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Produced by the NASA Center for Aerospace Information (CASI)
Hughes Ref No. r0095 • SCG 85006' R

SPACE AND COMMUNICATIONS GROUP

am

FINAL REPORT

NASA WELDING ASSESSMENT PROGRAM

(NASA-CR-175512) NASA WELLING ASSESSHEHT N85-21661

PLOGHAM Final Belort (Hugbes Aircraft Co.)

44 n HC A03/dF 101 CSCL 13H

Unclas

G3/37 19382

JPL Contra :t No. 956038

December 1984

. `'"'1

AW IS?-5

wak so b`mqy .Z, MW M^
Hughes Ref No. FU095

It 49

SPACE AND COMMUNICATIONS OROUP

EL SEGUNDO, CALIFORNIA

FINAL REPORT

NASA WELDING ASSESSMENT PROGRAM

Prepared for:

JET PROPULSION LABOPATORY

CALIFORNIA INSTITUTE OF TECHNOLOGY

4800 Oak Grove Dive, Pasader:a, Ca;iforria 91109

JPL Contract No. 956038

December 1984

This work was- performed for the Jet Propulsion Laboratory. California Institute of

Technology sponsored by the National Ae ronautics and Space Administration unaer

Contract NAS7-918

Prepared By: E. J. Stofe!
This document contains information prepared by Hughes Aircraft Company under JPL sub-contract. !ts content is not necessarily endorsed by the Jet Propulsion Laboratory, California

Institute of Technology, or its sponsors.

ii
ACKNOWLEDGMENT

This author thanks the Hughes personnel who have contributed to tn.s project:

E. Browne, Jr., who was the engineer in charge of the ultrasonic welding, and R. Ritter,

who designed and fabricated the Kevlar substrate. Their insight into fabrication

processes used on many present spacecraft ensured that the module design selected for

testing is representative of a production approach using readily available materials Lnd

processes. The author also acknowledges the extensive contributions made by

Spectrolab, Inc., as the rnajc: subcontractor in this project. N. Mardesich and

M. Cillanders of Spectrolab made substantial contributions to the design of the modules

as well as supf-rvising solar cell manufacturing, parallel gap welding, and the soldering of

Spectrolab's portion of the modules and coupons. Technical discussions with

J. Scott-Monck of the Jet Propulsion Laboratory and with C. Bararona, J. Bo-..ek, and

R. Hart of the ;NASA Lewis Research Center also contributed significantly to the

establishment of test guide lines.

85006 l W P
ABSTRACT

A long duration test has been conducted for comparing various methods of

attaching electrical interconnects to solar cells for near earth orbit spacecraft.

Representative solar array modules have been thermally cycled for 36,000 cycles

between -800 and +800C on this JPL and NASA Lewis Research Center sponsored work.

This test simulates the environmental :,tress --): more than 6 years on a near earth 

spacecraft as it cycles in and out of the earth's shadow. Fva!;,ations of the integrity of

these modules were made by visual and by electrical examinations before starting the

cycling and then at periodic intervals during the cycling tests.

Modules included examples of parallel gap and of ultrasonic welding, as well as

4 soldering, which were fabricated by the Space and Communications Group of Hughes Aircraft Company and by Spectrolab, a subsidiary of Hughes. The materials and fabrication

processes are state of the art, suitable fo- forming large solar arrays of spacecraft

quality. The modules survived this extensive cycling without detectable degradation in

their ability to generate power under sunlight illumination.

1'1? ,'G1;D1NQ d'.AGE ; [.A.NK NO.T WILME-D

PA&E_,^,_INTE11nONALIY 6^IK

850061WP

-i 'Al 

a
IWOEDING .PA JE 13L U,4K NOT

850061W`P vii ^1Cr ^.: l

fN1^ITTOMALLY eLAI^IK
ILLUSTRATIONS

Page

1 Module Overview 6

2 Ultrasonically Welded Module Front 7

3 Ultrasonically Welded Module Back 7

4 Parallel Gap Welded Module Front 8

5 Parai,el Gap Welded Module Mack S

6 Module Cross Section 10

7 Photovoltaic Response of Representative Solar Module at

28 0C, AMO Illumination 11

8 Ultrasonic Weld Station Schematic 15

9 Ultrasonic Weld Front 15

10 Ultrasonic Weld Cross Section 16

11 Ultrasonic Weld Detail 16

12 Parallel Gap Weld 18

13 Parallel Gap Weld Cross Section 18

14 SCG Pull Tab Configuration 22

15 Ultrasonic Weld Pull Tabs 22

16 Spectrolab Pull Tab Configuration 24

17 Parallel Gap Weld Pull Tab 24

18 Parallel Gap Welded Cell Front 26

19 Parallel Gap Welded Ceti Back 26

20 Sunlight Simulation Testing Module Fixture 28

21 Thermal Cycling Equipment Schematic 30

22 Thermal Cycling Machine 30

23 !Module Test Tray 31

24 Module Test Tray Cross Section 31

25 Partially Assembly 'Module Tray 32

850061 WP viii
850061 WP ix

TABLES

Page

1 Test Requirements 2

2 Components for Modules 9

3 Component Weights for 9-Cell Modules 10

4 Arrav Power Density Performance 11

5 Pull Test of Ultrasonic Bonds 22

6 Pull Test of Parallel Gap Bonds 73

7 Fill Factors at 28 0C for 12 Modules During Thermal Test _4

8 Fill Factors at 80oC for 12 Modules During Thermal Test 35

w
1. INTRODUCTION

Future spacecraft are expected to have longer operating life than has been

typical of past spacecraft. NASA is now giving serious consideration to missions that

will operate for up to 10 years in near earth orbit. During this time, solar arrays will be

subjected to as many as 60,000 thermal cycles as they pass in and out of the earth's

shadow. Since any solar array is fabricated from a variety of materials hiving differing

coefficients of thermal expansion, this thermal cycling will produce repeated cycles of

strain and stress within the array. If suff:cier tly severe, the resulting fatigue will rupture the continuity of the delicate electrical mite connects between cells. The lots of

electrical continuity will mean loss of electrical power to the spacecraft.

NASA has embarked ret-ently on an experimental investigation of the extEnt to

which such thermal cycling poses a threat to the long duration operation of solar arrays.

The comparison between welded inter -onnects and soldered interconnects is receiving

particular attention. The metallurgical simplicity inherent in welds gives them a theoretical advantage over the more massive, multimetal composite structure of soldered

joints. However, practical space experience with welded interconnects operating for

many years in near earth orbit does not exist. NASA, therefore, is assessing the present

state of the art of welded and of soldered solar cell interconnects by sponsoring a series

of laboratory tests that are cycling representative solar a-.-ray modules under conditions

that thermally simulate long duration exposure in space. This study is one of these NASA

sponsored investigations.

The procedure for this investigation is as follows:

1) Review candidate solar array materials and fabrication processes as they

exist today.

850061 WP
2) Select an optimum combination of these state of the art materials and

processes to achieve a lightweight solar array design that emphasizes long

term tolerance of repeated thermal cycling.

3) Fabricate representative small modules of the selected design.

4) Measure electrical performance of these modules.

5) Subject the modules to extended, repetitive thermal cycling.

6) Periodically interrupt the cycling long enough to remeasure electrical performance to dei:ect any thermally induce electrical degradation and record

it as a function of the increasing number of cycles.

The general configurations of the test modules and of the test parameters, specified by the Jet Propulsion Laboratory acting in behalf of NASA, are summarized in

Table 1. These test conditions are similar to conditions that NASA specified for concurrent companion solar array cycling investigations sponsored by NASA at &-ockheed Missiles and Space Company and at TRW. The results of these three investigations, each

conducted independently, provide a valuable survey of the durability of state of the art

solar arrays.

TABLE I. TEST REQUIRErENTS

Parameter Comment

Simultaneous testing of at

least 8 modules

Nominal temperature limits

Uniformity over test plane at

temperature limits

Atmosphere

Thermal cycles

Scheduled inspection cycles

3 SCG we l ded modules

3 Spectrolab welded modules

I SCG soldered module

I Spectrolab soldered module

# 80c to -8C0C

+50C

Dry N2

36,000

0

500

1,000

2,000

4,000

8,000

12,000

24,000

30,000

36,000

350061% P
The rresent study was conducted by the Spare and Communications Group (SCG)

of Hughes Aircraft Company and by Spectrolab, a subsidiary of Hughes. Altnouph both

organizations are pert of Hughes, they operate independently under two separate profit

centers with differinL, background experience on solar array fabrication. During the

initial phase of this investigation, it became apparent that this difference was leading to

alternative design approaches. Rather, than eliminate one a p proach in favor of the

other, Hughes decided to try both. The cycling test would then be used to determine

whether one approach excelled over the other. The approach selected -)y SCG features a

cell to cell electrical interconnect that is as thin and compliant as possible (0.011 mm

thick silver mesh), accommodating in a low stress manner the differences of thermal

expansion and contraction of differing adjacent materials. The approach selected by

Spectrolab features a cell interconnect material that matches the thermal expansion of

silicon (silver plated Invar or molybdenum), minimizing stress by minimizing expansion

differences. The SCG interconnect lends itself best tc ultrasonic welding between the

interconnect and cell, demonstrated earlier in the thin cell welding part of this JPL

contract.* The Spectrolab interconnect bonds best to a cell by parallel gap welding.

Both types of interconnects can be readily soldored. For all other components of the

solar array modules, agreement was made to use materials and processes in common.

*E. J. Stofel, Ultrasonic Seam Welding on Thin Silicon Solat Celis, Final Report, for JPL

Contract No. 956033, Hughes Aircraft Company SCG 82041 -9-P, June 1982.

850061 % P
^. MODULE DESCRIPTION

2.1 MODULE OVERVIEW

The modules tested in this investigation have the configuration depicted

schematically in Figure 1. Figures 2 through S are examples. All modules are similar

except for differing methods of forming the interconnects. Comparison of performance

after cycling is a comparison of the curability (-f the examined types of interconnects.

The module components, which are s pace , ,jality, are summarized in Tab:e 2.

Their arrangement is shown in Figure 6, and the weights are given in Table 3. Figure 7

presents the photovoltaic power response of a representative module. The NASA

electrical performance requirements for the modules were exceeded as summarized in

Table 4.

State of the art fabrication processes were used to assemble these modules. The

cells, covers, and adhesives arc typical of those used extensively in space. Soldering is

also commonly used on spacecraft arrays. While the welding on these modules has not 

been used by SCG or Spectrolab in space, extensive laboratory development has been

performed by both organizations with results demonstrated and evaluated on several

thousand cells. The welding processes used for fabricating the samples of this present

thermal cycling program are those optimized for antici pated space application.

Twent y -three modules were fabricated: 9 welded and 3 soldered SCG and

7 welded and 4 soldered by Spectrolab. The slight differences, noted in Table 4 in areal

and in mass densities between the ultrasonic and the parallel gap welded modules, are

caused by the differences of interconnect th i ckness and flexibility. Not onl y was the

interconnect used on the ultrasonically welded module! lighter in weight than the interconnect on the parallel gap welded modules, its tninness also permitted less adhesive to

he used between the cover and the cell (Table 3). Ors the other hand, the rigidity of the

interconnect used with parallel gap welding allowed the cell to cell spacing gap to be

3 50061 W P 5 PAC, - 

.._L T&TWAu.Y d AN

1 110!('El )ING 1'Af;E BLANK NOT FILNIF0

- 
9 SILICON SOLAR CELLS, 21 1 x 40.3 x 0.2 MM.

INTERCONNECTED 3 IN PARALLEL BY 3 IN S1 RIES

L_

SERIES

DIRECTION

OVERVIEW

6

held consistertly to a smaller value Than for the r-nore flex ; ble, ultrasonically welded

interconnect which resulted in a higher areal density with the parallel gap welded

module% (Table 4). in all other respects, the modules were similar.

m

127 MM v

bi
0

8

N

Cy

FIGURE 7 ULTRASONICALLY WELDED MODULF FRONT

m

0

0m

.0

FIGURE 3. ULTRASONICALLY WELDED MODULE BACK
pnl^^uryuu^l

4 5

mLr

FIGURE 5. PARALLEL GAP WELDED MODULE BACK

FIGURE 4. PARALLEL GAP WELDED MODULE FRONT

8 URIGINAL PAGz

OF POOR QUALITY 
TABLE 2. COMPONENTS FOR MODULES

Si

Component Requirement

Solar cells 14% efficient

Silicon

10 ohm-cm P base with dual antireflective coating (DAR),

back surface reflector (BSR), and back -urface Field

(BSF ).

21.1 x 40.3 x 0.20 mm

Covers C7940 silica glass

MgF antireflective overcoating

L; treflecting ?aviolet undercoating

21.1 x 40.3 x 0.15 mm

Cover to cell adhesive DC 93-500 silicone

Interconnects

SCG ultrasonic welding 0.011 mm thick pure silver foil

Mesh configuration

Chemically etched

Spectrolab parallel gap welding 0.035 mm thick Invar with 0.007 mm thick silver plate on

each side

Mesh configuration

Chemically etched

SCG soldering 0.011 mm thick pure silver foil

Mesh configuration

Chemically etched

Spectrolab soldering 0.030 mm thick molybdenum with 0.005 thick silver plate on

each side

Mesh configuration

Chemically etched

Solder 38% lead

2% silver

60% tin

Screen printed pads on Spectrolab soldered modules

Cell to substrate adhesive DC 93-500 silicone

Substrate Kevlar fabric, rigidized with epoxy resin

850061 WP 9
r .

COVER 0° SILICONE

ADHESIVE

N WELD OR

SOLDER

NINTERCONNECT SOLAR CELL

\rs' ! / P

INTERCONNECT

SILICONE ADHESIVE

P WELD OR

SOLDER

KEVLAR FABRIC

KEVLAR AND EPDXY

SUPPORT BAR

FIGURE 6. MODULE CROSS SECTION

TABLE 3. COMPONENT WEIGHTS FOR 9-CELL MODULES

Component Ultrasonic, gm Parallei Cap, gm

Cells, 21.1 x 40.3 x 0.20 mm silicon with 3.80 3.80

DAR, BSR, BSF

Covers, 21.1 x 40.3 x 0.15 mm C7940 silica 2.82 2.82

Cover adhesive, DC-93-500 0.30 0.70

Series interconnects 0.26 0.38

Parallel interconnects x,.09 0.18

End connectors 0.09 0.30

Cell to substrate adhesivf . , DC-93-500 0.40 0.40

Substrate, Kevlar 1.35 1.35

Total 9.1 9.9

850061 VI'F 10
1.2

1.0

OA

Q

H

z

cc 0.6

U

0.4

0.2

W

0

0

O

18c = 1.029 A

V OC = 1.799 V

I Mp = 0.960 A

V Mp = 1.500 V

AMAX = 1.440 W

CF - 0.778

0 1 1 1 1 1 1

0 0.5 1.0 1.5 2.0

VOLTAGE,V

FIGURE 7, PHOTOVOLTAIC RESPONSE OF REPRESENTATIVE

SOLAR MODULE AT 2EPC, AMO ILLUMINATION

TABLE 4. ARRAY POWER DENSITY PERFORMANCE

Achievement

Parameter Requirement Ultrasonic Parallel Gap

Areal density, w/m 2

>140 173 178

Mass density, w/kg > 90 169 157

Achievement:

Module power at 28 0C, AMO, W 1.44 1.44

Active Module area, mm 8343 8073

Weight of active area, gm 8.5 9.1

Weight of total module, gm 9.1 9.9

? 5006 l W P
2.2 MATERIALS

For all modules, the selected solar cells were of Spectrolab K6700 type. These

cells have a shallow, planar N on P junction, a front dual antireflective (DAR.) coating, an

aluminum rich back surface field (BSF), and an aluminum back surface reflector (BSR).

The contact metalization is the conventional titanium-palladium-silver combination.

These cells have an efficiency of 14 percent at 280C under air mass zero (AMO) solar

illumination. This type is already in standard production and would be readily available

for larg: area solar arrays. The use of a back surface field is an appropriate means of

achieving high power density since at near earth orbits the radiation environment is

relatively benign. The enhanced power achieved with the BSF would therefore not be

lost as a result of radiation damage. The planar front surface was selected over the

nominally more efficient (at standard 28 0C) sculptured surface because the planar cell

has a lower solar absorbtivity which produces a relatively cool operating panel. The

cooler operation enhances power output in orbit.

The coverglass for the cells is a space qualified type manufactured by the

Optical Coating Laboratories, Inc. (OCLI). The base material is Corning 7940 fused

silica. Both bottom and top surfaces are polished flat. The top surface has a magnesium

fluoride coating for antireflection of visible light. The bottom surface has OCLI's

multilayer, ultraviolet reflective coating. This coverglass is used routinely on solar

arrays assembled by Hughes and by several other spacecraft manufacturers.

The covers were bonded to the front of the solar cells with DC-93-500 clear silicone adhesive which is used conventionally to bond covers to cells on all flight programs

at SCG and at Spectrolab. Alignment of covers over the cells was achieved using

mechanical fixtures, as used for present flight hardware. The coverglass covered each

cell completely (including the front contact area). Therefore, all cells had their interconnects attached to the front contact, either by welding or by soldering, prior to bonding the coverglass to the cell. Tie subassembly resulting was a covered, interconnected

cell (CIC). This type of subassembly, with welded interconnects, is used on the solar

array for Intelsat VI. SCG also is planning to use this assembly configuration on future

welded solar arrays. This approach differs slightly from present practice on SCG

soldered solar arrays wherein covers are bonded to the front of the cells prior to

soldering the interconnect to the front contact. For this latter configuration, the covers

are slightly undersize so that the front contact is exposed for subsequent soldering.

850061 WP 12
There is no apparent reason why future soldered arrays could not iust as well b^

assembled using the CIC configuration. For the SCG modules of this test program, solder

was applied by hand. For the Spectrolab modules, solder was applied as screen printed

pads.

The substrate for all modules was formed from a sheet of Kevlar woven fabric.

Prior to bonding cells to this fabric, it was impregnated with a small amount of liquid

epoxy. The fabric was held flat while the epoxy cured. This process produced a substrate with in-plane rigidity but with out-of-plane flexibility, similar to a sheet of heavyweight paper. It no longer had the flexibility and drape of the untreated fabric. The

stiffness imparted by the epoxy made it practical to bond cells to it en masse. There was

no possibility of local distortion of the interconnect during bonding which would have

occurred had the fabric maintained a clothlike drape. It was as easy to handle this epoxy

impregnated Kevlar as it would have been to handle a sheet of Kapton, a more traditional

material for substrate use on lightweight solar arrays. Kevlar has the advantage over

Kapton of being more tear resistant, a lower coefficient of expansion, and a surface

sufficiently porous to readily provide 3 strong, interlocking bond t-) the silicone adhesive

that was used to bond the cells to the substrate.

The electrically interconnected solar cell strings were bonded to the substrate

using DC-93-500 silicone adhesive, a material selected from among several candidates.

An alternative candidate, considered in detail, was precut pieces of uncured sheet epoxy.

The intent with this epoxy material was to limit the adhesive to a selected area of the

back of each cell so that no adhesive would flow into the interconnect. By keeping the

interconnect free of any restraining epoxy, the thin mesh style interconnect could flex in

a low stress, uninhibited manner to accommodate thermal expansion and t-ontraction

most readily. However, the application of precut adhesive film was found to be less

convenient as a mass production method than the simple method of spreading a thin,

uniform layer of silicone adhesive onto trio subaa rate. Furthermore, the use of epoxy

film, which was precut to be small enough to avoid the interconnect area on the back of

the cells, also significantly reduced the area of heat transfer between the cell and the

back of the solar array. The array would therefore have operated at a higher, less

efficient temperature.

Of the various candidates for a silicone adhesive, DC-93-500 was selected

because its transparancy allows visual inspection of the interconnect before and after

850061 WP 13
cycling even for those modules containing sufficient adhesive to engulf the

interconnects. Furthermore, this material is low in outgassing and is completely

acceptable for any flight array. Indeed, besides being the adhesive used regularly

oetween covers and cells, it has already been used to bond solar cells to substrates on a

few past spacecraft. For fabricating most modules of this test project, the amount of

DC-93-500 applied between the cells and substrate was :ontrolled to a low level, an

amount insufficient to allow excess accumulation between the cells at the cell

interconnects. However, to test for the effect on thermal cycle endurance that an

excess would cause if it became an accidental manufacturing variable under future mass

production, sufficient adhesive was intentionally added on a few modules so that excess

adhesive arose between cells, covering the cell interconnects. Thus, comparing the

durability of modules with the two configurations of adhesive provided information as to

whether the presence or absence of adhesive enbedment of interconnects affects the

longevity of the interconnect under extended therrnal cycling.

2.3 WELDING DESCRIPTION

The ultrasonic welds were made using a rotating. seam welding machine. This

machine and its operation were described in a previous report of this JPL contract and at

the Sixteenth IEEE Photovoltaic Specialist Conference.* The principle of its operation is

depicted in Figure 8. The ultrasonically active wheel produces a track of weld spots (see

Figure 9) as it rolls over the mesh st), le interconnect foil. Such welding produces an

intimate bond between the interconnect and the silver metallization on the solar cell. A

photomicrograph of a metallurgical cross section of one of these weld spots is shown in

Figure 10. There were 45 such bonds formed in parallel redundancy on the front of each

cell and 45 on the back. The intimacy of th-- bond is emphasized in Figure 11, which

shows a magnified view of the central part of Figure 10. The small void in the band of

metal at the center of this figure provides an unusual opportunity to reveal the location

of the interface between the applied interconnect and the original silver contact. The

source of this void was a shallow depression that had existed on the surface of the silicon

*E. J. Stofel, Ultrasonic Seam Welding on Thin Silicon Solar Cells, Final Report, for JPL

Contract No. 956038, Hughes Aircraft Company SCG 8204198, June 1982. E.J. Stofel,

E.R. Browne, R.A. Meese, and C.J. Vendura, "Welded Solar Cell Interconnection," 16th

Photovoltaic Specialists Proceedings, 1982, pp. 45-50.

850061 WP 14
co

C,

0

0

m

0

UHjUlvi NL V 

OF. POOR QUALIV v,

/ WHEEL TRANSLATION 0

0

WHEEL ROTA-1 ION 9

INTERCONNECT t'

CLAMP WELD TIP

/WHEEL AXIALLY DIRECTED

ULTRASONIC MOTION

TRANSDUCER

INTERCONNECT — _

/AXIS

f/ /TRANSDUCER

ALIGNMENT

STEP

I \ I `- /

SOLAR I /

CELL

WELDING

ANVIL

MOTION OF

HORIZONTAL CLAMP —

HORIZONTAL CLAMP'

FIGUPF 8. ULTRASONIC WELD STATION SCHEMATIC

FIGURE 9. ULTRASONIC WELD FRONT

850061WP 15
coT

c

0

m

I pw

INTERCONNECT

qkw

METALLIZATION

M,

CELL

I J

0.010 MM

go

a.

ORIGNAL PAGE IS

OF POOR QUALITY

FIGURE 10. ULTRASONIC WELD CROSS SECTION

0

0

0

0

0

FIGURE 11 ULTRASONIC WELD DETAIL

16
prior to applying the silver front contact. The contact metallization, applied by the cell

manufacturer, followed the contour of the silicon; therefore, it also had a shallow

depression at this location. The subsequent ultrasonic welding of the interconnect to the

cell did not deform the interconnect sufficiently to fill this depression. This depression

thus remains as a visually distinct mark that not only reveals the location of the original

interface between the interconnect and the cell contact, but by contrast with the

material to the left and to the right, also emphasizes the metallurgical continuity across

the weld interface that exists over the majority of the welded area.

The parallel gap welds were made by pressing two small, closely spaced

electrodes against the top surface of the interconnect as it lay upon the cell contact. A

pulse of current was conducted down one of the electrodes, through the interconnect

material bridging the small gap between the electrodes, then up the other electrode. The

current pulse produced sufficient localized heating in the interconnect to cause the

interconnect to bond to the cell metallization in the interface region under the action of

the high mechanical pressure applied by the electrodes. The size of the electrical pulse

was carefully controlled to provide a temperature sufficiently near melting to ensure

strong bonding while at the same time not overheating to a degree tn.it would have

produced extensive melting and loss of interconnect integrity. This parallel gap method

has been a common welding process for several years on a variety of electrical circuits.

Specifically, it has been the welding method most commonly used by several

organizations for welding interconnects to solar cells, including several flight solar

panels that have been constructed in Europe. A pair of weld spots from the present

program is shown in Figure 12. A photomicrograph of a metallurgica! cross section of

one of these weld spots is shown in Figure _:. Eight such spots were used in parallel

redundancy on the front of each cell and eight on the back.

The two types of welding methods, each with its distinctive interconnect, were

chosen to provide emphasis on comparing two different approaches toward achieving long

fatigue life. Either of these approaches must provide tolerance for two types of

thermally induced stress. The first type occurs as a shear stress at the interface bond

between the interconnect and the solar cell, the result of differing thermal coefficients

of expansion between the metal interconnect and the underlying silicon. The second type

of stress occurs from changes in cell to cell spacing caused by the difference in

expansion between the substrate and the interconnect. This puts the interconnect

alternately in tension and compression. Not only does this fatigue the portion of the

850061 WP 17

W,
UKIGINAL PAGE M

OF POOR QUALITY

FIGURE 12. PARALLEL GAP WELD

0

m

w

OD

N8 O^

N

INTERCONNECT

Vir

WELD

CELL

FIGURE 13. PARALLEL GAP WELD CROSS SECTION

18
^' w

W f

interconnect bridging between cells, it also aggravates the shear stress acting on the

bond line between the interconnect and the solar cell.

The ultrasonically welded modules feature a very compliant interconnect. The

thinness of the foil (Figure 10) allows large deflections to occur in the out of plane loop

(Figure 6) without producing large fatigue strain within the metal. The change in cell to

cell spacing caused by expansion and contraction under thermal cycling therefere is

accommodated without large bending stress in the interconnect nor large force being

applied to the welded area. Thermally induced shear at the bond line also :s kept small

by having the metal thickness overlying the silicon as thin as possible, thus partially

alleviating the problem of their differing coefficients of expansion. The relatively large

holes in the diamond shaped mesh pattern (Figure 9) aid in providing compliance in the

plane of the foil and retard tear propagation. The narrow strips of metal around the

holes result in a large number of small weld areas. All act electrically in parallel

redundancy; however, the com p liancy of the mesh allows each weld area to be

mechanically decoupled from its neighbors. The accidental debonding of one small weld

area therefore does not cause propagation of debonding to adjacent weld areas.

Since silver is an excellent electrical conductor, this low cross section, highly

compliant interconnect cortiguration is permitted withOL I t introducing significant series

electrical resistance. Ultrasonic welding is a good production method of rapidly making

many weld bonds to such thin material without the possibility of accidentally overhez r ting 

and melting it.

The parallel gap welded modules feature an interconnect with a very low coefficient of thermal expansion, one approximately the same as the underlying silicon. By

matching coefficients, :he contact shear stress can be kept low over a range of operating

temperatures. The parallel gap welded modules of this program use a silver plated Invar

j i-terconnect. The Invar provides the low expansion; the silver is required for adequate

jelectrical conductivity. This combination is relatively thick (o)mpare Figure 13 with

I , Figure 10) since it is difficult to obtain thinner Invar foil. The resultant interconnect is

much sti f fer than the simple, thin silver interconnect. This stiffness is caused by the

greater thickness and higher modulus of elasticity of Invar. The out of plane s!: ess relief

loop (Figure 6) thus dces not provide m • 1ch compliance within the Invar interconnec'.. To

prevent large stress in the interconnect, the substrate, therefore, ,- g ust also have a low

thermal coefficient of expansion. Kevlar was chosen for this substrate.

850061 W P 19

• .^s;i-
2.4 STRENGTH OF WELDS

Prior to fabric?zing the thermal cycle test modules, tabs of representative interconnect material were we i dej to severa! sample cells to demonstrate the production

readiness of both welding methods. The stren g th of these welds was evaluated by pulling

the interconnect tabs to destruction. Four types of separation could occur:

l) Simple separation at weld interface between the interconnect and cell

metallization.

2) Fracture of underlying silicon material.

3) Rupture of interconnect material immediately adjacent to the weld where

the interconnect had been deformed by the % • gilding process.

4) Plastic stretching of interconnect tab and its eventual rupture somewhere

along the length of the tab away from the w gilded region.

I` a large pull force was required to cause destruction, than no matter what its

failure mode that weld was considered to nave been of adequate strength for typical

flight applications. If a small force caused uestruction, then the different fa i lure modes

were indicative of different problems with welding. Type l failure, parting at the weld

interface, could arise at a low force if insufficient energy had beer delivered during

welding which would leave ar. incomplete weld interface. Contaminated surfaces prior to 

welding could produce a similar result. T y pe 2 failure, rupture of the underlying silicon,

could indicate the introduction of microcracks at the time of welding due to too much

weld pressure or energy. Type 3 failure, rupture of the interconnect adjacent to the

weld, arises when too much weld energy is used and the cross sectional area is greatly

reduced due to extensive plastic deformation of the interconnect metal. Type 4 failure,

plastic stretching of the interconnect, is not a failure mode associated with the weld. If

stretching had occurred at an abnormally small force, it would have indicated a defer. in

the interconr,+.ct or an improperly aligned pull test.

For testing the ultrasonically welded cells, a mesh interconnect was first welded

along the full width of the cell (as depicted in Figure 9) in the same manner that was

used on the thermal cycle test modules an;; will be used on flight solar arrays. The mesh

was then cut with a sharp razor blade into several parallel 3.0 mm wide tabs, each with

550061Al P 20

E•
^N

'i

I

the dimensions indicated in Figure 14. Tabs prepared this way had welds that were

representative of rolling seam welding, with each tab having f0L 1 r small, independent

weld areas.

An example of the result obtained by pulling four adjacent tabs is presented in

Figure 15. Two tabs failed by type 3, two by type 4. Table 5 summarized pull results of

all tested ultrasonic welds. Some of these welds were pulled at an angle of 45 0 with

respect to the plane of the front surface of the cell; the remainder were puiled normal to

this plane. The 450 pull produ :e. a combination of shear and peel at the weld; the 900

pull produces only peel. The former condition is closer to the stress that can be expected

during normal handling of welded cell strings during array assembly operations. The latter condition makes a more sensitive distinction of strength variation among nominally

similar welds. Most tests ended by breaking the interconnect from 60 to 80 grams, the

inherent strength of the interconnect. Only at 90 0 did a portion of welds fail by peeling

apart at the weld interface. These pull results indicate an optimum set of weld

parameters had been used. If more weld energy or pressure had been used, there would

have been fewer welds separating in 90 0 peel. However, the resultant greater defor ► nation of the interconnect during the welding would have allowed some type 3 failures to

occur at less than 60 grams. On the other hand, with less welding energy, more 90 0 peeling would have occurred. The observed type 4 failure of the interconnect is the desired

result, since the interconnect .vas designed to have sufficient compliancy to

accommodate movement with easy strain rather than resisting movement by developing

high stress.

The front contact pull tab configuration for the parallel gap welded cells is

depicted schematically in Figure 16. A photograph of a back contact pull tab is shown in

Figure 17. Table 6 summarizes the pull tab results. Since these tabs are thick (see

Figure 13) and therefore have intrinsically high strength, none of the tabs broke. Instead,

separation occurred either at the weld interface or by fracturing the underlying silicon.

Since destruction force was large for all welds, the tests showed that parallel gap

welding produces strong bonds.

2.5 COUPONS

Twenty-four welded cell assembly coupons and 12 welded coupons also were

fabricated and sent to the NASA Lewis Research Center for evaluation per their

850061 W P 21
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ORIGINAL PAGE IS

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• TAB CONFIGURAT ION, MM • TAB LOCATION

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FIGURE 14. SCG PULL TAB CONFIGURATION

kv--"

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. ASHED LINES INDICATE EDGES OF I

TABS PRIOR TO PULLING

1

FIGURE 15. ULTRASONIC WELD PULL TABS

22
TABLE 5. PULL TEST OF ULTRASONIC BONDS*

Parameter At 450 Pull At 900 Pull

Front

Tabs tested 50 30

Tabs broken 50 26

Tab strength, gm 60 to 80 60 to 80

Welds separated 0 4

Separation force, gm N/A 10 to 60

Back

Tabs tested 30 10

Tabs broken 30 7

Tab strength, gm 60 to 80 60 to 80

Welds separated 0 3

Separation force, gm N/A 20 to 40

*Test with Unitron pull test machine.

TABLE 6. PULL TEST OF PARALLEL

GAP BONDS*

Parameter At 450 Pull

Front Contact

Tabs tested 14

Cells broken 5

Breaking force, gm 270 to 620 

Welds separated 9

Separation force, gm 240 to 610 

Back Contact

Tabs tested 10

Cells broken 1 

Breaking force, gm 400

Welds separated 9

Separation force, gm 240 to 500 

*Tested with Unitron pull test machine.

850061 WP 23
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URIGINAL PAGL IS

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• TAB CONFIGURATION, MM • TA3 LOCATION L, 0

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DIRECTION

FIGURE 16. SPECTROLAB PULL TAB CONFIGURATION

FIGURE 17. PARALLEL GAP WELD PULL TAB

24
request. Half of each type was welded at SCG, the other half at Spectrolab. Each cell

assembly coupon comprised a covered cell which had interconnects welded to both the P

and the N contacts, bonded to a segment of substrate. Each cell coupon consisted of an

! uncovered cell assembly coupon with interconnects welded to both the P and the N

contacts, bonded to a segment of substrate. Examples are shown in Figures 18 and 19,

prior to bonding to the substrate to depict the location of the welds on both the front and

back of the cells.

k

850061 W P 25
ORIGMIAL PAG:. 'r

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METRIC 1 2 3

FIGURE 18. PARALLEL GAP WELDED CELL FRONT

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FIGURE 19. PARALLEL GAP WELDED CELL BACK

26
3. THERMAL TESTING

Twelve modules were thermally cycled: 4 ultrasonic welded (SCG), 4 parallel gap

welded (Spectrolab), 2 SCG and 2 Spectrolab soldered. Two additional modules were used

as noncycled control moduies. Although not cycled, their electrical performance was

measured each time performance was measured for the cycled cells. These control

modules thereby sarved as a verification that electrical measurements were made in the

same manner each time even though several months elasped between measuring times.

Prior to starting the thermal cycling, all modules were examined optically

(at 10x) and anomalies were recorded. The full current voltage photovoltaic output curve

for each module at 28 0C under air mass zero illumination also was recorded. For these

measurements, each module was held by vacuum against a temperature controlled plate,

as shown in Figure 20. A reference cell was mounted permanently on this plate to verify

the level of illumination; two thermocouples verified the temperature of the plate.

These photovoltaic measurements and optical examinations were repeated after

cycle 500, 1,000, 2,000, 4,000, 12,000, 18,000, 24,000, 30,000, and 36,000. To make these

measurements, the cycling apparatus was stopped, the module tray disassembled, and

current-voltage measurements were repeated in the same manner as the precycling

measurements.

Photovoltaic measurements also were made at 80 00 before starting the thermal

cycling and then again after 12,000, 18,000, 24,000, 30,000, acid 36,000 cycles. These

higher temperature measurements were undertaken on the assumption that they provide

a more sensitive means of early detection of incipient failure. Expansion of the adhesive

between the cover and the cell would tend to pull the interconnect away frim the cell.

Thus, if a weld or solder joint had become loosened from thermal cycling, this high temperature expanded condition might have detected an open circuit whereas measurement

850061WP 27
28

ORIGINAL PAGE ,S

OF POOR QUALITY

F'GURE 20. SUNLIGHT SIMULATION TESTING MODULE FIXTURE
at 280C would not if the interconnect at this lower temperature remained touching but

was otherwise unbonded from its cell. These measurements at 80 0C were made with an

illumination several percent higher than AMO in order to exaggerate any changes in

series resistance.

To accomplish the thermal cycling, the 12 modules were supported on an aluminum tray stack that moved alternately from a warm chamber to a cold chamber at the

rate of approximately one cycle every 10 minutes. The mechanism is enclosed within a

housing that maintains a dry nitrogen gas environment. The mechanism is shown schematically in Figure 21. The housing, with its doors open to reveal the mechanism, is

shown photographically in Figure 22. After loading the test tray, the doors of the

housing were closed. The housing then was purged with dry nitrogen gas to remove all

water vapor prior to starting the thermal cycling. A positive pressure of nitrogen was

maintained throughout the cycling to ensure that no water or frost ever condensed on the

modules.

The tray stack consisted of four sheets of aluminum. The center two sheets each

supported six modules; the outer two aided in providing a uniform temperature

distribution across the entire area of the stack. This arrangement is depicted

schematically in Figures 23 and 24, photographically in Figure 25. The modules were

held in tension by metal coil springs attached along their edges, thus preventing curling

of the modules during the alternating expansion and contraction of each cycle. A leaf

spring of low thermal conductance held a low mass thermocouple against the center cell

of each module. A small amount of thermally conductive compound (Dow Corning 340

silicone heat sink compound) was used to ensure the thermocouple was at the same

temperature as the module. This compound retains its greaselike consistency over the

temperature range of this t est. It was easy to remove the compound from the modules

every time the module tray was removed from cycling to make the periodic currentvoltage measui- ements. The tray also contained additional thermocouples located so as

to detect any undesired thermal gradients. The temperatures of all thermocouples were

recorded periodically to form a record of module temperatures. Uniformity of t5°C over

all modules was specified by NASA at both +80 0 ana -800C. In practice, uniformity at

the high temperature limit usually was within t3 0C, at the low temperature within

t4°C. On rare occasions, the temperature set limits drifted up or down a couple of

degrees when operating automatically over an evening or weekend. While this led to a

850061 W P 29
SHAFT

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URIGINAL PAGE 1S

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COLD SWITCH 'REVERSING

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FIGURE 21. THERMAL CYCLING EQUIPMENT SCHEMATIC

FIGURE 22. THERMAL CYCLING MACHINE

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FIGURE 23. MODULE TEST TRAY

SPACER THERMOCOUPLE SOLAR CELL

FOR TRAY MODULE

FIGURE 24. MODULE TEST TRAY CROSS SECTION

31
URIGINAL PAGE IS

OF POOR QUALITY

FIGURE 25 PARTIALLY ASSEMBLED MODULE TRAY

oUO

N

{II

32
temporary excess of a degree or two beyond the nominal tS oC limit, such drifts affected

both the low and high limits in the same direction so that the total swing in temperature

remained close to 1600C.

The visual examinations and electrical measurements made periodically during

this cycling test did not reveal any large changes. The only change noted visually was a

slight roughening of the solder surface on some of the soldered cells after 8,000 cycles.

There was no evidence that this roughened condition led to any performance

degradation. The power output at AMO illumination and 28 0C remained unchanged.

The most likely electrical damage to search for on a thermally cycled module

would be an increase of st:ries resistance or a decrease in shurting resistance. Slight

changes in either would manifest themselves as a change in the fill factor of the currentvoltage curves for the modules. Table 7 shows there has been no change in fill factor at

28 0C. Table 8 shows there also was little, if any, change in the fill factor at 80 0C under

measurement conditions that should have exaggerated such change. Interpreting these

high temperature measurements with precision is difficult since 80 0C was near the upper

limit readily achievable with the test fixture shown in Figure 20. The data of Table 8

therefore may have more experimental scatter than that of Table 7. Only the soldered

Spectrolab cells show any suggestion of possible degradation. However, measurement of

these modules at high temperature was particularly difficult since the configuration of

electrical pressure contacts on the test fixture was not as compatible with the

contacting pad configuration of these particular modules as they were with the pads on

the other modules. Before concluding that an electrical degradation has been started on

these soldere l modules, thermal cycling would need to be extended well beyond the

present 36,000 cycles to see if the fill fact;.r established a more clearl y defined

downward trend or merely fluctuated from measurement to measurement because of

measurement difficulty.

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35
4. CONCLUSION

These tests have indicated that solar array modules constructed with state of the

art components and materials are able to w i thstand at least 6 years equivalent of near

earth orbit thermal cycling without fatigue indu--ed de;radation of power.

PIWEDING PAGE PLANK NOT F",V. PD

PAGE 3^ 

MEN MALD BLANK

85006 l W P 37

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