland, Ohio 

F. Ernst 

Case Western Reserve University, Cleveland, Ohio 

P. Pirouz 

Case Western Reserve University, Cleveland, Ohio 

A. Hepp 

National Aeronautics and Space Administration, Cleveland, Ohio 

S. Bailey 

National Aeronautics and Space Administration, Cleveland, Ohio 

Introduction 

The success of exploratory missions in outer space often depends on a highly efficient renewable energy supply, 

as provided by solar cells. Figure 1 shows a well-known example: The robotic vehicle “Rover,” constructed for 

NASA’s “Mars Pathfinder” mission. The solar cells for such applications not only need to have high conversion 

efficiency, but must possess a high specific power, thus a high power output per unit mass. Since future missions 

will demand for large aggregates of solar cells and space flights are expensive, the solar cells must furthermore 

be available at low costs (per unit power output) and – very important in outer space – have a long lifetime and a 

high resistance against structural damage introduced by irradiation with high-energy electrons and protons. 

The photovoltaic materials that are presently available only partly fulfill all these requirements. Therefore, we 

propose to explore a new method for fabricating thin-films for cost-efficient solar cells with very high specific 

power, high irradiation resistance, and long lifetime, based on the α-phase of the Cu–In–Se system, “α-CIS.” 

Fig. 1: The robotic vehicle “Rover,” which NASA 

sent to the surface of planet Mars (exhibit at NASA 

Glenn Research Center, Cleveland). 

NASA/CP—2005-213431 273
The stoichiometric composition of α−CIS corresponds to 

CuInSe2. It has long been known that this compound 

possesses outstanding properties for thin-film solar cells. 

So far, however, one cannot fully exploit the advantages of 

this material, because it is not possible to make α-CIS 

layers with a sufficiently low concentration of structural and, 

consequently, electronic defects. In this paper, we describe 

a research program to overcome these limitations and to 

optimize the fabrication of α-CIS thin-films for solar cell 

applications in outer space. 

Presently, the CIS films with the best photovoltaic 

performance are grown by physical vapor deposition 

(PVD). However, such films contain high concentrations of 

structural defects, which degrade their efficiency in 

photovoltaic devices: point defects, dislocations, largeangle grain boundaries, pores, anti-phase boundaries, 

cracks, or inclusions of other phases and, correspondingly, 

interphase interfaces [1]. 

Figures 2, 3, and 4 present cross-sectional TEM 

(transmission electron microscopy) bright-field images 

revealing these kinds of defects in CIS thin-films grown 

under a variety of different conditions by the group of H.-

W. Schock at the Institut für Physikalische Elektronik, 

University of Stuttgart, Germany. The 

microcharacterization, including TEM, was carried out as 

part of the Ph. D. project of T. Haalboom at the Max-PlanckInstitut für Metallforschung (Materials Science) in Stuttgart, 

Germany, under the supervision of F. Ernst. 

As apparent from the micrographs, the microstructure 

sensitively depends on the processing conditions. For a 

Cu:In:Se ratio of 1:1:2 in the vapor, high temperatures 

(between 460 and 520 °C) lead to relatively large grains. 

Low temperatures (430 °C), in contrast, lead to the 

formation of extended, plate-like crystallites. 

In addition to the defects mentioned before, thin-films 

grown with a Cu:In:Se ratio near 1:1:2 in the vapor always 

exhibit a high density of stacking faults and twin 

boundaries. 

Another problem of PVD, apart from the high defect density 

in the films, is to control the chemical composition. Analysis 

of the films in Figs. 2, 3, and 4 by XEDS (X-ray energydispersive spectroscopy) revealed that the actual, local 

chemical composition of the thin-films may substantially 

differ from the nominal composition of the vapor (indicated 

in the figure captions). As demonstrated by the X-ray scans 

in Fig. 5, even slight deviations from the ideal composition 

CuInSe2 lead to the formation of secondary phases. The 

two X-ray scans were obtained from stoichiometric and Cudeficient CIS films. Both diffraction patterns exhibit 

reflections of the (desired) α-CuInSe2 phase [2] and Moreflections of the substrate. Additional reflections at 2θ = 

Fig. 2: Cross-sectional TEM image of a CIS film 

grown at 460 to 520 °C by PVD. Cu:In:Se ratio in 

the vapor: 1:1:2. Composition measured by XEDS: 

Cu25.5In24.8Se49.8. 

Fig. 3: Cross-sectional TEM image of a CIS film 

grown at 460 to 520 °C by PVD. Cu:In:Se ratio in 

the vapor: 1:3:5. Composition measured by XEDS: 

Cu24.0In25.9Se50.2 . 

Fig. 4: Cross-sectional TEM image of a CIS film 

grown at 430 °C by PVD. Cu:In:Se ratio in the 

vapor: 1:2:3.5. Composition measured by XEDS: 

Cu16.6In32.2Se51.2. 

Fig. 5: X-ray scans of PVD-grown thin-films with

stoichiometric (CuInSe2) and with Cu-deficient

composition. 

NASA/CP—2005-213431 274
21.8° and 42.6°, however, indicate the presence of a second phase (defect phase β [3]) in the Cu-deficient 

material. Since the presence of second phases has a particularly adverse effect on the conversion efficiency [4], 

this material is not suitable for photovoltaic devices. 

The present record of conversion efficiency with PVD α-CIS thin-films is at 18 .8 % [5]. While this number is 

impressive, the theoretical conversion efficiency is still 30 % higher (24 %) [6]. Moreover, the efficiency of standard 

PVD material is way below 18 %. There is no doubt that the high concentration of grain boundaries and phase 

boundaries encountered in PVD-grown material accounts for the major part of the discrepancy between actual an 

theoretical conversion efficiency [4]. Annealing can somewhat improve the structural properties, but usually of the 

film composition (particularly the Se content) and causes segregation [7-10]. 

The conversion efficiency of PVD-grown α-CIS thin-films is limited by structural defects, which form because 

these films grow far from thermodynamic equilibrium, under high driving forces, but with small atomic mobility. A 

promising approach for improving the structural quality of α-CIS, therefore, is by growing it from the melt, close to 

thermodynamic equilibrium. 

CIS Growth from the Melt 

Near-equilibrium growth from the melt will not only allow the fabrication of single-phase, pure α-CIS, but should 

also enable the growth of macroscopic α-CIS single crystals. Single crystals of α-CIS are valuable for studying the 

intrinsic properties of α-CIS, and to find out how much room there really is for improving the conversion efficiency 

of α-CIS thin-film solar cells by improving the structural quality of the material.

Following this idea, C. R. Schwerdtfeger at the National 

Renewable Energy Laboratory (NREL) employed highpressure, liquid-encapsulated directional solidification to 

grow (Cu-deficient) α-CIS single crystals and obtained 

apparently coarse-grained macroscopic crystals [11, 12]. 

However, no study of the microstructure of this material 

has been published. More recent attempts of M. Hornung 

to grow α-CIS single-crystals from the melt by the vertical 

gradient freeze method also failed [13]. SEM (scanning 

electron microscopy) and TEM studies on the largest 

grains of these “single crystals” revealed that they were 

never even single-phase α-CIS, but always contained a 

second phase (β) and an enormous density of α/β phase 

boundaries [14]. Fig. 6 presents an example from this 

material: a TEM image revealing particles of the β-phase in 

an α-CIS “single crystal”. Note that the β particles did not 

form because the solidified material had the incorrect 

composition. Careful analysis of chemical composition by 

XRF (X-ray fluorescence analysis) and ICP-OES 

(inductively-coupled plasma optical emission spectroscopy) 

confirmed that the composition of this material was 

Cu23.6In25.9Se50.5, thus sufficiently close to the stoichiometric composition of α-CIS for obtaining a single-phase 

material. The latest reports on α-CIS single crystals are by Cahen et al. [15], but again no detailed microstructural 

studies were published. 

Cu–In–Se Phase Equilibria 

To understand these drawbacks, we have recently investigated all phase equilibria of the entire Cu–In–Se system 

and determined the first complete and correct ternary phase diagram [16-18]. 

Fig. 6: TEM image revealing the presence of a 

second phase (“β”, arrowed regions) in a “single 

crystal” of Cu23.6In25.9Se50.5. 

NASA/CP—2005-213431 275
The new phase diagram, which constitutes one of the most complicated ternary phase diagrams known so far, 

explains why it is difficult (if not impossible) to grow single-phase α-CIS from a melt with the composition equal to 

or close to CuInSe2. Consider the Cu2Se–In2Se3 isopleth of Fig. 7, for which we have proven that it actually 

constitutes a quasi-binary vertical section of the Cu–In–Se phase diagram. On cooling a melt with the composition 

CuInSe2, the first solid phase is not α-CIS, but a high-temperature phase we denote as δH. On further cooling, αCIS should then form from δH via a phase transformation. However, the melting point maximum of the δH phase 

field lies at a Cu content of 23.5 %, thus considerably off (on the left hand side of α-CIS in Fig. 7) the 25.0% 

composition required for stoichiometric α-CIS (CuInSe2). Therefore, if one tries to grow single-phase α-CIS by 

cooling down a melt with a composition in the vicinity of the stoichiometric composition CuInSe2, one first obtains 

the phase δH, which has the “incorrect” composition Cu23.5In26.0Se50.5, surrounded by more Cu-rich material. On 

further cooling, this material cannot perfectly transform to α-CIS, owing to the limited kinetics. Instead, it will 

transform to α + δR. Consequently, one obtains a high density of phase boundaries. 

Starting out from a composition not equal to but in the vicinity of CuInSe2 does not yield single-phase material, 

either. As revealed in isothermal sections of the Cu–In–Se phase diagram [16-18], α-CIS establishes two-phase 

and three-phase equilibria with nine different phases! This variety of phase equilibria explains the difficulties other 

researchers have encountered when attempting to grow CuInSe2 single crystals. 

1.1 How to Solve the Problem 

Figure 8, taken from [18], represents the most important data for our project: the projection of the liquidus surface

of the Cu–In–Se phase diagram onto the composition triangle. (The liquidus surface separates liquid and solid 

phases in the ternary phase diagram. Figure 8 constitutes a topographical map of the liquidus surface, where 

contour lines indicate the temperature.) Figure 8 confirms in a more comprehensive (three-dimensional) way than 

the isopleth of Fig. 7 that a melt with the composition CuInSe2 does not precipitate α-CIS, but the phase δH – and 

this is the reason why previous attempts to grow single-phase α−CIS from the melt have failed. On the other 

hand, the liquidus surface exhibits four different faces that do enable direct, “primary crystallization” of α-CIS from 

the melt (shaded regions in Fig. 8). However, the corresponding compositions are far away from the composition 

CuInSe2. This explains that direct precipitation of α-CIS from the melt was not discovered “accidentally.” 

Summarizing, the new phase diagram explains the difficulties associated with growing single-phase α-CIS from a 

melt with the composition near CuInSe2, but also suggests melt compositions that do precipitate single-phase αCIS directly, without causing any phase transformation such as δH→α. 

Liquid-Phase Deposition of α-CIS 

According to Fig. 8, it should be possible to grow single-phase α-CIS thin-films by liquid-phase deposition (LPD) – 

by casting a melt with a composition within one of the shaded fields of Fig. 8 over a substrate, and cooling it 

slowly to a temperature somewhat below the liquidus temperature. According to Fig. 8, the melt will then directly

precipitate a film of α-CIS onto the substrate, without further phase transformations. After the film has reached the 

desired thickness, the melt is removed from the substrate. Subsequently, the system is cooled to room 

temperature. 

Since the material grown from the melt will be pure α-CIS and solidifies under near-equilibrium conditions, we 

anticipate substantially lower concentrations of structural defects than in PVD layers, and thus major advantages 

in photovoltaic efficiency (by definition, defects are non-equilibrium features). Indeed, the TEM image of Fig. 9, 

recently obtained from α-CIS material produced during the experiments carried out to determine the new phase 

diagram, reveals coarse-grained, pure α-CIS with a much lower concentration of defects than Figs. 2, 3, and 4. 

NASA/CP—2005-213431 276
Fig. 7: The Cu2Se–In2Se3 isopleth, a quasi-binary vertical section of the new phase diagram. 

NASA/CP—2005-213431 277
Fig. 8: Projection of the liquidus surface of the Cu–In–Se phase diagram onto the composition triangle. 

The four shaded areas indicate surfaces of primary crystallization of α-CIS – thus melt compositions from 

which α-CIS precipitates directly, with no further phase transformations. 

Contrary to the typical PVD-grown material, moreover, LPD-grown α-CIS will be (nearly) 

thermodynamically stable. Since it is single-phase α−CIS, its lifetime in a solar cell will not be limited by 

ongoing diffusional phase transformations. Note that at low temperatures, the phase field of α-CIS, 

shaded in Fig. 7, becomes very narrow. A material outside of this narrow compositional range, as typically 

encountered in PVD-deposited thin-films, is not thermodynamically stable and will undergo diffusional 

phase transformations that may limit its lifetime in a solar cell. This issue is even more important in outer 

space than on the ground, because the cosmic radiation (vacancies by knock-on events) [19] and higher 

operating temperature owing to lack of heat convective heat release (no air) promote diffusion. 

According to the isopleth in Fig. 10, it should be possible to precipitate α-CIS at temperatures below 

400 °C. This opens the perspective to deposit α-CIS thin-films onto temperature-resistant polymer films, 

enabling the fabrication of solar cells with very high-specific-power (power output per unit mass). 

Moreover, LPD will allow the fabrication of larger modules than PVD and can decrease the fabrication 

costs of α-CIS photovoltaic devices, because the deposition rate of LPD is much higher than for PVD, 

and LPD does not require UHV chambers, which are costly to set up and costly to maintain. Experiments 

of the Liquid-Phase Epitaxy Group at the Max-Planck-Institut für Festkörperforschung have shown that 

large (Ø = 4 in) wafers can be efficiently coated by centrifuging the melt by rotating the substrate. 

NASA/CP—2005-213431 278
Fig. 9: TEM image of α-CIS produced during the experiments we carried out to determine the new 

phase diagram. The diffraction pattern (inset) confirms that the region imaged here is a large grain of 

single-phase α−CIS. 

Fig. 10: Se–CuInSe2 isopleth of the Cu–In–Se phase diagram [16-18]. 

NASA/CP—2005-213431 279
Experimental 

For the experimental technique of depositing a solid thin-film from a ternary liquid phase we will draw on 

the experience that has been gained with the deposition of SixGe1-x thin-films from a Si–Ge–Bi melt in the 

Liquid-Phase Epitaxy Group at the Max-Planck-Institut für Festkörperforschung [20-22]. 

Fig. 11 presents a schematic cross-sectional view of a first, basic experimental setup we have installed 

for LPD of α-CIS. The experimental procedure is very similar to the one we have successfully applied for 

DTA (differential thermal analysis) in our study of phase equilibria of the Cu–In–Se system [16-18]. The 

material for the melt, a few grams of high-purity Cu, In, and Se, are encapsulated in a quartz tube. After 

evacuating the tube, it is flushed with Ar at 820 mbar. Encapsulating the material in quartz tubes allows us 

to observe the changes occurring during processing. In particular, by observing the diffusion zone at Cu 

stripes protruding from the Se or In melt we can carefully control the temperature. Owing to the large heat 

of formation and the high vapor pressure of Se, fusion of Cu–In–Se alloys normally causes excessive 

agitation of the components. The Ar-filling of the tubes, however, solves this problem [13]. 

After fusing the metal pellets, the newly formed solid solution is annealed in a muffle furnace for up to one 

hour at 950 °C. To improve the homogeneity of the temperature distribution during this treatment, the 

silica tube is embedded in a large body of heat-resistant steel. Subsequently, the tube is cooled in air and 

the ingot inspected. By mechanical agitation, we detach the newly formed material and any left-over of 

the ingot from the tube walls, put the tube back into the muffle furnace and melt it a second time. 

Depending on the composition, the melting point may vary between 950 and 1020 °C. After the fusing the 

ingot in an induction furnace, the ingot material is removed from the quartz tubes and broken up into 

smaller sizes, each weighing approximately 250 mg. 

To confirm the homogeneity of the ingots, three 

separate pieces of the ingot taken from different 

regions of the ingot are prepared for XRD, DTA and 

XEDS analysis. By performing XEDS measurements 

the stoichiometry of the ingots are confirmed not only 

locally within the ingot but throughout. DTA confirms 

the liquidus projection, shown in Fig. 8, for any given 

composition. 

A commercial tube furnace, equipped with a largediameter (10 cm) quartz tube and a good temperature 

control (± 0.5 K), serves for the actual LPD 

experiments. After the liquidus temperature and the 

composition of the ingot are confirmed, a piece of 

ingot, • 250 mg, is loaded (along with the substrate) 

into another quartz tube, where the actual LPD 

experiment takes place (Fig. 11b). By slowly heating 

the furnace to a temperature between 550 °C and 

1020 °C (depending on the composition of the ingot), a 

homogeneous melt results. Then the quartz tube is 

tilted by 90°, such that the melt wets the substrate. By 

means of an electronic temperature control, we cool 

the melt to a well-defined temperature somewhat 

(ideally about 1°) below the temperature of the 

liquidus surface. Accordingly, the melt will begin to 

precipitate α-CIS. 

In the presence of the substrate, nucleation of α-CIS 

does not begin homogeneously, in the melt, but 

heterogeneously, at the surface of the substrate. 

Subsequently, a closed layer of α-CIS grows and 

Fig. 11: Schematic cross-sectional view of the 

setup for liquid-phase deposition of α-CIS. 

NASA/CP—2005-213431 280
Fig. 12 DTA of ingot I1b revealing a liquidus projection of 

667°C. 

covers the substrate. After the layer has grown to the desired thickness, the quartz tube is tilted back to 

the original position (Fig. 11b) and the system cooled down to room temperature. 

If we choose the composition of the ingot in one of the four regions of the composition triangle where 

Fig. 8 indicates primary crystallization of α-CIS, the above procedure should produce a thin-film of singlephase α-CIS on the substrate. 

First Results 

Ingots were prepared from all 

primary phase fields represented on 

the liquidus projection of the Cu-InSe system (shown in Fig. 8). Many 

initial attempts to produce thin-film αCIS failed due to the inability of the 

melt to wet the substrate. For this 

reason, we used quartz substrates 

sputter-coated with 1 μm of Mo, 

followed by 100 nm of Cu. The 

intention was to increase the surface 

energy of the Mo layer. The Cu layer, 

deposited after the Mo layer without 

breaking vacuum to the chamber, 

served to inhibit surface 

contamination on the Mo. During the 

liquid-phase deposition process, the 

Cu layer diffused into the ingot 

material without altering the ingot 

composition to any significant extent, 

revealing a contaminant-free Mo 

layer. With the incorporation of these 

Cu coated substrates, we were able 

to deposit films from all four 

composition fields suitable for 

primary precipitation of α-CIS, as 

indicated in Fig. 8. 

The liquidus projection of Fig. 8 gave 

us the temperature in which 

nucleation of α-CIS would occur. 

However, to confirm the exact 

liquidus temperature of the ingots 

that were fused together, DTA was employed. Fig. 12 shows the DTA plot for ingot designated I1b. From 

Fig. 12 the liquidus projection is found to be 667 °C. This is in good agreement with the projection given in 

Fig. 8. A small deviation results from making a composition up which lies slightly off one of the 700 °C 

contour lines. 

550 600 650 700

-30

-20

-10

0

10

20

30

Liquidus projection 667 deg C 

676 deg C

658 deg C

DTA / uV

Temperature (deg C)

 (Heating 10 deg/min)

 (Cooling 10 deg/min)

NASA/CP—2005-213431 281
Ingot Ingot 

(weighed) 

Atomic % 

Ingot Measured 

(EDS) 

Atomic % 

Film Measured 

(EDS) 

Atomic % 

I1a 

Cu 12.0 

In 47.5 

Se 40.5 

Cu 8.7 

In 41.4 

Se 49.9 

Cu 4.19 

In 52.69 

Se 43.12 

I1b 

Cu 32.5 

In 32.5 

Se 35.0 

- 

Cu 24.76 

In 27.10 

Se 48.14 

I1c 

Cu 46.0 

In 18.0 

Se 36.0 

- 

Cu 59.50 

In 15.58 

Se 24.92 

I2 

Cu 55.0 

In 38.5 

Se 6.5 

- 

Cu 16.46 

In 41.47 

Se 42.08 

I3 

Cu 46.0 

In 4.0 

Se 50.0 

Cu 45.40 

In 2.66 

Se 51.94 

Cu 7.90 

In 1.33 

Se 90.77 

I4 

Cu 2.85 

In 3.55 

Se 93.6 

Cu 4.46 

In 2.81 

Se 92.73 

Cu 5.74 

In 0.80 

Se 93.46 

 Table 1 is a compilation of compositions, 

 given in atomic percentages, for all ingots and films. 

Fig. 13 XRD pattern of the LPD α-CIS from 

Ingot I1b. 

Films from all primary 

phase fields of the liquidus 

projection were deposited. 

Several of the deposited 

films simply exhibited the 

composition of the ingot 

material – especially those 

from the uppermost, Serich composition field. The 

fact that no change in 

composition was observed 

indicates that the 

precipitation of singlephase α-CIS was not 

successful with these 

particular fields. However, 

several deposited films 

contained compositions 

that did reflect 

considerable compositional 

changes, indicating the 

precipitation of a 

thermodynamically stable 

phase. A summary of 

compositions of the ingots 

made (measured by an 

analytical balance and 

confirmed by XEDS) and 

the films (also measured 

by XEDS) are given in 

Table 1. From Table 1, the 

ingot designated I1b 

initially had a composition 

corresponding to 1:1:1 

Cu:In:Se and resulted in 

the deposition of a film 

with a stoichiometry of 

1:1:2. 

To verify that that the 

stoichiometry of the film 

made from ingot I1b did 

correspond to the α-CIS 

phase, and not to a sum of 

several phases adding up 

to a 1:1:2 composition, 

XRD was utilized. 

Figure 13 is an XRD 

pattern from one of the 

films grown from the I1b 

ingot. The presence of a 

dominant peak at 26.6° 

confirms the presence of 

the (112) plane. This peak 

is the dominant peak 

within the spectrum as it is 

20 30 40 50 60 70 80

-1000

0

1000

2000

3000

4000

5000

6000

7000

8000

Counts

2 Theta

112

Mo

Mo

Mo

NASA/CP—2005-213431 282
one the two preferred growth planes of α-CIS along with the (220) plane (not predominately present). 

Conclusion 

Based on the results described in the previous section, we believe that it is possible to fabricate singlephase α-CuInSe2 layers with high structural and electronic quality by liquid-phase deposition (LPD). 

Some experimental problems of this method have been identified, such as insufficient wetting of the 

substrate, high viscosity of the melt below the liquidus temperature, and the necessity of a very precise 

temperature control. However it seems that these problems can be overcome. As the experimental 

techniques are further developed, we hope that LPD will become an attractive method for fabricating highquality α-CuInSe2 thin films at low cost. 

 References 

1. A. Rockett, F. Abou-Elfotouh, D. Albin, M. Bode, J. Ermer, R. Klenk, T. Lommasson, T. W. F. 

Russell, R. D. Tomlinson, J. Tuttle, L. Stolt, T. Walter and T. M. Peterson: Structure and chemistry of 

CuInSe2 for solar cell technology: current understanding and recommendations, Thin Solid Films 237 

(1994) 1-11. 

2. D. Suri, K. Nagpal and G. Chadha, J. Appl. Cryst. 22 (1989) 578-583. 

3. W. Hönle, G. Kühn and U.-C. Boehnke, Cryst. Res. Technol. 23 (1988) 1347-1354. 

4. J. R. Sites, J. E. Granata and J. F. Hiltner: Losses due to polycrystallinity in thin-film solar cells, 

Solar Energy Materials and Solar Cells 55 (1998) 43-50 

5. NREL, www.nrel.gov. 

6. J. R. Sites: Separation of voltage loss mechanisms in polycrystalline solar cells, in: IEEE 

Photovoltaic Specialists Conference 06160-8371 (1988), IEEE, USA, p. 1604-1607. 

7. I. Dirnstorfer, W. Burkhardt, W. Kriegseis, I. Osterreicher, H. Alves, D. M. Hofmann, O. Ka, A. Polity, 

B. K. Meyer and D. Braunger: Annealing studies on CuIn(Ga)Se2: the influence of gallium, Thin 

Solid Films 361-362 (2000) 400-405. 

8. S. Kuranouchi and A. Yoshida: Annealing effects of CuInSe2 films prepared by pulsed laser 

deposition, Thin Solid Films 343-344 (1999) 123-126. 

9. S. Kuranouchi and A. Yoshida: Annealing effects of CuInSe2 films prepared by pulsed laser 

deposition, Thin Solid Films 343-344 (1999) 123-126. 

10. S. Niki, I. Kim, P. J. Fons, H. Shibata, A. Yamada, H. Oyanagi, T. Kurafuji, S. Chichibu and H. 

Nakanishi: Effects of annealing on CuInSe2 films grown by molecular beam epitaxy, Solar Energy 

Materials and Solar Cells 49 (1997) 319-326. 

11. C. R. Schwerdtfeger, T. F. Ciszek and R. Noufi, 1995, High-Pressure, Liquid-Encapsulated 

Directional Solidification of Copper-Deficient Copper Indium Selenides. Crystal Growth of Novel 

Electronic Materials: Proceedings of the 97th Annual Meeting of the American Ceramic Society 

(Cincinnati, Ohio: American Ceramic Society), p. 95-103. 

NASA/CP—2005-213431 283
12. C. R. Schwerdtfeger: Crystal Growth of Copper-Deficient Copper-Indium-Selenides Using ComputerControlled High-Pressure Liquid-Encapsulated Directional Solidification, Ph. D. Thesis, National 

Renewable Energy Laboratory (1995). 

13. M. Hornung: Züchtung von Kristallen im System Cu–In–Se und deren mikrostrukturelle 

Charakterisierung, Dissertation, Universität Freiburg (1996). 

14. T. Haalboom: Mikrostruktur und Konstitution von Kupfer-Indium-Selen-Mischkristallen für Solarzellen 

(Microstructure and phase equilibria of copper-indium-selenium solid solutions for solar cells), 

Doctoral Thesis, Universität Stuttgart (1999). 

15. V. Lyahovitskaya, S. Richter, F. Frolow, L. Kaplan, Y. Manassen, K. Gartsman and D. Cahen: 

Growth of single CuInSe2 crystals by the traveling heater method and their characterization, Journal 

of Crystal Growth 197 (1999) 177–185. 

16. T. Gödecke, T. Haalboom and F. Ernst: Phase Equilibria of Cu-In-Se. I. Stable States and NonEquilibrium States of the In2Se3–Cu2Se Subsystem, Zeitschrift für Metallkunde 91 (2000) 622-634. 

17. T. Gödecke, T. Haalboom and F. Ernst: Phase Equilibria of Cu-In-Se. II. The Cu–Cu2Se–In2Se3–In 

Subsystem, Zeitschrift für Metallkunde 91 (2000) 635-650. 

18. T. Gödecke, T. Haalboom and F. Ernst: Phase Equilibria of Cu-In-Se. III. The In2Se3–Se–Cu2Se 

Subsystem, Zeitschrift für Metallkunde 91 (2000) 651-662. 

19. A. F. W. Willoughby: The control of radiation resistance in space solar cells, International Journal of 

Electronics 76 (1994) 865-882 

20. H.-P. Trah: Flüssigphasen-Heteroepitaxie auf Silizium-Substraten, Ph. D. Thesis, Universität 

Freiburg (1988). 

21. V. I. Tagirov, N. F. Gakhramanov, A. G. Guseinov, F. M. Aliev and G. G. Guseinov, Soviet Journal of 

Physics and Crystallography 25 (1980) 237-239. 

22. H.-P. Trah: Liquid phase epitaxy in the ternary system Si-Ge-Bi, Journal of Crystal Growth 102 

(1990) 175-182. 

NASA/CP—2005-213431 284
RECENT PROGRESS IN CuInS2 THIN FILM SOLAR CELL RESEARCH AT NASA GLENN

M. H.-C. Jin, K. K. Banger, and C. V. Kelly

Ohio Aerospace Institute, Brook Park, OH 44142 

J. H. Scofield

Oberlin College, Oberlin, OH 44074

J. S. McNatt, J. E. Dickman, and A. F. Hepp

NASA Glenn Research Center, Cleveland, OH 44135

INTRODUCTION

The National Aeronautics and Space Administration (NASA) is interested in developing low-cost highly efficient

solar cells on light-weight flexible substrates, which will ultimately lower the mass-specific power (W/kg) of the cell

allowing extra payload for missions in space as well as cost reduction (1). In addition, thin film cells are

anticipated to have greater resistance to radiation damage in space, prolonging their lifetime (2). The flexibility of

the substrate has the added benefit of enabling roll-to-roll processing (3).

The first major thin film solar cell was the “CdS solar cell” – a heterojunction between p-type CuxS and n-type CdS

(4). The research on CdS cells started in the late 1950s and the efficiency in the laboratory was up to about 10 %

in the 1980s (5). Today, three different thin film materials are leading the field. They include amorphous Si, CdTe,

and Cu(In,Ga)Se2 (CIGS) (6). The best thin film solar cell efficiency of 19.2 % was recently set by CIGS on glass

(7). Typical module efficiencies, however, remain below 15 % (8).

Several groups have attempted to fabricate thin film solar cells on metal foils and polymers. Efficiencies up to

12.8 % on polyimide (9) and 17.1 % on stainless steel (10) have been reported. Polymer substrates cannot

tolerate temperatures as high as those allowed by metal foils. Cell fabrication and absorber layer synthesis on

polymer substrates must be accomplished at temperatures below 450 °C (11). This low-temperature constraint

introduces challenges for achieving high-quality absorber layers and high-efficiency solar cells. To facilitate low

temperature deposition, single-source organometallic precursors with low decomposition temperatures (< 250 °C)

were synthesized and used in an aerosol-assisted chemical vapor deposition (AACVD) setup to create

chalcopyrite thin films (12,13). This paper gives a brief description of the first systematic effort at NASA Glenn

Research Center (GRC) to fabricate CuInS2 thin film solar cells.

EXPERIMENTAL

As previously reported, CuInS2 SSP, (PPh3)2CuIn(SEt)4 was synthesized following a modified version (12) of the

procedure reported by Kanatzidis (14). Because of their sensitivity to O2 and H2O, all reagents were handled in an

argon-filled glove box. The precursor was dissolved in toluene and the solution was delivered to the heated

substrate in the form of sprayed aerosol in a horizontal, atmospheric-pressure, hot-wall AACVD reactor equipped

with a plate-type ultrasonic nebulizer (Sonaer Ultrasonics, 2.5 MHz). A detailed description of the film growth

process with a schematic of the reactor can be found elsewhere (13,15). For this study, (112)-oriented CuInS2

films were deposited using 0.01 M precursor solution under 4 L/min of Ar flow. The evaporation and hot zone

temperatures were 120 °C and 395 °C respectively. Typical film thickness was around 0.7 μm and Cu/In ratio was

1.0 ± 0.1. The detailed characterizations of the films were reported previously (13,16).

The solar cell structure used in this study was Al/ZnO:F/CdS/CuInS2/Mo/glass. Prior to CuInS2 film growth,

molybdenum (bottom metal contact) was deposited on glass substrates (Fisher catalogue no. 12-550A) at room

temperature using radio-frequency (RF) magnetron sputtering. After CuInS2 film was grown by AACVD, the film

NASA/CP—2005-213431 285
was etched in 1.5 M KCN solution for 1 min at room temperature prior to the CdS chemical bath deposition (CBD).

CBD was performed with a solution consisting of 0.001M CdSO4, 1.5M NH4OH, and 0.0075M thiourea. The bath

was heated to 70 °C then the sample was immersed. After the CBD, the particulates on the surface of the sample

were removed in an ultrasonic bath. Subsequently a ZnO:F layer was deposited from a 6.5” (16.5 cm)-diameter Fdoped ZnO target in an RF sputter system. The sputter gas was pure Ar for both Mo and ZnO:F deposition.

Following ZnO:F growth, a 0.2 μm thick Al top metal contact was deposited by thermal evaporation through a

shadow mask. The cell area was defined by mechanical scribing after Al deposition with a typical cell area of 0.4

cm

2

. A solar simulator at GRC (Spectrolab X-25 Mark II) was used to characterize the solar cells under AM0

radiation, calibrated using either a standard GaAs single-junction cell or a Si single-junction cell.

In order to validate the fabrication procedure, a chalcopyrite thin film deposited by physical vapor deposition (PVD)

on Mo-coated glass substrate was acquired and a solar cell was fabricated through the same procedure

described. In an effort to improve the cell performance, GRC collaborated with Oberlin College to try a double

window layer scheme (n-type ZnO/intrinsic ZnO) after CdS deposition. In addition, a Ni/Al double metal top

contact was also used instead of a single Al contact. The intrinsic and the Al-doped ZnO layers were deposited

using a reactive DC magnetron sputter process (Sloan) in a mixture of O2 and Ar gases. The Ni was evaporated

from an alumina-coated tungsten boat and the Al from a single-hearth e-gun (Sloan) mounted in a 24-in. diffusionpumped bell jar. The GRC and Oberlin cell structures are illustrated in Figure 1.

Figure 1. Schematic of the solar cell structures made at GRC and Oberlin College.

The film thickness was determined using a profiler (Sloan Dektak IIA) and the optical transmittance was measured

by a UV/VIS spectrophotometer (Perkin-Elmer Lamda-19). The electrical measurements were performed with a

four-point probe system (Bio-Rad HL5500PC) operated in the Van der Pauw configuration. For the CdS film, Xray diffraction (XRD) (PANalytical X’Pert Pro) was used to identify phases and glancing-angle X-ray diffraction

(GAXRD) was further performed to probe phases in the film and on the surface of the film separately.

RESULTS AND DISCUSSION

The surface resistance of Mo bottom contact was about 0.5 ohm/sq. with a typical thickness of 0.7 μm. XRD

confirmed that the Mo had a body-centered cubic phase with a (110) texture showing the strongest diffraction at 2θ

= 40.5°. The optical transmittance of CBD-grown CdS was 65 % at a photon energy of 1.5 eV and it had a strong

signature of cubic phase with CdO as an impurity phase (Figure 2). GAXRD confirmed that the CdO phase was

present both in the bulk and the surface of the film (not shown). There have been many reports regarding

secondary phase formation including CdO in the CdS CBD (17,18), but no XRD was reported to support the CdO

formation in as-deposited CdS prepared by CBD. A typical 1 μm-thick GRC n-type ZnO:F layer had a sheet

resistance of about 200 ohm/sq and a transmittance of 85 % at a photon energy of 1.5 eV. In comparison, a 560

nm thick n-type ZnO layer deposited at Oberlin College had a sheet resistance of about 46 ohm/sq. with a

transmittance of about 70 % at a photon energy of 1.5 eV. The transmittance of the 70 nm thick intrinsic ZnO was

NASA/CP—2005-213431 286
over 90 % at photon energy of 1.5 eV and the sheet resistance of the double window layer, n-ZnO/i-ZnO was

about 200 ohm/sq. Typical thickness of single Al top contact deposited at GRC was about 700 nm and the sheet

resistance of the double top contact layer (2 μm Al/50 nm Ni) made was about 0.18 ohm/sq.

Figure 2. XRD of CdS film deposited by chemical bath deposition.

Figure 3 shows the characteristics of the reference solar cells completed with both the GRC and Oberlin College

window/contact layers. The chalopyrite absorber layer deposited by PVD was used for this validation process of

the device fabrication at GRC. The AM0 efficiencies of 5.3 % and 6.7 % were obtained with minimum device

optimization. The use of intrinsic ZnO and Ni/Al top contact layers, which is known to improve the quality of this

type of solar cell (6), lowered the series resistance and improved the fill factor (FF) as shown in the Figure 3.

Although the short circuit current (Isc) was comparable to that of any other reported chalcopyrite solar cells (7,8),

open circuit voltage (Voc) and FF have to be further improved. Further optimization of the fabrication process will

be performed in the future.

Figure 3. Characteristics of the reference solar cells made at GRC and Oberlin College.

A solar cell with a structure of Al/ZnO:F/CdS/CuInS2/Mo/glass was fabricated with a CuInS2 thin film deposited by

AACVD at GRC and its output characteristics were shown in Figure 4. The cell had both low shunt resistance and

high series resistance. The intrinsic ZnO and Ni/Al top contact will be incorporated for further improvement in the

future. In addition, further increase of CuInS2 film thickness from 0.7 μm should minimize the optical loss. It

should be noted that the best efficiency reported for a solar cell prepared from AACVD-made CuInS2 was 2.66 %

(20). However, the cell had a superstrate structure – Cd(Zn)S was deposited on the glass substrate first and light

NASA/CP—2005-213431 287
passes through from the glass side - and there was no report of a working device with a typical chalcopyrite solar

cell structure, CdS/CuInS2 except for the authors’ previous results (13).

Figure 4. Characteristics of the CuInS2 solar cell made at NASA GRC.

CONCLUSION

Chalcopyrite solar cell fabrication process has been established at NASA GRC. A reference solar cell with a cell

structure of Al/ZnO:F/CdS/CuInS2/Mo/glass was fabricated and an efficiency of 5.3 % under AM0 illumination was

achieved. Further improvement was made by incorporating an intrinsic ZnO layer and Al/Ni top metal contact at

Oberlin College, increasing the efficiency up to 6.7 %. The series resistance and fill factor were clearly improved.

In addition, GRC has fabricated CuInS2 thin films solar cells prepared from AACVD using the single source

precursor, (PPh3)2CuIn(SEt)4. Although the solar cell performance still needs to be improved, Isc and FF were

better than results reported previously (13). The major loss was due to the low Voc and it remains as a challenge,

which has been observed in solar cells prepared with AACVD deposited films (19,20).

REFERENCES

(1) D. J. Hoffman, T. W. Kerslake, A. F. Hepp, M. K. Jacobs, and D. Ponnusamy, “Thin-Film Photovoltaic Solar

Array Parametric Assessment,” Proc. of the AIAA 35

th

Intersociety Energy Conversion Engineering

Conference, pp. 670-680, Las Vegas, NV, 2000.

(2) V. K. Kapur, M. Fisher, and R. Roe, “Fabrication of Light Weight Flexible CIGS Solar Cells for Space Power

Applications,” Mat. Res. Soc. Symp. Proc., 668, pp. H3.5.1-H3.5.6, 2001.

(3) G. M. Hanket, U. P. Singh, E. Eser, W. N. Shafaran, and R. W. Birkmire, “Pilot-scale Manufacture of

Cu(InGa)Se2 Films on a Flexible Polymer Substrate,” Proc. of the 29

th

IEEE Photovoltaic Specialists

Conference, pp. 567-570, New Orleans, LA, 2002.

(4) A. Rothwarf and A. M. Barnett, “Design analysis of the thin-film CdS-Cu2S solar cell,”IEEE Trans. Electron

Devices, ED-24, pp. 381-387, 1977.

(5) R. B. Hall, R. W. Birkmire, J. E. Phillips, and J. D. Meakin, “Thin-film polycrystalline Cu2S/Cd1-xZnxS solar

cells of 10% efficiency,” Appl. Phys. Lett., 38, pp. 925-926, 1981.

(6) M. D. Archer and R. Hill, “Clean Electricity from Photovoltaics,” Imperial College Press, London, 2001.

(7) K. Ramanathan, M. A. Contreras, C. L. Perkins, S. Asher, F. S. Hasoon, J. Keane, D. Young, M. Romero,

W. Metzger, R. Noufi, J. Ward, and A. Duda, “Properties of 19.2% Efficiency ZnO/CdS/CuInGaSe2 Thin-film

Solar Cells,” Prog. Photovolt: Res. Appl., 11, pp. 225-230, 2003.

(8) K. Kushiya, “Progress in Large-area Cu(InGa)Se2-based Thin-film Modules with the Efficiency of over 13%,”

Proc. of the 3

rd World Conference on Photovoltaic Energy Conversion, 2PL-C1-02, Osaka, Japan, 2003.

NASA/CP—2005-213431 288
(9) A. N. Tiwari, M. Krejci, F. J. Haug, and H. Zogg, “12.8% Efficiency Cu(In,Ga)Se2 solar cell on a flexible

polymer sheet,” Prog. Photovolt: Res. Appl., 7, pp. 393-397, 1999.

(10) M. Contreras, B. Egas, K. Ramanatan, J. Hiltner, F. Hasoon, and R. Noufi, “Progress toward 20% efficiency

in Cu(In,Ga)Se2 polycrystalline thin-film solar cells,” Prog. Photovolt: Res. Appl., 7, pp. 311-316, 1999.

(11) F. Kessler, K. Herz, M. Powalla, M. Hartmann, M. Schmidt, A. Jasenek, and H. W. Schock, “Flexible and

monolithically integrated CIGS-Modules,” Mat. Res. Soc. Symp. Proc., 668, pp. H3.6.1-H3.6.6, 2001.

(12) K. K. Banger, J. Cowen, and A. F. Hepp, “Synthesis and Characterization of the First Liquid Single-Source

Precursors for the Deposition of Ternary Chalcopyrite (CuInS2) Thin Film Materials” Chem. Mater., 13, pp.

3827-3829, 2001.

(13) J. D. Harris, K. K. Banger, D. A. Scheiman, M. A. Smith, M. H.-C. Jin, and A. F. Hepp, “Characterization of

CuInS2 films prepared by atmospheric pressure spray chemical vapor deposition,” Mat. Sci. & Engr., B98,

pp. 150-155, 2003.

(14) W. Hirpo, S. Dhingra, A. C. Sutorik, and M. G. Kanatzidis, “Synthesis of Mixed Copper-Indium

Chalcogenolates. Single Source Precursors for the Photovoltaic Materials CuInQ2 (Q = S, Se),” J. Am.

Chem. Soc., 115, pp. 1597-1599, 1993.

(15) J. A. Hollingsworth, K. K. Banger, M. H.-C. Jin, J. D. Harris, J. E. Cowen, E. W. Bohannan, J. A. Switzer, W.

E. Buhro, and A. F. Hepp, “Single source precursors for fabrication of I-III-VI2 thin-film solar cells via spray

CVD,” Thin Solid Films, 431-432, pp. 63-67, 1993.

(16) M. H.-C. Jin, K. K. Banger, J. D. Harris, and A. F. Hepp, “The Effect of Film Composition on the Texture and

Grain Size of CuInS2 Prepared by Chemical Spray Pyrolysis,” Mat. Res. Soc. Symp. Proc., 763, pp. B8.23.1-

B8.23.6, 2003.

(17) P. J. Sebastian and H. Hu, “Identification of the Impurity Phase in Chemically Deposited CdS Thin Films,”

Adv. Mater. Opt. Electron., 4, pp. 407-412, 1994.

(18) T. Nakada, H. Fukuda, A. Kunioka, and S.Niki, “XPS analysis of chemically deposited CdS films on

polycrystalline CIS surface and glass substrates,” Proc. of the 13

th

European Photovoltaic Solar Energy

Conference, pp.1597-1600, Nice, France, 1995.

(19) P. R. Ram, R. Thangaraj, A. K. Sharma, and O. P. Agnihotri, “Totally sprayed CuInSe2/Cd(Zn)S and

CuInS2/Cd(Zn)S solar cells,” Solar Cells, 14, pp. 123-131, 1985.

(20) O. Kijatkina, M. Krunks, A. Mere, B. Mahrov, and L. Dloczik, “CuInS2 sprayed films on different metal oxide

underlayers,” Thin Solid Films, 431-432, pp. 105-109, 2003.

NASA/CP—2005-213431 289
A SILICONE/POSS COATING FOR THIN FILM SOLAR CELLS AND CONCENTRATORS 

Henry W. Brandhorst1

, Mark J. O’Neill2 and Joseph D. Lichtenhan3

1

Space Research Institute, Auburn University, 2ENTECH, Inc., 3Hybrid Plastics, Inc. 

Introduction:

Significant effort has been devoted to the development of thin film solar cells for use in space. One of the 

reasons for this enthusiasm is based their exceptional resistance to bombardment by electrons and protons of 

various energies in the laboratory. This resistance is due in part to their structure and in part to their inherent 

performance mechanisms. Unfortunately, laboratory testing is limited to a few energies of both electrons and 

protons. The energies and doses do not correspond well to those found in space. In space of course there is a 

spectrum of particle energy of both protons and electrons as well as a wide range of fluences as functions of 

particle and energy. In general, the test results to date do not include very low energy protons due to facility 

limitations. Because all solar cells including thin film cells are seriously affected by ionizing radiation that is 

absorbed in their junction region it is important to shield against these highly damaging particles. The protons 

are the most damaging due to their greater mass and they leave a large ionization trail. 

In the past, many have believed that the antireflection coating and/or the transparent conducting oxide layers on 

the surface of thin film cells would provide protection against this spectrum of ionizing radiation. However, from 

many laboratory tests using very low energy particles and from the earliest tests of unshielded GaAs cells in 

space, it is clear that additional protection will be needed if thin film cells are to withstand the rigors of the earth 

orbital environment. The potential solution of using the traditional rigid cover glass is not desirable for at least 

two reasons: they are heavy and they are not flexible. Both these factors will reduce or eliminate the advantage 

of thin film cells in space compared to crystalline cells. The desired coating must be flexible, it must be resistant 

to atomic oxygen for low earth orbits below about 700 km and it must be resistant to ultraviolet radiation for all 

earth and near-sun orbits and, of course, it must withstand the damaging effects of space radiation. 

These same attributes of the protective layer are directly applicable to the Stretched Lens Array (SLA) 

developed by ENTECH, Inc. and Able Engineering. The SLA has demonstrated array specific masses of 180 

W/kg for arrays sized to 7 kW and the potential exists to increase that impressive value by at least twofold. 

Because it is a concentrator system, the cells can be protected against radiation with a cover glass. However, 

the stretched lens must also be flexible and still survive the damaging effects of ultraviolet light, atomic oxygen 

and space radiation. The lens is made from DC 93-500 silicone material which is also used to secure cover 

glasses to the cells. 

Building on this experience, one material that has many of the desired attributes of a universal protective coating 

is the Dow Corning silicone resin DC 93-500. Of all the potential optical plastics, it appears to be the most 

suitable for use in space. As noted above, DC 93-500 has been extensively used to attach cover glasses to 

crystalline solar cells and has worked exceptionally well over the years. It is flexible and generally resistant to 

electrons, protons and ultraviolet (UV and VUV) radiation; although a VUV-rejection coating or VUV-absorbing 

ceria-doped cover glass may be required for long mission durations. It can also be applied in a thin coating (< 25 

μm) by conventional liquid coating processes. 

Unfortunately, when exposed to atomic oxygen (AO) DC 93-500 tends to develop a frosty surface. Such frosting 

can lead to a loss of light transmitted into the cells and destroy the essential clarity needed for a concentrator. 

The damage also erodes the surface and hence to an ultimate failure as a radiation protective coating. This 

NASA/CP—2005-213431 290
Figure 1: Anatomy of a POSS®

 nanostructured® chemical

frosting would be exceptionally harmful for lenses used in a concentrator application. Therefore, solving the 

protection problem for thin film solar cells may also provide the solution for the stretched lens and vice versa. 

A New Option:

A new class of materials has recently been discovered during the search for improvements in optical plastics. 

This class of materials may eliminate the limitation of DC 93-500. These new polymeric building block materials 

are called “polyhedral oligomeric silsequioxalanes” and were developed by the Air Force Research Laboratories 

at Edwards Air Force Base in 1998. The technology is exclusively manufactured by Hybrid Plastics Inc. 

(Fountain Valley, CA) and is being marketed under the name of POSS®

 nanocomposites. This technology has 

several significant advantages that are relevant to solar cell use. The glass-like composition of POSS provides 

excellent resistance to UV and much reduces or eliminates degradation by atomic oxygen. POSS nanobuilding 

blocks can be incorporated into all known plastics using conventional polymerization or compounding 

techniques to afford tailorable optically transparent materials with and entirely new performance levels. 

POSS®

 building blocks contain hybrid (organic-inorganic) compositions in which the internal frameworks are 

comprised primarily of inorganic silicon-oxygen bonds. As shown in figure 1, the exterior of the nanostructure is 

covered by both reactive and nonreactive organic functionalities (–R) 

which ensure both compatibility and 

tailorability of the nanostructure with 

organic polymers. The resulting 

nanoscopic chemicals have low density, 

range in diameter from 0.5 to 3 nm and 

can be tailored through variation in the –

R groups and the size of the nanocage 

reported by Gilman1

 and more 

extensively by Gonzalez2

, the 

molecularly dispersed POSS readily 

forms a passivating silica layer when 

attacked by atomic oxygen. This layer in 

turn protects the virgin material from 

degradation. Furthermore, the silica-like composition of POSS provides enhanced UV and VUV resistance. The 

UV and VUV resistance provided by POSS can even be enhanced further through replacement of metals like 

Ce for the silicon atoms at the vertices in the nanocage. 

Previous Experimental Results:

Siloxanes (e.g. DC 93-500) are generally considered stable under photo-aging or photo-oxidation conditions. 

Although they can protect underlying material from oxidation, they do not significantly absorb incident UV 

radiation, especially if they are alkyl-substituted siloxanes. A related class of materials that are more highly 

oxidized and thereby more stable to oxidation and UV are silsesquioxanes. Previous research has shown that 

chemical manipulation of the organic groups surrounding the POSS cage enables dispersion of POSS 

nanostructures throughout the a polymer matrix at high POSS loadings via blending and co-polymerization 

techniques.3

 The 1nm level of dispersion attainable with POSS is controlled by the free energy of mixing 

equation ∆G = ∆H-T∆S.4

 Where the enthalpic ∆H term for POSS dispersion is controlled by the chemical 

compatibility of the R groups on the POSS nanostructure with the resin along with associated polymer-POSS 

interactions. The entropic ∆S term for POSS dispersion is always favorable for dispersion because of the 

monodisperse cage size of POSS. Note this is not the case for the dispersion of a polymeric resin because it is 

already a dispersion of molecules. 

POSS based siloxane copolymers have been explored for use as LEO-resistant coatings (Figure 2). 

NASA/CP—2005-213431 291
Figure 2: POSS-siloxane copolymers. 

Evaluation of the resistance of the POSS-siloxane copolymer to a simulated LEO environment was conducted 

under exposure to atomic oxygen with and without far-UV radiation. XPS analysis revealed that the surface of 

the copolymer samples became silicon- and oxygen-rich after exposure to atomic oxygen with and without farUV radiation. The results shown in Table 1 indicate that the POSS-siloxane copolymer has excellent resistance 

to the simulated LEO environment and actually gained weight during the exposures. The slight increase of 

weight is attributed to the loss of surface organics and formation of a passivating silica surface layer. Combining 

far-UV radiation with atomic oxygen does not appear to significantly change the results. 

Table 1: Exposure conditions and mass changes before and after exposure 

Exposure Conditions Far-UV Intensity 

(mW/cm2

) 

Total Fluence 

(0 atom/cm2) 

Net Mass Change 

(mg/cm2) 

Atomic oxygen 0 4.3 x 1021 + 0.01 

Atomic oxygen and far-UV 15 (7 earth suns) 4.3 x 1021 + 0.02 

A comparative SEM analysis of the exposed and unexposed POSS-siloxane copolymer samples, shown in 

Figure 3, reveals an interesting change in the surface morphology. The surface defects or microcracks in Figure 

4(a), resulting from sample preparation, were completely “healed” by exposure to the atomic oxygen beam [Fig. 

4 (b) 1. This is in contrast to the results from the Long Duration Exposure Facility (LDEF) experiment where all 

siloxane-coating samples showed extensive microcracking, frosting, and oxidation of underlying organic 

polymeric material due to penetration of A0 through the microcracked coating. We believe that this self-healing 

nature is a direct result of POSS compatibility and dispersion throughout the polymer matrix. This annealing 

phenomenon may be from surface heating, both from the heat of reactions and from absorption of the kinetic 

energy (2-3 eV) of the atomic oxygen. Such behavior provides a significant advantage for these materials since 

microcracking exposes a coated surface to the LEO environment and to accelerated degradation and 

contributes to loss of long-term durability for composites used in space. If the glassy silica layer erodes or 

suffers a microdefect, it would quickly reform due to the uniform POSS dispersion. Furthermore retaining optical 

characteristics is a great advantage – even a necessity – for solar arrays. 

Figure 3: SEM of (a) unexposed and (b) exposed POSS siloxane copolymer surfaces. The simulated LEO exposure has 

“healed” the microcracks present initially in the POSS-siloxane sample. 

NASA/CP—2005-213431 292
Approach:

In this new effort, we plan to improve the UV/VUV and AO resistance of DC 93-500 for both thin film solar cell 

and concentrator applications using two approaches. First we will incorporate an all-silicon containing POSS 

material to confirm a modest improvement in UV, VUV, and AO resistance with adequate flexibility and 

secondly, we will incorporate a Ce-containing POSS to confirm leapfrog-type advancements in VUV and UV 

stability. The above approaches will maintain the current manufacturing protocols used for DC 93-500. Typical 

materials are shown in Figure 4. 

Si

O

Si O

Si

O

Si

O Si

O

Si

O

Si

O Si

O

Si

Si O

O

O Si

O

Si

R

R

R

O

O

R

R

R

O O

O

O

R

R

R

R

R

R

Si

Si

O

O

Si

Si

Si

Si

O

O

O

O

Si O

Si

O

O

O

O O

R R

R

R

R

R

R R

 

Si

Si

O

O

Si

Si

Si

Si

O

O

O

O

Si

O

Si

O

O

O

O

O

R R

R

R

R

R

R R

Si

Si

O

O

Si

Si

Si

Si

O

O

O

O

Si

O

Si

O

O

O

O

O

R R

R

R

R

R

R R

O

O

Ce

 a b 

Figure 4: (a) 1.5nm diameter ocatmeric and dodecameric all Silicon POSS additives bearing aromatic R groups suitable for 

UV absorption (b) A Cerium containing POSS additives for maximum UV absorption. 

Incorporating POSS into DC 93-500 will be conducted using a narrow range of POSS systems that show 

particular promise and compatibility with this resin. Both silicon-only and Ce-containing POSS systems will be 

explored. In this effort, ENTECH, Inc. will conduct all molding, and testing work to determine the suitability of 

these materials for use in the SLA concentrator lens materials. Hybrid Plastics, Inc. will conduct all formulation 

development and fundamental materials characterization. The Space Research Institute will evaluate the use as 

coatings as thin films. 

Variables to be included in the compounding approach will be the composition and the loading level for the 

POSS systems. These will be evaluated in DC 93-500 in the form of films, test plaques or bars or other 

appropriately sized specimens for screening and characterization relative to the baseline resins. Specific 

characterization testing will be conducted relative to clarity, thermal transition heat distortion, abrasion 

resistance and moldability. Initial screening will be based on optical transparency. A down-select to at least 

three baseline formulations will be conducted and these will be evaluated with respect to the remaining criteria. 

Conclusions:

We are embarking on an entirely new approach for protecting solar cell arrays against UV and VUV radiation 

and atomic oxygen. If successful, this will lead to a major change in the solar array industry. 

References:

1. J.W. Gilman, D.S. Schlitzer and J.D. Lichtenhan, “Low Earth Orbit Resistant Siloxane Copolymers”, J. 

Appl. Polymer Sci. 60, 591-596 (1996) 

2. R.I. Gonzalez S.H. Phillips and G.B. Holland, “In Situ Oxygen-Atom Erosion Study of Polyhedral 

Oligomeric Silsesquioxane-Siloxane Copolymer”, J. Spacecraft and Rockets, 37, #4, 463-467, (2000) 

3. J. D. Lichtenhan, J. J. Schwab, W. A. Reinerth “Nanostructured Chemicals: A New Era in Chemical 

Technology,” Chemical Innovation January 2001 pp 3-5. 

(http://pubs.acs.org/subscribe/journals/ci/31/special/0101lead.html) 

4. W. Zhang, B. X. Fu, Y. Seo, E. Schrag, B. Hsiao, P. T. Mather, N-L. Yang, D. Xu, H. Ade, M. 

Rafailovich, and J. Sokolov “Effect of Methyl Methacrylate/Polyhedral Oligomeric Silsesquioxane 

Random Copolymers in Compatibilization of Polystyrene and Poly(methyl methacrylate) Blends”,

Macromolecules 35, 8029-8038 (2002). 

NASA/CP—2005-213431 293
Future Directions for Thin Films in Space Workshop at SPRAT XVIII

Aloysius F. Hepp and John E. Dickman

NASA Glenn Research Center

Photovoltaic and Space Environments Branch

Cleveland, OH 44135

Overview

For approximately ten years, the SPRAT conference series at NASA Glenn (formerly Lewis)

Research Center has devoted a workshop to the topic of thin-film solar cell technology and its

potential for space applications. While thin-film materials have been investigated for a number

of years, including copper sulfide research at NASA in the 1960’s, there has been a re-birth of

interest in this class of materials for space applications in the past dozen years or so. There are

several reasons for this renaissance. An important contributor is efficiency improvements

beyond ten percent. Another contributor is the increase in interest represented by funding

opportunities by NASA and several agencies in the U.S. Department of Defense (Missile

Defense Agency, Air Force, and DARPA). Finally, there have been several intriguing missions

identified through various means, these include: station-keeping for high-altitude airships, space

solar power, planetary surface power, and solar electric propulsion. To aid in leading the

discussion for this workshop, a series of seven questions were posed. These are reproduced

below as well as a summary of key points and conclusions from the workshop as well as an

attendees list and results of an informal poll related to long-term potential of thin films for space.

Workshop Discussion Questions

1 – Will thin film PV (for space) always be the technology of the future?

2 – Polycrystalline thin film materials co-exist with polymer? Quantum dots? Poly III-V?

3 – How do metal and polymer substrates compare for I-III-VI2 arrays? Array interconnects?

4 – Effect of operating temperatures (space environment) on TFPV? Self-annealing?

5 – How far can we push polymers? Are there inherent efficiency limiters? Sodium?

Deposition/processing temperature? Do we need new balance of system technologies?

6 – Optimal deposition methods? Important device components? Multi-junction structures?

7 – Specific missions for TFPV? Large area? Surface operations? Integrated power? DODrelated missions? Air Force? Commercial?

NASA/CP—2005-213431 294
Issues Raised During Discussions

Four space PV companies (Sharp; RWE; Emcore; and Spectrolab) but none that focus on

TF cells

Only TF cells (not arrays) have been flown; TF cells shown to be radiation hard

Question: is radiation hardness a function of low efficiency? Or is it the fact that there is

not very much material for reactions to take place

Difficult to find suppliers for polymer substrates (low volume for applications) most

likely will be a small company

TF development can’t rely on literature and research foundation like Si and III-V

materials; the basic physics has not yet been developed

Thin-film development will most likely require government funding; IRAD will not be

sufficient for long-term development

Key Conclusion(s) Derived from Discussions

The prevailing view is that CI(G)S must be deposited on polymer substrates to be a

competitive technology

Thin film on polymer must have an efficiency of at least eight percent AM0 for a large

area

Array technology that is developed must be revolutionary and developed in parallel with

cell technology

Acknowledgements

We would like to acknowledge the excellent support of our peerless administrative assistant,

Mrs. Barbara Madej, during SPRAT XVIII and the follow-on Third Conference on Thin Film

and Nano- Materials for Energy Conversion and Storage. Her cheerful disposition, efficiency,

and professionalism turned a burdensome commitment into a pleasant undertaking!

NASA/CP—2005-213431 295
Informal Survey

The workshop attendees were asked an unscientific survey question: Are you optimistic

about the long-term prospects for thin-film technology for space power?

Results: Yes – 58% No – 21% Uncertain – 21%

NASA/CP—2005-213431 296
SPRAT Workshop: Nanostructures for Solar Cells 

Ryne P. Raffaelle, Chair 

The workshop on nanostructures for solar cells focused on a discussion of a few of the 

general features associated with 0-dimension (quantum dots), 1-d (quantum wires), and 

2-d (multiple quantum wells) approaches to photovoltaics developments. A particular 

focus of the discussion was directed toward the features of these quantum confined 

structures which may be particularly useful in the development of space photovoltaics. 

After the identification of general features, their examination, and discussion of their 

utility the next phase of the workshop was to try and identify the best prospects for near, 

mid, and far term developmental efforts. 

The general features associated with nanostructured solar cells which were discussed are 

listed below. 

General Features

Spectrum Shifting Applications 

Minibands/Hot Carrier Transport 

Multi-Photon Processes 

Exciton Disassociation in Polymers 

During the discussion of these general features there were several important 

considerations and possible advantages identified. In addition to the sub-gap absorption 

in the well region of a MQW solar cell, the increase in oscillator strengths really allow 

the I region of the cell to do most of the necessary absorption. This allows the host layers 

to be very thin to improve radiation tolerance. This was considered to potentially a very 

important advantage in developing future space solar cells. In addition, they can be more 

heavily doped to assist in collection and photovoltage. 

The problems associated with transport in nanostructured solar cells received quite a lot 

of discussion. The consensus was that fast carrier extraction and long-lived states would 

be extremely important in order to avoid the deleterious the recombination and 

photoluminescence. A strong field across the MWQ region would be essential. 

The discussion over the fundamental aspects of nanostructured cells brought up some of 

the controversy surrounding the reported global efficiencies of various nanotechnological 

approaches to photovoltaic development. The debate over global efficiency 

improvements with nanostructures involves much in the way of theory and conjecture 

and some of the finer points aren’t necessary to analyze real cells, and have little to do 

with relative performance enhancement or ancillary advantages of using nanostructures. 

Therefore the discussion at the workshop focused on pragmatic approaches and relative 

improvements with once again a particular focus on ancillary advantages for space usage. 

A comparison was made between a QD approach as opposed to a MQW approach to cell 

development. It was agreed that at this point it is clear that especially in terms of solar 

cells, MQWs are a much more mature technology that QDs. The possible decrease in 

NASA/CP—2005-213431 297
thermalization due to more confinement and two-photon processes which are 

theoretically possible in the QD approach have not yet been established to the extent that 

bandgap shifting and transport in MQWs has. The absorption hit that the quantum dot 

approach will take needs to be overcome by these other positive attributes before relative 

improvements are established. 

After discussing many of the salient features of nanostructured solar cells, attention was 

focused of realistic implementation schemes and timeframes under which these schemes 

could be accomplished. Based on this discussion several near, mid, and far term 

prospects were identified and the workshop was concluded. The prospects for 

nanostructured solar cells development resulting from the workshop are summarized 

below. 

Near Term

MQW solar cells already can be made thinner than their single junction counterparts due 

to their enhanced absorption (less weight) and are more radiation tolerant and defect 

tolerant in general. Temperature coefficients are also better due to decreased 

thermalization and hot-carrier transport. 

Nanostructures have already been shown to expand the crystal growers palette (i.e., 0.1 to 

0.2 eV shift in absorption edge in a strain-balance MQW for III-Vs). This technology is 

ready for integration into multi-junction cells for improving the current limiting junction. 

This approach can provide a rather large photocurrent increase with only a small 

photovoltage penalty. 

Significant improvement in polymeric solar cell efficiency is possible through the 

introduction of nanomaterials. 

Mid Term

Use of quantum dots as up-converts could improve conversion efficiency in a wide 

variety of devices and concentrators. Amorphous silicon multi-junction is a technology 

that could potentially reap much benefit from this approach. The available materials 

systems need to be expanded. Some narrow gap and high electron affinity material 

system would be attractive. In addition, the two-photon processes need to be critically 

examined. 

Far Term

Single junction intermediate band solar cells (IBSC) with efficiency which can exceed 

even the best triple-junction cells of today. 

NASA/CP—2005-213431 298
Advanced Solar Cell Testing and Characterization 

SPRAT Workshop Summary 

Chaired by: 

Sheila Bailey, Henry Curtis 

and Michael Piszczor 

The topic for this workshop stems from an ongoing effort by the photovoltaic community and 

U.S. government to address issues and recent problems associated with solar cells and arrays 

experienced by a number of different space systems. In April 2003, a workshop session was held 

at the Aerospace Space Power Workshop to discuss an effort by the Air Force to update and 

standardize solar cell and array qualification test procedures in an effort to ameliorate some of 

these problems. The organizers of that workshop session thought it was important to continue 

these discussions and present this information to the entire photovoltaic community. Thus, it was 

decided to include this topic as a workshop at the following SPRAT conference. 

The Advanced Solar Cell Testing and Characterization Workshop was divided into two areas of 

discussion. The first topic focused on the Air Force effort to standardize qualification and testing 

procedures. Brad Reed from the Aerospace Corporation has led the Air Force effort and was 

invited to lead that portion of the workshop. After that discussion, Sheila Bailey discussed recent 

progress in test procedure guidelines at the international level. 

Brad Reed started the first part of the workshop with a short talk on the rationale behind the Air 

Force push to establish standardized test procedures. He noted that the Aerospace Corporation, 

under direction from the Air Force, is engaged in a project to create military-specifications for 

inclusion in future Air Force RFPs. He explained that this initiative was borne out of an 

investigation of recent on-orbit and ground performance problems with military and commercial 

“space-qualified” solar arrays. 

In support of this project, Aerospace is submitting a proposed set of qualification tests to the 

industry for comment. A preliminary version of these proposed tests were discussed during the 

workshop and are included in the following three tables. The tables address two primary areas: 1) 

combined effects testing for solar cell and/or solar panel coupons and 2) accelerated life testing 

for solar cells for the purpose of determining reliability. 

The discussion during the workshop session proceeded to address the individual points in the 

tables, with a focus on soliciting comments/improvements to the proposed guidelines from the 

workshop participants. Specifically, input was being sought on: 

1. Appropriateness of the tests. 

2. Acceptability of the test procedures and standards to the space photovoltaic community. 

3. Suggestions for new tests that would enhance the qualification process. 

4. Data that could be included by the solar cell manufacturers to enhance the design process. 

Brad concluded the discussion by noting that further input from the community would be 

solicited and this process would continue.

NASA/CP—2005-213431 299
Solar Cell Qualification Test

Environment/Reason Level Test Procedure Test Condition Criteria/Report

1. Weld or Solder Joint 

Qualification Cell

Forward and Reverse Dark 

Electrical Characterization 28C, sample size determined by TBD

Build Cell to manufacturer's specification and 

record process

Weld or Solder Interconnect Per Manufacturer's Weld or Solder Schedule

Forward and Reverse Dark 

Electrical Characterization 28C Measure 10x

Reverse Bias

Dark Reverse Bias @ 2.0 V for 10 

Seconds @ 80C

Forward and Reverse Dark 

Electrical Characterization 28C Measure 10x

NDE of Weld or Solder Joints

Pull Test Pull at 45 degrees

100% DPA

Evaluate weld or solder interface micro 

structure

2. Pre-Launch and On-Orbit 

Degradation or Failure

2 Coupons 

with Bypass 

Diodes

Hotbox Electrical 

Characterization/Visual 

Inspection

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight @ 80C, coupon size 

TBD. Visual inspect at 1x at 1 foot.

Build Coupons with Composite Substrate and 

Aluminum core, Standard Interconnect, and 6 

mil Ceria-doped microsheet Coverglass, using

standard production processes

Humidity Exposure

90 days @ 95+/- 4% Relative Humidity @ 

45C

Hotbox Electrical 

Characterization/Visual 

Inspection

AM0 @ 1353 w/m2, characterized by JPL 

Balloon Flight, 80C. Inspect with 1x 

eyeball at 1 foot.

Combined Effects Test

Thermal Cycle first coupon between -180 

to +80 for 15,000 cycles. Thermal cycle 

second coupon between -95 and +110 C 

for 5000 cycles. In hot cycle either (1) 

illuminate at AMO under load, or (2) dark 

forward bias.

Hotbox Electrical 

Characterization

AM0 @ 1353 w/m2, characterized by JPL 

Balloon Flight @ 80C. Visual inspect 1x 

at 1 foot. < 2% Maxiumum Power Degradation

NDE of Weld or Solder Joints

Pull Test Pull strength at 45 degrees

100% DPA TBD

3. Stack Reliability Measurement, 

Verify Ohmic Contacts, Verify 

Tunnel Junction performance Bare Cells X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight, 28C, 

Accelerated Life Test TBD

X-25 Electrical Characterization

AM0 @ 1353 w/m2, characterized by JPL 

Balloon Flight, 28C

DPA

NASA/CP—2005-213431 300
Solar Cell Characterization

Environment/Reason Level Test Procedure Test Condition Criteria/Report

1. Electron Radiation Effects

Bare Cell with ninterconnect, pinterconnect, 

diode 

interconnects X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C and 80C, sample size TBD

Baseline Test - Cell group shall not 

exceed 5% spread at Pmax

Electron Radiation

Radiation energies of .6, 1.0, and 12 

MeV at Fluences of 1e14, 3e14, 1e15, 

3e15, 1e16

X-25 Electrical Characterization 

after each Radiation Fluence

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C and 80C, sample size TBD

Cell group shall not exceed 5% spread at 

Pmax

Reverse Bias Characterization

Dark Reverse Bias @ 2.0 V for 10 

Seconds @ 80C after each fluence

Perform Reverse Bias/Electrical 

Characterization for maximum of 10 times

X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C

< 2% Average Efficiency Degradation 

from Baseline and < 3.5% for any 

Individual Cell

100% DPA TBD

2. Proton Radiation Effects

Bare Cell with ninterconnect, pinterconnect, 

diode 

interconnects X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C and 80C, sample size TBD

Baseline Test - Cell group shall not 

exceed 5% spread at Pmax

Proton Radiation

Radiation energies of 20 KeV, 50 KeV, 

100 KeV, 300 KeV, 1MeV, 3MeV, and 

10 MeV at Maximum Fluences of TBD

X-25 Electrical Characterization 

after each Radiation Fluence

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C

Cell group shall not exceed 5% spread at 

Pmax

Reverse Bias Characterization

Dark Reverse Bias @ 2.0 V for 10 

Seconds @ 80C after each fluence

Perform Reverse Bias/Electrical 

Characterization for maximum of 10 times

X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C

< 2% Average Efficiency Degradation 

from Baseline and < 3.5% for any 

Individual Cell

100% DPA TBD

3. UV Radiation Effects CIC X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C, sample size TBD

UV Soak

TBD Hours at 200 nm irradiation in a 

Vaccuum of 10-5 torr until electrical 

degradation plateaus < 1% per hour. 

Exposure for another TBD2 Sun-Hours

Report TBD1 hours, and TBD2 % 

degradation

X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C

< 2% Average Efficiency Degradation 

from Baseline and < 3.5% for any 

Individual Cell

4. Peeling Gridlines after 

thermal cycle Cell Tape Peel Test

Pull with 3M type 600 cellulose tape to 

Evaluate Contact Adhesion. Visual 

inspection at 10x under fluorescent 

illumination <1% AR and <1% metal peeling

5. Humidity in Storage Cell X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C, sample size TBD

Humidity Test 30 days @95% RH @ 45C

X-25 Electrical Characterization

AM0 @ 1353 w/m2 as characterized by 

JPL Balloon Flight Calibration Cell, 

28C

Average efficiency degradation of < 1.5% 

from baseline and no individual cell >2.5%

Visual Inspection 1x at 1 foot distance no obvious anomalies

6. Breakage of Cells on Curved 

Array CIC Bend Test multiple radius bend test measure bend radius up to cell fracture

7. Mechanical Strength - 

Handling Issues Cell

Yield Test as per ASTM E 855-

90 (2000), Test Method C TBD samples yield strength distribution curve

8. Photogeneration Response 

vs. Wavelength Cell

Internal or External Quantum 

Efficiency

Spectral Response from 250 to 2000 

nm for TBD energies after electrical 

radiation fluences of 1e14, 3e14, 1e15, 

3e15, 1e16

9. Off-Normal Electrical 

Response to Pointing Angle Cell

Vary Angle of Incidence and 

provide X-25 Electrical 

Characterization

Measure response at 5, 10, 20, 30, 40, 

50, 60, 70, 80, 85, 90 degrees. 

Measure cells at AM0 @ 1353 w/m2 as 

characterized by JPL Balloon Flight 

Calibration Cell, 28C

10. Temperature Coefficient Cell

Vary Temperature and Provide 

X-25 Electrical Characterization

Measure I-V curve to include Isc, Voc, 

Imp, Vmp between -180 and 150 C in 

10 degree increments on cells with no 

fluence, and after fluences of e14, e15, 

e16. Provide AM0 @ 1353 w/m2 as 

characterized by JPL Balloon Flight 

Calibration Cell, 28C.

11. Thermal Design CIC Hemispherical Emissivity from 5 to 50 um

12. Thermal Design CIC Solar Absorptance from 250 to 2500 nm

13. Electrical Characterization Cell

Dark I-V characterization 

between I = Isc to - Isc at 28C

14. Electrical Characterization Bypass Diode As per Mil-S-19000

NASA/CP—2005-213431 301
Solar Panel Qualification Test

Environment/Reason Level Test Procedure Test Condition Criteria/Report

1. Problems During Storage at Launch; 

Problems On-Orbit Coupon Electrical Characterization

AM0 @ 1353 w/m2 as characterized 

by JPL Balloon Flight, 28C, coupon 

size determined by TBD

Build Coupon with parts and processes to 

be used in flight

Humidity Exposure

30 days @ 95+/- 4% Relative 

Humidity @ 45C

Electrical Characterization

AM0 @ 1353 w/m2, characterized by 

JPL Balloon Flight, 28C < 2% Average Efficiency Degradation

Visual Inspection TBD TBD

Acoustic Test TBD

Electrical Characterization

AM0 @ 1353 w/m2, characterized by 

JPL Balloon Flight, 28C < 2% Average Efficiency Degradation

Combined Effects Test

Thermal Cycle between predicted low 

non-operational temperature -16C 

and predicted high non-operational 

+16C at 1.5 x mission cycles. In hot 

cycle either (1) illuminate AM0 under 

load, or (2) dark forward bias. 

Reverse bias TBD directly after 

forward bias.

Electrical Characterization

AM0 @ 1353 w/m2, characterized by 

JPL Balloon Flight, 28C < 2% Average Efficiency Degradation

2. Panel-Level VCM Panel Electrical Characterization

AM0 @ 1353 w/m2 as characterized 

by JPL Balloon Flight, 28C, coupon 

size determined by TBD

Build Coupon with parts and processes to 

be used in flight

Measure VCM on cold 

surface

Thermal Soak predicted high nonoperational temperature + 16C. Measure as per ASTM 595. Input VCM 

results to vehicle contamination analysis.

Electrical Characterization

AM0 @ 1353 w/m2, characterized by 

JPL Balloon Flight, 28C < 2% Average Efficiency Degradation

NASA/CP—2005-213431 302
Irving Weinberg Award 

Presented by Dr. Sheila Bailey 

The Irving Weinberg Award, in memory of Dr. Irving Weinberg (1919–1995), a leading contributor 

to the field of space photovoltaic research and development for most of his professional career, was 

established at the 14th Space Photovoltaic Research and Technology (SPRAT) conference in 1995. This 

award is to be given at every SPRAT meeting to persons who have made significant contributions to the 

field of space photovoltaics. This award is meant to be inclusive of all aspects of space photovoltaic 

research and technology, from fundamental investigations of semiconductor materials, to device 

improvements, and finally to innovations in hardware for actual mission applications. The recipients are 

14th SPRAT, October 24 to 26, 1995, Dr. Chandra Goradia 

15th SPRAT, June 10 to 12, 1997, Dr. Masafumi Yamaguchi 

16th SPRAT, August 31 to September 2, 1999, Dr. Dennis Flood 

17th SPRAT, September 11 to 13, 2001, Henry Curtis 

18th SPRAT, September 16 to 18, 2003, Dr. Bruce Anspaugh 

Irv would have been particularly pleased with the 

selection of Dr. Bruce Anspaugh. 

Bruce received his Ph.D. from the University of 

Nebraska and started work for the Jet Propulsion 

Laboratory (JPL), all in 1964. He has spent his whole 

career trying to damage things, starting with the 

electronics and solar cells onboard the Mariner IV 

spacecraft. This early work led to his involvement with 

the Solar Array group in 1968. His career has been 

particularly colorful beginning with a review of the very 

first addition of a Solar Cell Radiation Handbook 

written more for scientists than panel engineers, 

followed by in 1973 a “Black” handbook (chosen for 

the color of the textured surface solar cells developed in 

that time period); a “Purple” revised handbook was next 

in 1977 (chosen for the “Violet cells” developed by a 

Comsat group); next was the “Green” handbook in 1982 

that contained the first computer models (I’ve been 

letting my imagination run wild and believe that green 

was how he felt after punching all those computer 

cards), and finally the “Blue” GaAs radiation damage handbook in 1996 (here I’m thinking the space 

community was singing the blues wondering if this would be the last book). Amidst developing these 

essential industry standards, Bruce had begged a Dynamitron from the Fission Physics Group at JPL and 

has spent the last 30 years or so keeping it all together and working. I think he really became interested in 

the balloon calibrations because of his love of flying. After all, you get to chase the balloon package 

across the Texas countryside after calibration via a homing beacon. Needless to say, those JPL balloon 

standards are the foundation for all simulator measurements. The Space community will long remember 

the remarkable contributions of Dr. Bruce Anspaugh. You will be missed when you retire in November. 

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18th Space Photovoltaic Research and Technology Conference

Thomas L. Morton, compiler

Solar cell; Photovoltaics power; Radiation damage; Space environment effects

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Proceedings of a conference held at Ohio Aerospace Institute and sponsored by NASA Glenn Research Center in

Cleveland, Ohio, September 16–18, 2003. Responsible person, Thomas L. Morton, organization code RPV,

216–433–6287.

The 18th Space Photovoltaic Research and Technology (SPRAT XVIII) Conference was held September 16 to 18,

2003, at the Ohio Aerospace Institute (OAI) in Brook Park, Ohio. The SPRAT conference, hosted by the Photovoltaic

and Space Environments Branch of the NASA Glenn Research Center, brought together representatives of the space

photovoltaic community from around the world to share the latest advances in space solar cell technology. This year’s

conference continued to build on many of the trends shown in SPRAT XVII—the continued advances of thin-film and

multijunction solar cell technologies and the new issues required to qualify those types of cells for space applications.