NASA / CP-1998-208536 

The 1997 NASA Aerospace Battery 

J.C. Brewer, Compiler 

Marshall Space Flight Center, Marshall Space Flight Center, Alabama 

https://ntrs.nasa.gov/search.jsp?R=19990017662 2018-05-14T12:47:47+00:00Z
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The 1997 NASA Aerospace Battery 

Workshop 

J. C. Brewer, Compiler 

Marshall Space Flight Center, Marshall Space Flight Center, Alabama 

Proceedings of a workshop sponsored by the 

NASA Aerospace Flight Battery Systems Program 

and held in Huntsville, Alabama 

November 18-20, 1997 

National Aeronautics and 

Space Administration 

Marshall Space Flight Center 

July 1998 
Available from: 

NASA Center for Aerospace Information 

800 Elkridge Landing Road 

Linthicum Heights, MD 21090-2934 

(301) 621-0390 

National Technical Information Service 

5285 Port Royal Road 

Springfield, VA 221 61 

(703) 487-4650 
Preface 

This document contains the proceedings of the 30th annual NASA Aerospace Battery Workshop, 

hosted by the Marshall Space Flight Center on November 1 8-20,1997. The workshop was attended 

by scientists and engineers fiom various agencies of the U.S. Government, aerospace contractors, 

and battery manufacturers, as well as international participation in like kind from a number of 

countries around the world. 

The subjects covered included nickel-cadmium, nickel-hydrogen, silver-zinc, nickel-metal hydride, 

and lithium- based technologies. 

I99 7 NASA Aerospace Battery Workshop 
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Introduction 

The NASA Aerospace Battery Workshop is an annual event hosted by the Marshall Space Flight 

Center. The workshop is sponsored by the NASA Aerospace Flight Battery Systems Program which 

is managed out of NASA Lewis Research Center and receives support in the form of overall 

objectives, guidelines, and funding from Code S, NASA Headquarters. 

The 1997 Workshop was held on three consecutive days and was divided into five sessions. The 

first day consisted of a General / Secondary Battery Session and a Lithium / Lithium-Ion Battery 

Focused Session. The second day consisted of a Nickel-Hydrogen Session and a Nickel Electrode 

Design Focused Session. The third and final day was devoted to an Advanced Nickel-Hydrogen 

Technology / Nickel Electrode Characteristics Session. 

On a personal note, I would like to take this opportunity to thank all of the many people that 

contributed to the organization and production of this workshop: 

The NASA Aerospace Flight Battery Systems Program, for their financial support as well 

as their input during the initial planning stages of the workshop; 

Rao Surampudi, Jet Propulsion Laboratory, and Joe Stockel, Office of Research & 

Development, for serving as Focused Session Organizers, which involved soliciting 

presentations, organizing the session agenda, and orchestrating the session during the 

workshop; 

Huntsville Hilton, for doing an outstanding job in providing an ideal setting for this 

workshop and for the hospitality that was shown to all who attended; 

Marshall Space Flight Center employees, for their help in mailing the various 

correspondence, registering attendees, handling the audience microphones, and flipping 

transparencies during the workshop. 

Finally, I want to thank all of you that attended andlor prepared and delivered presentations for this 

workshop. You were the key to the success of this workshop. 

Jeff Brewer 

NASA Marshall Space Flight Center 

1997 NASA Aerospace Battery Workshop 
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Table of Contents 

Preface.. ................................................................... iii 

Introduction ................................................................. v 

General I Ni-Cd I Ni-MH Battery Session ......................................... 1 

An Historical Summary and Prospects for the Future of Spacecraft Batteries 

G. Halpert and S. Surampudi, Jet Propulsion Laboratory .............................. .3 

Battery System Management -- Optimized Electrokinetic Control in Galvanic, Electrolytic, 

and Storage Operating Modes 

Floyd Williamson, ZAE Research, Inc. ........................................... .29 

Advanced Lithium-Ion Energy Technology 

Ricky J. Roberson and Greg Tyler, Huntsville Integrated Voltaic Equipment .............. 55 

Performance of Nickel-Cadmium Batteries on the GOES I-K Series of Weather Satellites 

Sat P. Singhal, Computer Sciences Corporation; Walter G. Alsbach, Jackson & Tull; and 

Gopalakrishna M. Rao, NASA Goddard Space Flight Center .......................... .83 

Near Earth Asteroid Rendezvous Flight Battery Performance 

Jason E. Jenkins, JHU Applied Physics Laboratory; Jeff W. Hayden and David F. Pickett, Eagle 

PicherIndustries,Inc. ........................................................ 113 

Crane Cell Testing Support of Goddard Space Flight Center 

Steve Hall and Harry Brown, Naval Surface Warfare Center, Crane; and Gopalakrishna Rao, 

NASA Goddard Space Flight Center ............................................. 149 

Status of Nickel Metal Hydride Cell Development 

Y. Sone, H. Kusawake, K. Koga, and S. Kuwajima, National Space Development Agency of 

Japan ..................................................................... 161 

Battery Systems for X-38 Crew Return Vehicle (CRV) and Deorbit Propulsion Stage (DPS) 

Eric Darcy, NASA Johnson Space Center ......................................... 179 

Lithium-Ion Focused Session ................................................ -2 13 

Improved Li-BCX Primary Cells for Space Applications 

D.M. Spillman, N.M. Waite, M.F. Pyszczek, and E.S. Takeuchi, Wilson Greatbatch Ltd. .. .215 

Development of Ultra Low TemperaGre, Impact Resistant Lithium Battery for the Mars 

Microprobe 

H. Frank, F. Deligiannis, E. Davies, B.V. Ratnakurnar, and S. Surampudi, Jet Propulsion 

Laboratory; P.G. Russel and T.B. Reddy, Yardney Technical Products, Inc. ............. .225 

1997 NASA Aerospace Battery Workshop -vii- 
Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications 

Lynn Marcoux and Gregg Bruce, BlueStar Advanced Technology Corporation .......... .265 

Rechargeable Lithium-Ion Battery Development at Eagle-Picher Industries 

Chad Kelly, Eagle-Picher Industries, Inc. ........................................ -299 

Evaluation of 20 Ah Li Ion Cells 

Marshall Smart, B.V. Ratnakumar, C.K. Huang, and S. Surampudi, Jet Propulsion Laboratory; 

Carole Hill and Dan Radzykewycz, Air Force Research Lab, Albuquerque; and R.A. Marsh, Air 

Force Research Lab, Dayton .................................................. .3 1 1 

The Li-Ion Battery Qualification Programme for STRV-ld 

Vijay V. Thakur, Defence Evaluation and Research Agency ......................... .341 

Rate Capability Analysis of Li-ion Cells 

Esther Sans Takeuchi and Randolph A. Leising, Wilson Greatbatch Ltd. ............... .365 

Nickel-Hydrogen Session .................................................... .387 

Mars Pathfinder Battery Performance 

B. Otzinger, D. Perrone, S. Dawson, T. Valdez, S. Surampudi, R. Ewell, and M. Shirbacheh, Jet 

Propulsion Laboratory ....................................................... .3 89 

Cycle Life Expectations for IPV Nickel-Hydrogen Cells 

Lawrence H. Thaller, The Aerospace Corporation ................................. .411 

Life Cycle Testing of Spaceflight Qualified Nickel-Hydrogen Battery Cells 

Chris Fox, Jennifer Francisco, Kevin Arnes, and Ron Repplinger, Eagle-Picher Industries, Inc. 

........................................................................... 429 

EOS-AM1 Nickel Hydrogen Cell Interim Life Test Report 

C.W. Bennett, Lockheed Martin Missiles and Space; D.J. Keys and G.M. Rao, NASA Goddard 

Space Flight Center; H.E. Wannemacher, Jackson and Tull; and H. Vaidyanathan, COMSAT 

Laboratories ............................................................... .457 

Characterization and Initial Life-Test Data for Computer Designed Nickel Hydrogen Cells 

A. Zimrnerman, J. Matswnoto, A. Prater, D. Smith, and N. Weber, The Aerospace Corporation 

........................................................................... 471 

Development of a Large Diameter (5.5 Inch) Nickel-Hydrogen (NiH,) Individual Pressure 

Vessel (IPV) Battery Cell 

Dwight Caldwell, Jim Morgan, and Jack Brill, Eagle-Picher Industries, Inc. ............. .485 

I997 NASA Aerospace Battery Workshop -viii- 
Charge Temperature Characterization of Nickel-Hydrogen Batteries 

D. Coates, R. Smith, D. Judd, and J. Brill, Eagle-Picher Industries, Inc.; J. McDermott and G. 

Gates, Lockheed-Martin ....................................................... .SO3 

Meeting the Demands of the Commercial Aerospace Market: Verification Testing and 

Product Qualification at a New Battery Manufacturing Facility 

Jeff Dermott and Dwight Caldwell, Eagle-Picher Industries, Inc. ...................... .523 

Nickel Electrode Design Focused Session ...................................... .545 

Cause for Second Plateau Discharge in Nickel Electrodes 

A. Zimmerman and N. Weber, The Aerospace Corporation .......................... .547 

Fabrication of Nickel Electrodes Using Ethanol Solutions 

David F. Pickett, Jr., Eagle-Picher Industries, Inc. ................................. .559 

Effects of Cycling Conditions of Active Material from Discharged Ni Positive Plates 

Studied by Inelastic Neutron Scattering Spectroscopy 

Juergen Eckert, Los Alamos National Laboratory; Ravi Varma, Lisa Diebolt, and Margaret Reid 

........................................................................... 567 

Study of Parameters Causing Second Plateau Effect on Nickel Hydrogen Cells 

Y. Borthornieu, SAFT Defense and Space Division; N. Sac-Epee, CNRS; and T. Jamin, CNES 

........................................................................... 587 

Performance Comparison Between NiH2 Dry Sinter and Slurry Electrode Cells 

J.D. Armantrout, Lockheed-Martin ............................................. .627 

Advanced Ni-H2 1 Nickel Electrode Session ..................................... .663 

Lander and Orbiter Battery 

William Cook, Eagle Picher.Inc. ............................................... .665 

IRIDIUM Satellite Batteries: Flight Observations "By the Dozen" 

Mark R. Toft, Motorola Space Systems Technology Group .......................... .689 

Behavior of 2-Cell CPV Ni-H, Battery During Pulse Discharge 

Hari Vaidyanathan, COMSAT Laboratories; and Gopalakrishna Rao, NASA Goddard Space 

Flightcenter ............................................................... 713 

Nickel-Hydrogen (NiH,) Single Pressure Vessel (SPV) Battery Development Update 

Steve Sterz, Beth Parrnley, Dwight Caldwell, and John Bennett, Eagle-Picher Industries, Inc. 

........................................................................... 729 

Mechanical Behavior of Nickel Oxide Electrode Under Shallow Cycles and Overcharge 

G. Davolio, E. Soragni, P. Zannini, and P. Baraldi, University of Modena .............. .733 

I997 NASA Aerospace Battery Workshop -ix- 
Characterization of Electrochemically Impregnated Fiber Nickel Electrodes 

D . Coates. J . Francisco. D . Chiappetti. and J . Brill. Eagle-Picher Industries. Inc ........... 755 

Nickel Electrode Corrosion During Nominal On-Orbit Operation 

Chuck Lurie and Bob Tobias. TRW Space and Electronics Group ..................... -777 

1997 NASA Aerospace Battery Workshop Attendance List ........................ 791 

1997 NASA Aerospace Battery Workshop 
General / Ni-Cd 1 Ni-MH Battery Session 

1997 NASA Aerospace Battery Workshop 
Page intentionally left blank 
AN HISTORICAL SUMMARY AND 

PROSPECTS FOR THE FUTURE OF 

SPACECRAFT BATTERIES 

G. HALPERT 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 

AND 

S. SURAMPUDI 

DEVICE RESEARCH AND APPLICATIONS SECTION 

NASA BATTERY WORKSHOP 

HUNTSVILLE, ALABAMA 

NOVEMBER 18,1997 
4- 

'9 

S 

JPL 

OUTLINE 

HISTORICAL EVOLUTION OF BATTERIES IN SPACE 

EVOLUTION AND STATUS OF NI-CD AND NI-H2 

PRESENT APPLICATIONS 

FUTURE APPLICATIONS 

ADVANCED BATTERIES FOR FUTURE MISSIONS 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
4- 

s JPL 

I HISTORY OF BATTERIES 

IN SPACE 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
SPUTNIK AgIZn 

1 W FOR 3 WEEKS 

VANGUARD ZnlHgO 

EXPLORER I ZnlHgO 

VAN ALLEN RAD BELT 

EXPLORER 6 Cyl NiICd 

FIRST EARTH PHOTOS 

RANGERS Pris NiICd 

MOON PHOTOS 

ARlEL I Pris NiICd 

First LEO MISSION 

SYNCOM-2 Cyl NilCd 

First GEO 

APOLLO CM Ag/Zn 

LTD CYCLE LIFE 

1997 NASA Aerospace Battery U 'orkshop -6- General / Ni-Cd / Ni, -MH Battery Ses 
- JPL 

I APPLICATIONS - FIRST SPACE MISSIO

cll 

It20 
NiIH, 

GEO 

I 

1997 NASA Aerospace Battery Workshop -8- 

1990 

1995 

2000 

GeneraI / Ni-Cd / Ni-MH Battery Session 

4/25/90 HST 

NASA LEO 

b 

611 0190 LEASAT 

Super NllCd 

1/25/94 CLEMENTINE SPV NiIH, 

LUNAR MAPPING 

1 125194 TU BSAT-B 2 Cell CPV 

STORE MESSAGES 

511995 CENTAUR Li-SOCI, 

28V, 250AH LAUNCH VEHICLE 

5/5/97 IRIDIUM-1 50Ah SPV 

LEO 34 TO DATE 

NOV 97 FLIGHT EXP Na/S 

WAKESHIELD 
+- 

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EVOLUTION OF 

NI-CD BATTERIES 

IN SPACE 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
4- 

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V, 

'4 

JPL 

5 

2 

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1 

I-CD SPACE BATTERY EVOLUTION 

4 

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m 

!$' 

3 

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% 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 

' Technology 

Performance 

Manufacturers 

1958-69 

3-6 Ah Cells 

Pelion 

GTM Seals 

2-5% DOD 

< 1000 cycles 

Leaks 

Const. I Charge 

Gould 

Gulton 

Saft 

Sonotone 

1970-79 

5-20 Ah Cells 

Teflonation 

Ceramic Seals 

NASA Std Cells 

10-20% DOD 

NASA VT 

EPI 

G.E.. 

Gulton 

Saft 

1980-89 

NASA 50 Ah 

E-l Process 

Lt Wt Designs 

Passivation 

>I 0 Years GEO 

40K Cycles LEO 

EPI 

G.E../Gates 

Saft 

Sanyo 

1990-97 

Super Ni-Cds 

Pellon 2536 

'MATURE' 

EPI 

Hughes 

Saft 

Sanyo 

Acme 
EVOLUTION OF 

NI-H2 BATTERIES 

IN SPACE 

ADVANCED TECHNOLOGY PROGRAM OFFICE 
4- 

% JPL 

( NI-H, SPACE BATTERY EVOLUTION 

Technology 

Manufacturers 

35-50 Ah IPV 

E-l Aq. Process 

Back to Back(C) 

Recirculating(AF) 

LEO 25% DOD 

2000 Cycles 

Polar 40% DOD 

EPI 

Yardney 

ERC 

SAFT 

GE 

E-l Alc. Process 

100 Ah 

40K Cycles LEO 

I0 YEARS GEO 

EPI 

Hughes 

JCI 

Saft 

Gates 

CPV & SPV 

120 Ah 

26% KOH 

50-100K LEO 

CYCLES 

EPI 

Hughes 

JCI 

Saft 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
4 - 

5 

JPL 

KEY EVENTS IN SPACE BATTERY HISTORY 1 

NI-CD 

LITHIUM 

'59 EXPLORER-6 

(CYLINDRICAL) 

'62 ARIEL-I LEO 

( PRISMATIC) 

'63 SYNCOM2 GEC 

'77 NTS-2 & AF 

FIRST IPV USE 

' '80 SOLAR MAX 

NASA 20 AH STD 

'82 LANDSAT -D 

STD 50 Ah 

INTELSAT 5 

IPV GEO 

'83 STS- LI-BCX 

'90 LEASAT GEC 

SUPER Nl-CD 

'90 HUBBLE IPV 

'94 CLEMENTINE 

SPV 

'94 TUBSAT CPb 

'95 CENTAUR 

'96 PATHFINDER 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
_I 

2 

S 

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NASA MISSIONS 

TODAY 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
4- 

'0 S 

JPL 

2 b 

2 

k 

a 1 

m 

!? LAUNCH MISSION NO. NO CELL CELL 

d 

=4 

DATE NAME TYPE BATS CELLSIBAT CAPACITY TYPE 

a 5 

43 

11/97 TRMM LEO 2 22 50Ah EPI-SUPER NI-CD 

12/97 TRACEISMEX LEOISS 1 22 9Ah EPI- SUPER NI-CD 

?I97 ACE LIBATION 1 18 12Ah NI-CD 

I 

w 

Y' 5/98 LANDSAT-7 LEO 2 17 50Ah EPI - AXIAL Ni-H, 

6/98 EOS-AM LEO 2 54 50Ah EPI-RE Ni-H, 

7/98 GOES-L GEO 2 28 12Ah GATES / SAFT NI-CD 

F 

3 9/98 WIREISMEX LEO I 22 9Ah EPI-SUPER NI-CD 

k 

\ 

3 ?I99 IMAGE LEO I 22 21Ah CS SUPER NI-CD \ 1 2 

3 ?I00 MAP FULL SUN I 1 112CELL 23Ah CPV EPI-J RE NI-H2 

t 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
L1- 

\n JPL 

NEAR TERM NASA MISSIONS - JPL I 

1 LAUNCH 9 MISSION NO. 

5 DATE 

0 

NAME TYPE BATS ' 1996 .a0 MGS MARSORB 2 

1996 MPF LANDER 1 

1996 MPF ROVER 3 I 

? 1997 CASSlNl PROBE 3 

1998 NEW MIL DS-1 1 

1998 NEW MIL DS-2 2 

k' 

z 1998 & MARS SURVEYOR198 2 

$ 1999 STARDUST 

'2 

2 

3 ' 1999 STARDUST SAMP. RET 3 

& 

NO CELL CELL 

CELLSIBAT CAPACITY TYPE 

I 112CELL 23Ah CPV EPI-J RE NI-H2 

18 40Ah AGIZN RECHARGE 

3 12Ah LI-SOCL, 'D' 

LI-SO, 

1 112CELL 12Ah SAME AS MSTI-3 

4 2Ah LI-SOCL, FLAT PLATE 

13 8AH LI-SO2 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
s - 

U 

JPL 

5 

2 

b 

a 

2 

1 

2 d LAUNCH MISSION NO. NO CELL CELL 

P 8 DATE NAME TYPE BATS CELLSIBAT CAPACITY TYPE B 

G 1998 AXAF LEO 3 22 CELLS 40Ah EPI IPVNI-H2 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
'r- 

..I 

JPL 

THE FUTURE OF 

BATTERIES IN SPACE 

ADVANCED TECHNOLOGY PROGRAMS OFFlCE 
- -- 

POTENTIAL NASA SPACE 

NASA MISSIONS AIR FORCE MISSIONS 

JPL LEO 

MARS LANDER AND ROVER -2001 NPOESS -2007 

MARS LANDER AND ROVER -2003 Surveill. Platforms 

MARS SAMPLE RETURN MISSION - 2005 SBlRS Low 

GSFC 

CHAMPOLION MISSION - 2003 

SOLAR PROBE - 2005 GEO 

Milsatcom - 2002 

DSP 

SATELLITE SERVICING TOOLS 

AIRCRAFT 

LIBATION POINT SPACRAFT 

(MAP-2000,NGST 2007) 

GEO SPACECRAFT(G0ES) 

LEO SPACECRAFT(E0S) 

AVIATION 2001 

UAVS- 2002 

JSC 

SHUTTLE APPLICATIONS 
PAST PRESENT 
Planetary Orbiters Planetary Lander GEO Spacecraft 

Planetary Rover Astronaut Equipment LEO Spacecraft 
TECHNOLOGY PROGRESSION 

I00 mAh 

< 200 cycles 

An od eWMsturL 

Electrolyte 

Cat. Mat. 

Separator 

Binder 

1-20 Ah Cells 

I00 Whlkg 

500 cycles 

i Cell Design i 

Battery ~esign 

I 

Database i 

Charge Control 

28-270V 

10-1 00 Ah 

I00 Whlkg 

2000 GEO & 

30,000 LEO 

CYCLES 
311 313A3 

1997 NASA Aerospace Battery Workshop -23- General / Ni-Cd / Ni-MH Battery Session 
k- 

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3 

JPL 

5 

t2 

h 

a 

b" 

OF NASA I AF LI-ION PROGRAM 

$ 

P 

x 

9 

q DEVELOP ADVANCED LITHIUM ION CELLS AND 

B 

4 SMART BATTERIES FOR AEROSPACE AND DOD 

APPLICATIONS 

ESTABLISH U.S. PRODUCTION SOURCES 

DEMONSTRATE TECHNOLOGY READINESS FOR: 

SATELLITE SERVICING TOOLS BY 1999 

ROVERS AND LANDERS BY 2000 

LIBATION POINT MISSIONS BY 2000 

GEO MISSIONS BY 2001 

MILITARY TERRESTRIAL APPLICATIONS BY 2001 

LEO MISSIONS BY 2003 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
4 

I 

8 

'4 

JPL 

5 

k? 

TECHNOLOGY CHOICES 

FOR FUTURE MISSIONS 

NEAR TERM !>2000!: 

LI-ION LIQUID ORGANIC ELECTROLYTE CELLS 

1 -20Ah CELL SIZES 

>I000 CYCLES 

SUPERIOR LOW TEMPERATURE PERFORMANCE 

COMMERCIAL APPLICATIONS USE 

LEVERAGE OF FUNDS SEVERAL PROGRAMS 

LONG TERM MISSIONS !>2007) 

LI - ION POLYMER 

HIGHER SPECIFIC ENERGY 

ADAPTABILITY TO SEVERAL CONFIGURATIONS 

ADVANCED TECHNOLOGY PROGRAMS OFFICE 
TECHNOLOGY APPROACH 

TO NASA I AF LI-ION PROGRAM 

DEVELOP ADVANCED ELECTRODE MATERIALS AND ELECTROLYTES TO 

ACHIEVE IMPROVED LOW TEMP. PERFORMANCE AND LONG CYCLE LlFE 

OPTIMIZE CELL DESIGN TO IMPROVE SPEC. ENERGY, CYCLE LlFE AND 

SAFETY 

ESTABLISH MANUFACTURING PROCESSES TO ENSURE PREDICTABLE 

PERFORMANCE 

DEVELOP AEROSPACE LITHIUM ION CELLS IN 10,20,50, AND 200 AH SIZES 

DEVELOP BATTERIES IN 28,100 AND 270 V CONFIGURATIONS 

DEVELOP ELECTRONICS FOR SMART BATTERY MANAGEMENT 

DEVELOP A PERFORMANCE DATABASE FOR VARIOUS APPLICATIONS 

DEMONSTRATE TECHNOLOGY READINESS FOR VARIOUS NASA AND AIR 

FORCE MISSIONS 
4 

'n 

'n 

'4 

5 

2 

b 

? 8 

R 

m 

'a f? 

3 

F 

THIS PAPER INCLUDED: 4 

A CHRONOLOGICAL HISTORY OF BATTERY FLIGHT 

L FROM 1959 TO THE PRESENT 

A LIST OF THE NEAR TERM FLIGHT MISSIONS FROM 

1997-2000 

9 

- \A PLAN FOR AN INTERAGENCY (NASA I AF) PROGRAM b TO DEVELOP LI-ION BATTERIES FOR PLANETARY, ? 

AVIATION, LEO AND GEO MISSIONS FROM 2000-2003 & 

'a 9 

d 
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tu 

B 

d 

T 

B 

Battery System Management 

.a0 

a Optimized Electrokinectic Control in 

tL 

Y= Galvanic, Electrolytic, and Storage 

Operating Modes 

B 

B d Floyd Williamson ZAE Research, Inc. 

r/, 

2 

s 

555 Sparkman Dr. Suite 450 Huntsville, AL 358 16 3 205-890-0507 / FAX 205-890-0857 email: flw@hiwaay.net 
Battery System Management 

Physics 

1 Inherent 

Discharge contro 

1 missing element / 

Charge control 
Faraday's law: 

Theoretical capacity: 

Coulom bic efficiency: 

Tafel equation: 

Total charge passed: 

The potential developed across the cell is equal to: 

V = qs (anode) + q, (anode) + IR - q, (cathode) - q, (cathode) 

The current density: 

The effective capacitance C, (Stern capacitor) of the double layer structure per the equation: 

Page 2 

1997 NASA Aerospace Battev Workshop General / Ni-Cd / Ni-MH Battev Session 
mportant concep 

Tafel equation shows the exponentia 

relationship between current and activation 

overpotentia 

Ions respond exponentially to the 

ication of DC currents. 

Exchange current is the measure of 

freedom from kinetic limitations. 

Overpotentials arization act to impede 

ectric fie d that is driving the reaction 

rate. 
I 

Current + 

Current + 

Two views of discharge polarization. 

1997 NASA Aerospace Battery Workshop General/ Ni-Cd / Ni-MH Battery Session r 
1997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Battery Session 
1997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Battery Session 
Inactive material 

Depleted material 

Depleted material 

Inactive material 

High interfacial 

surface area 

Less active 

material 

Cathode 

Separator 

Anode 

Cross-sectional view of spiral wound cell 

1997 NASA Aerospace Battery Workshop General / Ni-Cd/ Ni-MH Battery Session 
s 

-4 

Large surface area 

2 b 

E' 

Depleted 

Material 

Cathode 

Large material mass 

Small interfacial 

surface area 

Anode 

(+I Current Collector 

Cross-sectional view of a battery 
Char 

Disch 

I Ext. Circuit 

- Current collector 

Interstitial site 

.al hydride electrode 

NiMH Conduction 

Wion 

000 

Interstitial diffusion 

or migration 
Disch 

Cha~ 

Current collector 

Interstitial site 

Liathiated transition metal 

intercalation compound 

Liathiated carbon 

intercalation material 

Current collector 

Lithium Ion Conduction 

Li ' Ion 

ogo 

0 0 

Interstitial 

000 

diffusion 

or migration 
Random movement of particle with no electric field applied 

Electric force 

Random movement of particle under influence of a DC electric field. 

1997 NASA Aerospace Battery Workshop General / Ni-Cd/ Nj-MH Battery Session 
ion 

k' 

4. 

8 POROUS ELECTRODE 
dipole symbol 

Cation Anion 

o d x (nmf "Ik 

Dynamic nature of the electrical double layer 

1997 NASA Aerospace Batteiy Workshop -42- General / Ni-Cd/ Ni-MH Battery Session 
Double Layer 

IHP on the order of 

OHP on the order of 3 e 

capacitance 1 0uF/cm2 to 5 0uF/cm2. 

00 mV across nrn yields 06 V/m. 

The potentia is a kinetic resistance. 

Outside the He rnhortz layer the reactant 

species are too distant from surface to react. 
I Water Dipole dipole symbol I 

I Cation Anion I 

Anode 

Ch Cgc 

Cathode 

Cgc Ch 

Electrical double layer showing effective capacitance. 
- Current 1.1. ions 

Charge pulse 

- Current ..... ions 

Depolarization pulse 
Charge pulse removed 17 

EDL (IHP) field strength based on 7 

Ir] 

Field Strength 

ee=loomv 

Field Strength 

a R=-~OO~V 

0 

100% 

100 x lo6~/rn 

300 x 106v/rn 

1 

36.8% 

36.8 x to6v/m 

110 x 106v/rn 

2 

1 3.5% 

27 x 106v/m 

40.5 x 106v/m 

3 

5% 

5 x 106v/tn 

15 x 106v/rn 

4 

1.8% 

1.8 x IO~V/~ 

5.4 x 10~~/rn 

5 

0.7% 

0.7 x IO~V/~ 

2.1 x 106v/rn 

6 

0 

o 

0 
Current rate-of-rise 

1 second charge pulse with 15mS depolarization pulse 

Polarization resul tinq from - 

DC rate-of-rise 

q decays to 98.5% in 15mS 

e-X with x = 15mS 

'Is 

EDL voltage caused by the charge current 

time I I I I I I I I I I I I 

I 

5mS / DIV 

EDL with 1 second time constant 
Ideal ECpulseTM waveform 
Time A Time B Time C Time D Time E 

Electrical double layer at each time interval 

'QY6 

Water Dipole dlpole nymbol 

Cation Anion I 

I I I I 

A B D E I 

C I > time 

Electrical double layer reversal with ECpulseTM. 
Time 

- ECpulseTM EMF 

- ECpulseTM ion displacement 

- DC ion displacement - PRIOR ART 

ECpulseTM versus DC 
0 10 20 30 40 50 60 70 80 90 100 

Time 

- 

- ECpulseTM EMF w/ -0.1 A DC offset 

ECpulseTM ion displacement 

- DC ion displacement wl-0.1A DC load 

ECpulseTM versus DC 
-6 

0 10 20 30 40 50 60 70 80 90 100 

Time 

- 

- ECpulseTM EMF wl -0.5 A DC offset 

ECpulseTM ion displacement 

- DC ion displacement w/ -0.5 A DC load 

ECpulseTM versus DC 
NiMH 1200 mA-Hr Discharge 
NiMH 1200 mA-Hr 

Normal 
& 

'Q 

'0 

'4 

s 

2 

b 

a 

% 2 

ttr 

sa 5 

F 

Advanced Lithium-Ion ENergy 

B !z +i Technology 

Ricky J. Roberson 

Greg Tyler 

Huntsville Integrated Voltaic Equipment 

(HIVE) 
4 

'0 

\O 

-4 -. 

s 

2 

HIVE Goals: $ 

W 

P 

t! 

d 1. Develop modular, scalable Li-Ion support technology applicable to 

8 F any aerospace mission: 4 

low to high voltages 

low to high currents 

cln 

F high reliabilitylredundancy 

2. Produce an integrated prototype Li-Ion power supply to meet NASA 

9 

8 ProSEDS mission requirements: B 

\ 5 

Q '. 120 V DC, 3 Ah nominal 

2 3. Implement Li-Ion technology in another important aerospace project: =: in 

g . 

a YOURS 


3 

b 

b 

Ei 

d 

b 2 I?. 

0 

3 

HIVE Li-Ion Electronic Integration Approach: _. 

Use off-the-shelf 18650-size Li-Ion cells 

Avoid constant-voltage regulators to minimize capacity waste 

Stack Li-Ion cells to achieve required high voltages 

Isolate Li-Ion cells in stack from each other with solid-state relays (SSRs) 

Provide SSR-controlled alternate path for current flow around any given cell 

Implement microprocessor voltageltemp monitoring and SSR switching control 

Power local cell control electronics from upstream cells in stack 

Utilize redundant isolated Li-Ion cells to power electronics at top of stack 

* Switch out unneeded cells from stack for controlled system discharge 

Switch out all cells from stack for individual cell recharge 
a VIVE Li-Ion Electronic Integration Approach: % 

S 

- 

CHARCER BOOT-UP CNDN4 
HIVE Ma-jor Electronic Component Selection: % - 

2 

T 

iz 

3. 

4 Battery Cell: Moli Energy ICR-1.8650 Li-Ion 

I Temperature Sensor: Dallas Semiconductor DS1820 

E 

a Charger Control Chip: Maxim MAX846A 

Slave Microprocessor: $ 

National Semiconductor COPSACC \ 5 

2 -Relay: 

3 International Rectifier PVN012 

ki 

s 8 a And a variety of supporting capacitors, resistors, etc. 

3 B 
HivePartsList: 4 

8 

m 

PART 

1 LITHIUM ION BATTERY CELL 

2 TEMPERATURE SENSOR 

3 INTEGRATED CIRCUIT - BATTERY CHARGER 

4 MtCROPROCESSOR 

5 CRYSTAL - 4MHZ 

6 OPTORELAY 

7 OPTORELAY 

8 OPTORELAY 

9 OPTORELAY 

10 OPTORELAY 

11 TRANSISTOR 

12 DIODE 

13 RESISTOR - POWER - 0.15 OHM 3W 1% , 

14 RESISTOR - SMT - 680 OHMS 5% 

15 RESISTOR - SMT - 400K 1 % 

16 RESISTOR - SMT - 1 M 5% 

17 RESISTOR - SMT - 10K 5% 

18 RESISTOR - SIP - 8 PIN IS0 10K 2% 

19 RESISTOR - SIP - 10 PIN IS0 560 OHM 2% 

20 RESISTOR - SIP - 6 PIN IS0 IM 2% 

21 RESISTOR - SIP - 6 PIN IS0 18OK 2% 

22 RESISTOR - SIP - 10 PIN BUS IOOK 2% 

23 CAPACITOR - 

24 CAPACITOR - 

25 CAPACITOR - 4.7UF 16V TANT 

26 CAPACITOR - 4.7UF 35V TANT 

27 CAPACITOR - 10UF 16V TANT 

28 CAPACITOR - 22UF 35V TANT 

29 CAPACITOR - 0.1 UF CERAMIC 

30 CAPACITOR - O.OIUF CERAMIC 

31 CAPACITOR - O.01UF CERAMIC 

32 CAPACITOR - O.OIUF CERAMIC 

33 CAPACITOR - 33PF CERAMIC 

34 CAPACITOR - 33PF CERAMIC 

35 HEADER4 PIN 

36 HEADER-? PIN 

37 CONNECTOR4 PIN 

38 CONNECTOR-7 PIN 

39 SOCl(FT 

18850 

DS1820 

MAX846A 

COP8ACC 

X405 

PVNOI 2 

PVN012 

PVNOI2 

PVN012 

PVN012 

ZTX749 

FR301 CT I 

43F0.15 

P680VCT 

P402KFCT 

PI .OMVCT 

PI OKVCT 

750-83-RI OK 

750-103-R560 

750-63-R1 MEG 

750-63-R180K 

750-1 01 -RlOOK 

POWER 

ZETEX 

VSET 

RESET 

ISET 

LED 

ADC 20-3 

ADC 5-2 

ADC 

VSET 

ISET 

BATT 

COP8 

RESET 

MAX 

CS+ 

ZETEX 

CCI 

ccv 

XTAL 

XTAL 

GROUND 

PWRICOMM 

GROUND 

PWRICOMM 
G - Y. ."*~---k.-.~~-.-~~~- .---..I. 

3 !dnis~~~~l~~fl~~ PROOUCT DATA SHEET -. 

5 

5.2 ~NFORMATION MODEL. ICR-I 8650 

k HIGH ENERGY 

L ITHIUM-ION RECHARGEABLE BATTERY 

- 

Nominal Voltage 

Capacity 

r 

Energy 

- 

Weight 

Energy Density 

Volumetric 

Gravimetric L 

STORAGE SPECIFICATIONS 

1 i 

3.7 V 

1 500 mAh 

5.6 Wh 

42 grams 

325WM 

132 WM<S 

A: 18.3mm 6: 65mm 

OPERATING SPECIFICATIONS 

( Temperature ~ange . 1 -20 *C to 60-C I 

Operating Voltage 

Charge Voltage 

Cutoff Voltage. 

Temperature Range 

Discharge 

Charge 

Maximum Discharge Current 

(Continuous) 

Maximum Charge Current 

I Recommended Voltage Range \ 1 4.2 V to 2.5 V 1 

4.2 V to 3.0 V 

4.2 V 

2.5 V 

-20-C to 60-C 

O°C to 45% 

3.0 A 

. 

1.5 A 4 
-.- - -- 

Discharge Characteristics . 

*.-,Cm".,m.,rm..-,n 

1.3QAh Rated Capacity 

0.6 0.8 

Capacity (Ah) 
k 

- 

z 

Gl l--'-‘---- Discharge Characbristics 

MQLICEL @ ICR486.5Q Call 

a 

m 

1 +30Ah Rated Capacity 0.20A Discharge 
DS1820 

I-wirerM Digital Thermometer 

FEATURES 4 PIN ASSIGNMENT 

* Unique l-WireTM interface requires only one port pin 

for communication 

Multidrop capability simplifies distributed temperature 

sensing applications F BOrrOM VIEW 

Requires no external components 

Can be powered from data line 

@ Zero standby power required 8 

% 

\ 

Measures temperatures from -55°C to c125OC in 

0.5OC increments. Fahrenheit equivalent is -67°F to 

+257"F in 0.9"F increments & UUU 

1 * Temperature is read as a +bit digltal value. g88 Q 3 3 " 2 Converts temperature to digital word in 200 ms (typ.) OSI 

6' PR35 PACKAGE 3 See Mech. Drawings * User-definable, nonvolatile temperature alarm set- sedi~ 

tings 

Alarm search command identifies and addresson 
' Battery-Charger System 

I@ (LOW - TURN OFF CHARGE) 

1, (HGH - DISABLE FLOAT V) 

PWWMC (CONTROL FLOAT V) 

PWMMC (CONTROL CWE I) 

MICROCONTROLLER 
EJ. National Senzic'onduct~r 

COP8ACC7 &Bit One-Time Programmable (OTP) 1' 

Microcontroller with High Resolution 

AID Conversion Schmitt Trigger inputs on porls G and L 

I Packages: 28 DIPIS0 with 23 110 pins 

20 SO with '15 110 pins 

General Description 

The COP8ACC7 is a member df the COP8* 8-bit OTP mi- CPUAnstruction Set Features 

1 lps instruction cycle time crocontroller family. It is pin and software compatible to the mulli~source wctored inter,.,,pt sewicing mask ROM COPBACCS product family. (Continued) - External lnterrupl 

.~ey Features - ldle Timer TO 

- Timer TI associated Interrupts 

halog Function Block for high resoldion AID including - MICROWIRUPLUS 

- Analog comparator with seven Input muxes - Mulli-Input Wako Up 

- Constant Current Source and VccR Reference - Software Trap 

- 16-bit capture timer (upcounter) clocked from CKl with - Default VIS 

auto-reset on timer start-up - ND (Capture Timer) 

Quiet design (reduced radiated emlsslons) Versatile and easy to use lnstwction set 

1 4096 bytes on-board OTP EPROM with security feature I 8-bit Stack Polnter (SP) - dack In RAM 

1 128 bytes on-board RAM I Two &bit Registers indirect Data Memory Pointersv (B and X) 

Additional Peripheral Features 

I ldle Timer 

One 16-bit timer wlth two 16-bit registers supporting: 

- Processor Independent PWM mode 

- External Event counter mode 

- Input Capture mode 

1 Multi-Input Wake-Up (MIWU) wilh optlonal interrupts (4) 

4 WATCHDOG and dock monitor logic 

I MiCROWlRUPLUS serial 110 with programmable shifl clock 

polarity 

Fully Static CMOS 

IC Low current drain (typically < 5 pA HALT current) 

R Two power saving modes: HALT and IDLE 

Slngle supply operation: 2.7V to 5.5V 

r Temperature ranges: O'C to +70*C, -40'C to +85'C 

Development System 

Emulation device for COP8ACC5 

I Real time emulation and full program debug offered by 

MetaLink development system 

UO Features 

Memory mapped 110 Applications 

Software selectdble 110 optlone (Push-Pull Output, Weak Pull- . bat%^ Chargers . . 

Up Input, High impedance input) I Appliances 
I Smart Battery Charging 

I with COP8ACC 

Natlonal Semkondudor 

Application Note 1075 

Sayeed Rahman 

May 1997 

' TABLE 1. Clock Spwd w Rviolutlon, Convvmlon Tlmv, and AchlevaMe Accuracy 
MOSFET Photovoltaic Relay 

Micrmlmtr~nic 

Power IC Relay 

Sir~gle Pole, Normally Open, 0-20V, 2.5A AC/ 4.54 DC - - - 

PVN012 Features 

100rna On-Resistance m 

GenV HEXFET output m 

Bounce-free operation 

2.5 - 4.5 Amp capacity m 

Linear AC/DC operation m . 4,000 V, I/O isolation 

Solid-State reliability rn 

, UL recognized and CSA certified 
- 

a HIVE Li-Ion Mechanical Integra-qeproach: % 

R 

(D 

3 

.r! Prepare single epoxy "honeycomb" structure to hold Li-Ion cells 

F 9 8 

% Sandwich "honeycombv, Li-Ion terminals between two circuit boards 

Size chargerlcontrol circuitry to fit between these "power" boards 

Assemble via modular chargerlcontrol boards plugged into "power"boards 

9 a 

Utilize through-hole, allowing repeated UP reprogramming 1 reinsertion 

9 - a 

\ 

5 Encase assembly within lexan plastic shell 2 

HIVE Power Board Layout: 

Size = 12 x 12 inches 
HIVE Li-Ion ChargerIControl Circuit: 

Front 

Size = 2.50 inches high by 0.85 inches wide 

- Back 
HIVE Li-Ion ChargerIControl Circuit Parts Placement: 

Front 

RELAY 

RELRY 

RELFlY 

RELFlY 

RELAY 

CCC 

UP - Back 

Size = 2.50 inches high by 0.85 inches wide 
HIVE Li-Ion ChargerIControl Module: I 

Size = 2.50 inches hieh by 9.00 inches wide (qty 6) 

or 2.50 inches high by 10.70 inches wide (qty 2) 

..... ......... .... ... __-_...__._-_l.--..._. __ ----.---- ......... ........................ ,/ ;.\ .- - - ......... 

. . 

BETWEEN CONDUCTORS 

UNCOATED BOARDS 

0- 10,000 feet altitude 

UNCOATED BOARDS 

Above 70,000 feet 

I I 500 .i- .00012 in./V 

Minimum conductor spacing as a function of interconductor potential. (MIL-STD-275C) 
Specs Sheet firr RedcoTM Phenolic (High Pressure Imninates) 

(Dielectric strength (vpm 1~229 1 

.* .. P50 1200 1- 1300 - (550 
HIVE ProSEDS Point Design Solution: 

Power: Dual 0-120 V DC, 1.5 A busses 

Energy: 500+ Whr 

Weight: 19 pounds / 8.5 kilograms 

Volume: 0.25 cubic feet / 7.0 liters 

Size: 12x12~3 inches / 30x30~8 cm 

Ratios: 60 Whrtkg for integrated power system 

72 Whrniter for integrated power system 

Features: Individual voltage I temp monitoring on each of 84 Li-Ion cells 

Digital data storage / dump of complete discharge / recharge cycle 

Integrated non-metallic structure minimizes weight and volume 

Integrated control electronics for both charging and discharging 

Independent, adjustable power bus voltages 

Advanced voltage regulation minimizes energy capacity waste 
a - HIVE* Puture'Efforts: 

U 

B Build/program/test prototype outlined here s Z 

9 

P * Generate flight certification/documentation for components !z 

% 

Designlbuildlflight-qualify "Mod 2" unit with NASA/MSFC oversight on 

Investigate DCIDC converter modules to increase output currents 

9 = Implement mods allowing structure "stacking" to increase output currents 

il 

& '-. Develop ongoing joint partnership efforts with other potential users/projects 2 

\ 
Page intentionally left blank 
PERFORMANCE OF NICKEL-CADMIUM BATTERIES ON 

THE GOES I-K SERIES OF WEATHER SATELLITES 

Sat P. Singhal 

Computer Sciences Corporation 

Walter G. Alsbach 

Jackson & Tull 

Gopalakrishna M. Rao 

NASAIGoddard Space Flight Center 

Abstract 

The US National Oceanic and Atmospheric Administration (NOAA) operates the Geostationary 

Operational Environmental Satellite (GOES) spacecraft (among others) to support weather 

forecasting, severe storm tracking, and meteorological research by the National Weather Service 

(NWS). The latest in the GOES series consists of five spacecraft (originally named GOES I-M), 

three of which are in orbit and two more in development. Each of five spacecraft carries two 

. Nickel-Cadmium batteries, with batteries designed and manufactured by Space Systems Loral 

(SSLL) and cells manufactured by Gates Aerospace Batteries (sold to SAFT in 1993). The 

battery, which consists of 28 cells with a 12 Ah capacity, provides the spacecraft power needs 

during the ascent phase and during the semi-annual eclipse seasons lasting for approximately 45 

days each. The maximum duration eclipses are 72 minutes long which result in a 60 percent depth 

of discharge @OD) of the batteries. This paper provides a description of the batteries, 

reconditioning setup, DOD profile during a typical eclipse season, and flight performance from the 

three launched spacecraft (now GOES 8, 9, and 10) in orbit. 

INTRODUCTION 

The US National Oceanic and Atmospheric 

Administration (NOAA) operates the 

Geostationary Operational Environmental 

Satellite (GOES) spacecraft (among others) 

to support weather forecasting, severe storm 

tracking, and meteorological research by the 

National Weather Service (NWS). The latest 

in the GOES series consists of five 

spacecraft (originally named GOES I-M), 

three of which (GOES 8, 9, and 10) are now 

I997 NASA Aerospace Battery Worhhop 

in orbit. National Aeronautics and Space 

Administration (NASA) is responsible for 

procurement, launch, and checkout of these 

spacecraft before turning them over to 

NOAA for operational use. They were built 

by Space Systems Loral (SSL) with 

components from many vendors. 

A key element in overall mission success is 

the successfbl operation of the battery 

system. These spacecraft carry two NickelCadmium (Ni-Cd) batteries designed and 

manufactured by SSL with cells 

manufactured by Gates Aerospace Batteries 

General / Ni-Cd / Ni-MH Battery Session 
(sold to SAFT in 1993). The battery, which 

consists of 28 cells with a 12 ampere-hour 

(Ah) capacity each, provides the spacecraft 

power needs during the ascent phase and 

during the semi-annual eclipse seasons 

lasting for approximately 45 days. The 

maximum duration eclipses are 72 minutes 

long which result in a 60 percent depth of 

discharge @OD) of the batteries. 

This paper provides a description of the 

batteries including the design and 

manufacturing process through acceptance 

testing and pre-launch preparation 

(References 1-4). Results are also provided 

from life cycle testing at the Naval Surface 

Warfare Center, Crane (Reference 5). The 

tests used battery packs prepared from cells 

out of the batch manufactured for spacecraft 

use. We also provide on orbit performance 

data fiom ascent phase, reconditioning, and 

from the eclipse seasons for the three 

spacecraft in orbit. 

SPACECRAFT 

The GOES 8, 9 and 10 spacecraft are threeaxis stabilized geostationary satellites. Their 

dry mass is 977 kg and the nominal main bus 

power is 1 150 W. They carry two main 

scientific instruments (Imager and Sounder) 

and a number of other measuring devices. 

During sunlight, they are powered by a single 

wing, two-panel solar array (SA). The SA 

has an output of 1300 W at beginning of life 

(BOL) Equinox and 1050 W at end of life 

(EOL) Summer Solstice. During eclipse, 

two 12 Ah Ni-Cd batteries with 28 cells each 

sustain the spacecraft. Power transfer fiom 

the batteries to the spacecraft bus is achieved 

through diode coupling. The power control 

unit (PCU) is the principal element for 

management and control of spacecraft 

primary power. The power control 

electronics (PCE) consists of the PCU, one 

1997 NASA Aerospace Battery Workshop 

sequential shunt unit, and four electroexplosive device (EED) extension units. Key 

features of the PCE are as follows: 

Provides functions required for the 

primary bus. 

Provides direct energy transfer of SA 

power to the power distribution bus. 

Regulates and limits the voltage of 

the primary power bus to 42 + 0.5 V 

dc during sunlight operation. 

Minimizes primary bus ripple and 

voltage transients by use of SA 

regulation and primary bus filtering. 

Allows flexibility of battery-charge 

control to optimize battery energy 

balance, thermal control. 

Maintains minimum battery 

temperature control through use of 

thermistors and heaters in the 

associated batteries. 

Includes provisions for individual 

battery reconditioning. 

Provides fail-safe and redundant 

actuation control of spacecraft EED 

(pyrotechnics) functions. 

Eliminates single-part failure 

criticality by use of circuit 

redundancy, protective functions, 

fault protection of spacecraft heater 

loads, and alternate mode operation 

selectable by command. 

Permits operational flexibility and 

status monitoring of key subsystem 

parameters by command and 

telemetry functions. 

Redundancy Provisions. Power 

conditioning functions use both active and 

commandable redundancies. Commandable 

redundancy is provided for the following 

PCE functions: 

a. Battery discharge control 

b. Battery temperature control 

c. Voltage telemetry monitors 

General / Ni-Cd / Ni-MH Battery Session 
d. Battery relay commanding 

e. Battery charge control 

f EED ignition control (3 levels) 

g. Battery reconditioning control 

Some important functions for the battery 

control are listed below. 

Battery Charge Control. The battery 

charge control configuration provides 

capability of up to eight discrete battery 

charge rates. Battery charge current is 

supplied by six charge control arrays located 

on the SA wing. Within the PCU, charge 

control array sections A, B, and C are 

connected through relays and isolating 

diodes to battery 1. A similar arrangement is 

provided for charge control arrays D, E, and 

F for battery 2. 

Battery Temperature Control. The battery 

temperature control circuitry provides dual 

modes to connect and disconnect the heater. 

A total of four battery temperature controls 

are provided, one for each half of each 

battery. Automatic control of minimum 

battery temperaturee at 5 (2) f 1 OC is 

provided by heaters integrated with each 

battery assembly and separate temperature 

controllers in the power control unit. A 

precision thermistor on each battery provides 

temperature feedback to these controllers. 

Manual ovemde allows the heaters to be 

switched on or off by command. 

Value in parenthesis for GOES 8 only. 

Battery Reconditioning. Individual 

reconditioning of each battery is provided by 

command. A parallel group of resistors 

mounted external to the power control unit 

provide the reconditioning load. 

Commanded Load Control. Application of 

power to individual loads is through an 

on/off control input to each load dc/dc 

I997 NASA Aerospace Battery Workshop 

converter and by direct power bus switching 

of non-electronic loads (heaters). 

Eclipse Load Controller. The GOES PCU 

contains redundant load controller functions. 

Each load controller monitors the SA current 

and the shunt current. Each primary power 

bus load (excepting command functions) is 

connected to and disconnected from the 

primary bus by command. In addition, 

power is automatically removed from 

sunlight loads upon eclipse entrance and 

automatically restored upon eclipse exit. 

Command ovenide of this automatic 

function is provided. 

BATTERY DESIGN 

DESCRIPTION 

The battery subsystem of the GOES 

spacecraft consists of two assemblies, each 

containing 28 series-connected cells with a 

nominal capacity of 12 Ah and weighing 

approximately 12.9 kg each. Battery cell 

design has been guided by the life and 

reliability requirements of a 5-year geosynchronous satellite application and 

incorporates features that reduce the effects 

of unavoidable Ni-Cd cell degradation 

modes. Figure 1 shows one of the two 

batteries (SN 206) used on GOES-9. 

General / Ni-Cd / Ni-MH Battery Session 
Figure 1. 12 Amp-hour NiCd Battery (S/N 206) used on the GOES3 Spacecraft 
Electrical Design. The nominal capacity of 

each battery assembly is 12 Ah. The 60% 

DOD design limit for maximum duration 

eclipse operation represents load capabilities 

of 409 W at BOL and 398 W at five years, 

based on average cell discharge voltages of 

1.245 and 1.22 V, respectively, and a 1.0 V 

loss in the PCU. The 28 cells in series 

provide compatibility with charge voltage 

limits and deliver typical discharge voltages 

of 32.4 to 36.0 V. 

Reconditioning. Reconditioning of the 

batteries during the one-month period 

preceding each eclipse season is required to 

maximize subsequent voltage performance 

and consequently minimize DOD. 

Circuit Reliability. Circuit reliability within 

the battery is achieved by redundancy in 

wiring. Series connection between cells is 

provided by two parallel-connected, stranded 

copper wires soldered to the cell terminal 

lugs. Battery power connections are made 

to the terminals at the end of the cell series 

string by four redundant wires leading to the 

battery power connector. 

Telemetry. The battery wiring harness 

design includes provision for telemetry of 28 

individual cell voltages and battery 

temperatures. Additionally, overall battery 

voltage and current signals can be sensed. 

Mechanical Design. The structural and 

mechanical design of the GOES battery is 

optimized to efficiently perform two 

important functions: 

a. Cell support and restraint 

b. Battery to equipment platform mounting. 

Cell Support and Restraint. The basic 

mechanical design approach involves the 

restraint of two rows of 14 cells between 

1997 NASA Aerospace Battery Workshop 

two endplates, with each group of four cells 

supported by a rib structure. The key 

component to the battery design is the 

concept of supporting four cells on one rib, 

and compressing seven of these 

subassemblies between two endplatehie-rod 

assemblies. The design of the rib positively 

holds the corners of the prismatic cells in 

place. The compression force exerted by the 

endplates on the cells is controlled by the 

torque applied by the two tie rods. The cells 

are pre-loaded to 40 1bf7in2 in this manner. 

The end plates are designed to withstand a 

nominal cell overcharge pressure of 75 1bfinZ 

with minimal deflection. The endplates also 

provide mounting flanges for supporting the 

two connectors. The endplates are machined 

from heat-treated 7075-T73 aluminum, 

which alleviates stress corrosion issues, and 

the tie rods are titanium. The cell support 

ribs are cast from A357-T61 aluminum. 

Thermal Design. The thermal design of the 

GOES battery is such that temperature 

control is both passive and active. The 

passive control is primarily conductive 

through the cell support ribs, and the active 

control is via resistive heaters mounted on 

the same ribs. The seven aluminum ribs are 

sized to effectively conduct the heat 

dissipated by the four adjacent battery cells 

to the equipment platform of the spacecraft. 

The design is such that temperature gradients 

are only 3 OC. The ribs also have a mounting 

flange for the resistive heater elements. 

These flanges are centrally located such that 

an even distribution of heat throughout the 

battery assembly is achieved at times when 

the heaters are activated. The seven heaters 

are sized to dissipate the required heat for 

thermal control. Temperature sensing for 

each battery is accomplished by four 

precision wafer thermistors located on the 

battery assembly. Battery flatness is 

maintained within 0.020 inch during 

General / Ni-Cd / Ni-MH Battery Session 
assembly to allow for uniform thermal 

contact with the spacecraft. A silicone 

thermal grease is used between the battery 

and equipment platform on the GOES 

spacecraft to krther reduce gradients and 

contact resistance. 

Cell Design Summary. The Ni-Cd cell is 

designed to maximum margins for a greater 

than 5-year orbital lifetime. The prismatic 

Ni-Cd cells are packaged in a thin wall 

(0.012 in) 304L stainless steel container. A 

low profile terminal sealfcover configuration 

is used with double ceramic alumina-to-metal 

seals. Eleven positive and 12 negative plates 

are used. The specified negative-to-positive 

electrochemical capacity ratio is 1.70: 1. This 

ratio ensures that the cells will remain 

positive limited during the 5-year mission. 

The positive electrode group has a 

theoretical electrochemical capacity of 15.6 

k1.2 Ah and the negative group 26.4 53.6 

Ah. Active material loading is 1.25 H.06 

and 1.55 50.065 for the positive and 

negative plates, respectively. The negative 

plates are treated with Teflon to reduce 

cadmium migration and increase the quantity 

of electrolyte used within the cell. The 

balance of positive and negative loading, 

electrochemical capacity ratio, and negative 

plate treatment enhance long-term cell 

performance required for the 5-year GOES 

mission. Non-woven Pellon 2536 nylon 

filament material approximately 0.007-0.009 

inch thick is used for the separator. Weight 

concentration of 3 1% potassium hydroxide 

electrolyte provides for the transport of ions 

between the electrodes. This combination is 

being used for sealed Ni-Cd cells operated 

over the temperature range of 0 to 30 "C. 

Carbonate and nitrate levels are maintained 

at less than 2.0 gA and 1.0 mgA, respectively. 

Interfaces. The battery interfaces are rigidly 

controlled during battery assembly. The 

1997 NASA Aerospace Battery Workshop 

.. 

battery mounting surface and mounting hole 

pattern are controlled within close tolerances 

for an assembly of this type. Flatness of the 

battery is maintained within 0.020 inch, and 

the mounting hole pattern is drilled within 

0.020 inch. Two electrical connectors are 

provided on the battery assembly. The 

power connector is an eight-socket 

connector through which the battery is 

charged and discharged. Four bus wires 

each are connected to the positive and 

negative terminals. The second connector is 

a 50-socket connector. Through this 

connector the interface is provided for 

individual cells voltage sensing, thermistor 

measurement, and heater bus. 

Battery and Cell History. The following 

designations refer to the flight batteries. 

Spacecraft Battery Serial Number 

GOES-I 204,203 

GOES-J 205,206 

GOES-K 207,208 

Life-test. As with most NASA spacecraft, 

life-testing was started at Crane Naval 

Surface Warfare Center. A pack of cells for 

each satellite is undergoing real time or 

accelerated life-test. The test conditions are 

close to those in orbit with a 6.0 A discharge 

rate (equivalent to Cl2) and 0.9 A charge 

rate followed by a trickle charge period. 

Before entering a shadow period a 

reconditioning cycle is run. The cycles 

shown are (Figures 2-4): 

GOES I. Shadow Periods 2, 12, and 25; 

GOES J. Shadow Periods 2,9, and 25; 

GOES K. Shadow Periods 1 and 10. 

General / Ni-Cd / Ni-MH Battery Session 
82278 SHADOW 2. AVERAaE CUI. VOLTS 

62278 SHADOW 12. AVERAGE CELL VOLTS 

Figure 2. Battery DOD and Average Cell Voltages for Battery Pack 6227B 

(GOES-8) during shadow Periods 2,12, and 25 

1997 NASA Aerospace Battery Workshop General / Ni-Cd/ Ni-MH Battery Session 
6227C bl(ADOW 2. AVOUS CPI VOLTS. 

VOLTAGE UYCT LOW CVUES 72 48 

L~A.-. =%-_I .---- 

6227C sWIDOW 0. AVWAOE CaL VOLTS 

Figure 3. Battery DOD and Average Cell Voltages for Battery Pack 6227C 

(GOES-9) during Shadow Periods 2,9, and 25 

1997 NASA Aerospace Battev Workshop General / Ni-Cd / Ni-MH Battev Session 
GOESK SHADOW 1, AVERAGE CEU VOLTS 

moglEP~ 

m 0 AVERAGE . % DoD 

GOESK SHAWIcY 10. AVBlUQE CEU VOLTS 

TIIIR5SIECFigure 4. Battery DOD and Average Cell Voltages for Battery Pack GOESK 

(GOES-10) during Shadow Periods 1, and 10 

I997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Battery Session 
All curves show nominal behavior and are 

within specscation. There is also good 

agreement with the recorded flight data. 

Please note that shadow period 25 

corresponds to 12.5 years of mission, which 

is more than twice the specified lifetime. 

Manufacturing. The cells used for this 

battery were manufactured by Gates fiom 

1989 to 1993, designated as lot # 5 with 

serial numbers 12AB3 1 and 12AB35. Based 

on an original 9 Ah design, they were nameplated as 12 Ah cells with specified 

acceptance capacity at 12.6 Ah. Out of the 

500 originally mandctured cells 20 were 

taken into a 500 cycle test to address 

problems in two Aerospace Alerts. One alert 

fiom 1985, reported problems of negative 

electrode Mure; and other, fiom 1988, 

reported problems in the hot gas sinter 

process and with the Pellon 2536 separator. 

Storage. The cells were put in dry storage 

in a controlled environment 

Acceptance Testing. The basic acceptance 

testing consists ofseven steps (listed below), 

some of which were repeated until required 

results were achieved. 

1. Reconditioning 

2. Voltage Recovery 

3. Capacity(20 OC): 12.8 Ah at 1.51 V max 

4. Capacity(l0 OC): 12.0 Ah at 1.52 V max 

5.Capacity(0°C): 11.5A.h at 1.54Vmax 

6. Conditioning (20 "C): 12.8 Ah at 1.5 1 V 

7. Voltage Recovery 

Speciiied key values and acceptance test 

results for the GOES-I, -J, and -K batteries 

are listed in Table 1. 

Table 1. Cell Acceptance Data for GOES-I, -J, -K Batteries. 

1997 NASA Aerospace Battery Workshop -92- General / Ni-Cd /Ni-MH Battery Session 

GOES-K 

6/7/93 - 

511 1/94 

<1.479V 

11.9Ah - 

12.9A.h 

<1.52V 

12.0Ah - 

13.1 Ah 

< 1.53 V 

12.1Ah13.0 Ah 

Step 

3. Capacity at 20 OC 

MaxVoltage 

Capacity 

4. Capacity at 10 OC 

Max Voltage 

Capacity 

5. Capacity at 0 OC 

Max Voltage 

Capacity 

6. Conditioning at 20 OC 

MaxVoltage <1.51V 

7. Voltage Recovery > 1.145 V N/A N/A 

- 

> 1.190 V 

Requirement 

<1.51V 

> 12.8 Ah 

< 1.52 V 

> 12.0 Ah 

< 1.54 V 

> 11.5 Ah 

GOES-I 

9120188 - 

5/20/90 

<1.485V 

12.8Ah - 

13.5 Ah 

<1.510V 

13.0Ah - 

14.3 Ah 

< 1.530 V 

12.7Ah13.8Ah 

GOES-J 

7/16/90 - 

7/6/92 

<1.490V 

12.8Ah - 

13.2Ah 

<1.514V 

13.4Ah - 

14.5 Ah 

< 1.515 V 

12.7Ah14.3 Ah 
BATTERY OPERATION 

The two 12 Ah rated Ni-Cd batteries 

connected to the primary bus via redundant 

diodes and relays are designed for operation 

at a 60% maximum DOD for the mission. 

Criteria for battery management, including 

power subsystem design features to 

implement the same, are discussed below. 

Functions provided for each battery by the 

power subsystem to monitor, control, and 

protect the batteries are as follows: 

Battery voltage monitors 

Individual cell voltage monitors 

Battery charge current monitors 

Battery discharge current monitors 

Battery temperature monitors 

Battery heater status monitors 

Battery relay status monitors 

Battery reverse current monitors 

Battery relay control commands 

Battery charge rate control 

commands 

Battery reconditioning controls 

Battery charge voltage limiting 

Thermostatically controlled battery 

heaters 

Pre-launch and Launch. During the prelaunch and launch phases of the mission, the 

spacecraft is kept on external power until 

approximately 4 minutes before launch. 

Should a launch hold of more than 10 

minutes be encountered while the spacecraft 

is on mternal power, consideration should be 

given to recharging the batteries before 

continuing with the launch operation. 

Battery Charge Control. Batteries can be 

charged in either a continuous or sequenced 

mode. The primary consideration on the 

1997 NASA Aerospace Battev Workshop 

following recommended battery charge 

control implementation is to minimize battery 

stress while providing adequate charge 

return to ensure energy balance. This 

objective is satisfied using the lowest 

practical charge rate in the sequenced mode 

to minimize average battery temperature. 

The methodology recommended has been 

Dight proven on previous SSL programs and 

verified by life cycle testing. 

General Configuration Control. The 

electrical power subsystem permits 

considerable flexiiility in battery charge 

management. Charge power is derived fiom 

the SA using two identical groups of three 

charge control arrays. Group 1 consists of 

arrays A, B, and C; group 2 of arrays D, E, 

and F. The typical current output of these 

arrays is summarized in Table 2 for various 

conditions. 

Table 2. Battery Charge Array Current 

Two basic charge modes are available. First, 

with the normal recommended synchronous 

orbit sequenced mode, current is applied to 

General / Ni-Cd / Ni-MH Battery Session 
the two batteries in an alternating fashion in 

an approximately 10-minute cycle. Any 

combination of the six charge arrays can be 

switched between the batteries. Alternately, 

any combination of arrays A, B, and C can 

be used for battery 1 and any combination of 

arrays D, E, and F for battery 2. Second is 

the continuous charge mode. In this mode 

only the latter arrangement is practical. 

For one time only, a maximum DOD of 70% 

would be allowed for a battery temperature 

maintained between 5 OC and 25 OC. This 

maximum DOD, if approached, allows no 

margin for contingencies. To provide such a 

marein, a mission goal for maximum DOD 

should not exceed 60%. 

Normal Eclipsed Orbits Charge Control. 

Battery charge control performance is 

evaluated by use of the following telemetry 

provided for each battery: 

a. Battery voltage 

b. Individual cell voltages 

C. Charge current 

d. Battery temperature 

e. Battery heater status 

The 110% charge return at the firll rate, 

followed by trickle charge until the next 

eclipse, ensures sufficient recharge for all 

eclipse cycles. 

Battery Temperature Control. Battery 

temperature profiles are largely governed by 

the battery charge profiles. Thennostatically 

controlled 19.5 watt battery heaters are 

provided to maintain the minimum battery 

temperature* above +4 (1) OC. These heaters 

are designed to cycle on at +5 (2) +1 OC and 

off at +9 (5) +1 OC. Heater operation is 

controlled by a precision thermistor in each 

battery in conjunction with level detectors 

and relay drivers within the PCU. This 

fimction can be ovemdden by command to 

1997 NASA Aerospace Battery Workshop 

either turn the battery heater on or off 

regardless of battery temperatures. The 

intended use of this override function is to 

provide manual heater control in the event of 

a failure in the automatic heater control. 

* Values in parenthesis for GOES 8 only. 

Battery Reconditioning. To maximize 

battery discharge voltage during each 

eclipse, both batteries are reconditioned prior 

to the start of each eclipse season. These 

reconditioning cycles are performed as close 

as possible to the next solar eclipse to obtain 

the maximum benefit. 

General / Ni-Cd / Ni-MH Battery Session 
ON-ORBIT PERFORMANCE 

First of the GOES series I-M spacecraft was 

launched on April 13, 1994, and was 

renamed GOES 8 after achieving nominal 

operational orbit. The next two were 

launched on May 23, 1995 (GOES 9) and 

April 25, 1997 (GOES 10). The batteries 

provide spacecraft power needs fiom just 

before launch to the time of partial SA 

deployment soon after launch (the outer 

panel is deployed approximately ten days 

later providing fill power). The batteries also 

shoulder the power needs during eclipses, 

any maneuvers causing loss of solar power, 

and whenever the power produced falls short 

of the spacecraft needs. The sections below 

discuss battery performance during these 

various phases. Specific data is generally 

provided for GOES 10. However, data fiom 

GOES 8 and GOES 9 is also included 

whenever this data was usekl to establish 

trends. 

Ascent phase Battery Performance 

The batteries on the GOES spacecraft 

provide power to the primary bus starting a 

few minutes before launch when the 

spacecraft is switched to internal power. 

This first round of battery support is 

completed when the SA is partially deployed 

approximately 1.5 hours after launch. During 

the next few days, the batteries provide the 

spacecraft power needs during special events 

such as the magnetometer boom and full SA 

deployment phase and the dipole estimation 

phase. Also depending upon the time in orbit 

before the Apogee Maneuver Firing No. 1 

(AMF #I), the batteries provide support 

during several eclipses lasting approximately 

15 minutes each. 

For all cases when the batteries experienced 

a non-zero discharge current, data was 

1997 NASA Aerospace Battery Workshop 

retrieved from the archives at 5.12 second 

intervals. The charge removed during 

discharge (D) from the batteries was then 

calculated as an integral of the discharge 

current over time, i.e., 

The integral is evaluated numerically by 

using current values at 5.12 sec intervals. 

The DOD is then given as a percentage 

fraction of the nameplate capacity (Co) of 

the batteries (12 Ah each for a total of 24 Ah 

for the system). Thus 

gives the DOD value for the individual 

battery or the system as a whole depending 

on the values of the current and capacity 

used above. 

Table 3 lists all times during the ascent phase 

when the batteries on the GOES-K 

spacecraft were subjected to discharge. The 

maximum DOD was recorded during the 

Magnetometer boom and SA deployment 

phase; corresponding battery discharge 

currents are shown in Figure 5. 

General / Ni-Cd / Ni-MH Battery Session C 
Table 3. GOES-10 Battery Discharge History from Launch to Early On-orbit 

* Intermittent discharge currents (up to 0.6 A) from 17:30 to 18:OO. 
GOES- 10 Mag Boom and SA Deploy Sequence 

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Battery 1 & 2 Discharge Currents 

WBTIOISI: Min=0.2055 Max=4.711 Avq=2.881 Sld=1.889 Med-4.050 PIS-8203 MinO1=0.5110 

Figure 5. Battery Discharge Currents During the Magnetometer Boom and 

full Solar Array Deployment Phase for the GOES-10 Spacecraft. 
Battery Reconditioning 

Spacecraft Batteries are reconditioned prior 

to the start of each eclipse season. The 

batteries are individually reconditioned by 

use of the following sequence after vel-ifying 

that the other battery is connected to the 

spacecraft bus. 

a. Turn off battery charging 

b. Open battery discharge relay number 2 

c. Inhibit the battery under voltage 

protection 

d. Turn on battery reconditioning 

The 139.6 ohm resistive load is connected 

across the battery, resulting in an initial Cl48 

(0.25 A) reconditioning discharge rate. The 

individual cell voltages of the selected 

battery are monitored throughout the 

reconditioning discharge period. When the 

first cell voltage reaches 0.5 f 0.1 V, the 

reconditioning discharge is terminated. 

Figure 6 shows the battery reconditioning 

circuitry. 

On-orbit reconditioning has been perfbrmed 

prior to 7 eclipse seasons for GOES-8, and 5 

eclipse seasons for GOES-9. The batteries 

on GOES-10 were not reconditioned prior to 

the fall 1997 eclipse season. 

Charge removed from the batteries during 

reconditioning was calculated using a 

different approach than that described 

earlier. During reconditioning, the nominal 

discharge current has a value of 0.25 A 

(Cl48 rate). However, the step size for the 

discharge current telemetry is 0.06 A, too 

coarse to show discharge current changes as 

the battery voltage changes. Since the 

battery is being discharged by connecting it 

to a constant resistor (139.6 Ohms), the 

discharge current is given by the use of 

Ohm's law 

1997 NASA Aerospace Battery Workshop 

and the charge removed as an integral of 

battery voltage, i.e., 

D (Ah) = (1R) I V dt 

In addition, since the reconditioning process 

continues for 60 - 65 hours, the voltage data 

for integration is sampled at 1-minute 

intervals at the beginning and end of the 

process (where the voltage is changing 

comparatively rapidly) and at 5-minute 

intervals during the middle 48 hour period. 

Figure 7 shows the performance of GOES3 

battery 1 during its first reconditioning cycle 

(Fall 1994). The reconditioning was 

terminated when cell 12 voltage dropped to a 

value of 0.5 V. Corresponding data for 

battery 2 is shown in Figure 8. 

Table 4 compares the results from all 

reconditioning cycles to date: seven for 

GOES-8 and five for GOES-9. The data 

show that the battery capacity has improved 

with time. The table also shows the end of 

discharge (EOD) battery voltage for each 

case. 

General / Ni-Cd / Ni-MH Battery Session 
CURRENT 

SENSOR CURRENT 

TOMAIN 

BUS TO MAIN 

BUS 

- 

- BATTERY 1 1 " - I BA'ITERY~~ 

L li 

-' I - -. 

- - 1 

L.-- - 

4.2 

I 

I 

I 

AECONDITION ON 

RECONDITION ON 

BATTERY 1812 

RECONDITION OFF RECONDlTlONING = 

RESISTOR 888 OHMS EACH 

ti3e.e OHMS EautvI 

Hgure 6. Battery Reconditioning Circuitry 
Charge Removed (Ah) 

Figure 7. GOES-8 Fall 1994 Battery 1 Reconditioning: Battery and Cell Voltages for the Weakest (Cell 12) and Strongest 

(Cell 28) Cells as a Function of Charge Removed from the Battery. 
Charge Removed (Ah) 

Figure 8. GOES-8 Fall 1994 Battery 2 Reconditioning: Battery and Cell Voltages for the Weakest (Cell 14) and Strongest 

(Cell 25) Cells as a Function of Charge Removed from the Battery. 
Table 4. Battery Reconditioning Results for GOES-8 and -9. 

Fall 1994 

Spring 1995 

Fall 1995 

Spring 1996 

Fall 1996 

Spring 1997 

Fall 1997 

Battery 1 

Battery 2 

Battery 1 

Battery 2 

Battery 1 

Battery 2 

Battery 1 

Battery 2 

Battery 1 

Battery 2 

Battery 1 

Battery 2 

Battery 1 

Battery 2 

GOES-8 

Ah 

Removed 

14.14 

14.21 

15.25 

15.41 

15.32 

15.60 

15.89 

15.92 

15.57 

15.73 

16.00 

16.09 

15.68 

15.69 

GOES-9 

EOD 

Voltage 

25.1 

23.9 

19.7 

19.1 

20.3 

19.7 

18.1 

18.3 

19.3 

18.9 

16.3 

17.1 

18.9 

19.5 

Ah 

Removed 

14.06 

13.82 

15.23 

15.14 

15.30 

15.31 

15.77 

15.60 

15.40 

15.26 

EOD 

Voltage 

22.9 

27.5 

18.9 

19.7 

19.3 

20.1 

17.7 

19.5 

18.7 

19.7 
Battery Performance During The 

Eclipse Seasons 

The GOES spacecraft experience the loss of 

solar power during the semi-annual eclipse 

seasons that last approximately 45 days each 

centered on the Vernal and Autumnal 

equinox. The duration of the eclipses vary 

fiom a few minutes to a maximum of 72 

minutes on or near the equinox. We have 

analyzed the eclipse data for all 13 eclipse 

seasons (7 for GOES-8, 5 for GOES-9, and 

one for GOES-10) and found it to be 

consistent and predictable. We provide 

details of the latest (Fall 1997) eclipse season 

below along with some statistical data from 

the previous cases. 

Figure 9 shows the daily battery DOD during 

the Fall 1997 eclipse season for GOES-10. 

The figure shows the DOD for batteries 1 

and 2 and for the battery system as a whole. 

As the eclipse duration gets longer, the 

batteries are discharged for a longer period 

resuting in almost a linear increase in the 

battery DOD. The DOD curve, as shown, is 

not smooth because of the active load 

management that had to be performed to 

stay within the 60 % DOD limit. 

Operationally, we have the so-called 

"20-minute" rule that was imposed after a 

power amplifier failure on GOES-8. The 

uplink carriers fiom the ground station to the 

spacecraft provide extra heating to the 

power amplifiers during the eclipse but result 

in a higher DOD that would exceed the 60% 

limit during the longer eclipses. A balance is 

struck between the two requirements by 

keeping the carriers up during the eclipse 

except during the central portion of the 

season when the carriers are brought down 

at the start of the eclipse and then brought 

up 20 minutes before the end of the eclipse. 

During the central seven days of the eclipse 

1997 NASA Aerospace Battery Workshop 

season, this 20-minute rule was krther 

modified to a 10-minute rule. In addition, the 

Attitude and Orbit Control Subsystem 

(AOCS) team kept both Earth Sensors 

operating during the eclipse season except 

for the days when the power engineer 

required one of them to be turned off to 

avoid excessive drain on the batteries. 

Figure 10 shows the total DOD for the 

battery system for all three spacecraft for the 

Fall 1997 season. The data for GOES-8 and 

-9 also show the effects of the power 

management by modifying the times for the 

uplink camers. However, other aspects of 

active power management are evident only 

for the GOES-1 0 spacecraft since this was 

the first eclipse season for it. 

Figure 11 shows the battery discharge 

currents for Day of Year @OY) 266 

(September 23, 1997) for the three 

spacecraft. The eclipse times are separated 

by 2-hour intervals reflecting the fact that 

the three spacecraft are geostationary at 75 

degrees W (GOES-8), 105 deg W (GOES1 O), and 13 5 deg W (GOES-9) longitude. 

The figures show a sharp rise in the 

discharge currents near the end of the eclipse 

due to the uplink camers (approximately 90 

Watt increase in power consumption). 

Figures 12 through 14 show the battery 

minimum voltage versus DOD for batteries 

1 and 2 for GOES-8, -9, and -10 

respectively. GOES-1 0 data shows larger 

spread between the battery voltages at the 

same DOD (encountered before and after the 

maximum duration eclipse), probably 

reflecting the fact that these batteries had not 

gone through reconditioning, as mentioned 

earlier. Figure 15 shows the battery minimum 

versus DOD data for all 6 batteries together. 

The results clearly indicate that the batteries 

are behaving as a family. 

General / Ni-Cd / Ni-MH Battery Session 

" 
..-.-. Battl ---- Batt 2 - 

263 

Day Of Year 97 

Figure 9. GOES-10 Battery Depth of Discharge (DOD) During the Fall 1997 Eclipse Season. 
Day Of Year 97 

Figure 10. Battery Depth of Discharge (DOD) During the Fall 1997 Eclipse Season for GOES-8, 

GOES-9, and GOES-I 0. 
1997/266/04:00:00- 1997/266/06:00:00 

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1997/266/08:00:00- 1997/266/10:00:00 

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Tue, 4-Nov- 1997 10: 175 1 

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0. - . WBTZDISI 

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< 

2 

E I I 3 

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-WBTIDISI 

(A) 

- - - . WBT2DlSI 

(A) 

08- 

.I..! W .... 07.- I...! 

02s 

.... i

0s- 

-. rm [--I ' . 

Figure 11. Battery Discharge Currents During the eclipse on 

Day 266 (September 23) of the Fall 1997 Eclipse Season 

for (a) GOESS, (b) GOES9, and (c) GOES-10. 

1997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Battery Session 
20 30 40 

DOD (%) 

Figure 12. Battery Minimum Voltage as a Function of Battery Depth of Discharge (DOD) 

During the Fall 1997 Eclipse Season for GOES-8. 

Figure 14. Battery Minimum Voltage as a Function of Battery Depth of Discharge (DOD) 

During the Fall 1997 Eclipse Season for GOES-10. 

36 

35 

Z 

Q 

cn 

- 3 

8 

5 34. E .- 

C 

s 

2 

B 

m 

33 . 

32 i 

0 10 20 30 40 50 60 70 

DOD (%) 

- 

x+ + 

X 

+ 

x 

x+ 

xBatt1 +Batt2 

x+ 

X + 

+ + 

X X X+ +*X+ 

*X+ u +I. 

X W~SH 

X XX 

+ 

X H- + 

u X+ 

X X xw 

+ 
Figure 15. Battery Minimum Voltage as a Function of Battery Depth of Discharge (DOD) 

During the Fall 1997 Eclipse Season for all 6 Batteries on the GOES-8, -9, -10 Spacecraft. 
During the Fall 1997 eclipse season, the 

minimum battery voltages recorded were 

32.5 V, 32.5 V, and 32.9 V (both batteries) 

for GOES 8, 9, and 10 respectively. The 

minimum cell voltages recorded for the three 

spacecraft were: 1.160 V (Battery 1) and 

1.161 V (Battery 2) for GOES-8; 1.161 V 

(both batteries) for GOES-9; and 1.174 V 

(Battery 1) and 1.186 V (Battery 2) for 

GOES-1 0. 

The battery temperatures ranged fiom 

approximately 0 OC to 9 OC for GOES-8, 

1 OC to 12 OC for GOES-9, and 2 OC to 13 OC 

for GOES-10. Daily peaks in the battery 

temperatures were observed near the end of 

the discharge cycle while the minimums were 

recorded a few hours before the end of 

charge cycle, as expected. 

Table 5 lists the minimum battery voltages 

during all of the eclipse seasons encountered 

by the three spacecraft. As expected, the 

minimum battery voltages get lower with the 

aging of the batteries but they seem to have 

leveled off at 32.5 Volts. The battery 

temperatures have stayed in the same range 

for all eclipses (values as indicated above for 

the latest eclipse season). 

Table 5. Minimum Battery Voltages During the Eclipse Seasons 

1997 NASA Aerospace Battery Workshop -1 11- General / Ni-Cd / Ni-MH Battery Session 

Fall 1994 

Spring 1995 

Fall 1995 

Spring 1996 

Fall 1996 

Spring 1997 

Fall 1997 

GOES-8 GOES-9 

Battery 1 

33.3 V 

33.1 V 

32.7 V 

32.5 V 

32.7 V 

32.5 V 

32.5 V 

Battery 1 

33.1 V 

32.9 V 

32.7 V 

32.7 V 

32.5 V 

GOES-10 

Battery 2 

33.3 V 

33.1 V 

32.7 V 

32.5 V 

32.5 V 

32.5 V 

32.5 V 

Battery 2 

33.1 V 

32.7 V 

32.7 V 

32.7 V 

32.5 V 

Battery 1 

32.9 V 

Battery 2 

32.9 V 
SUMMARY AND CONCLUSIONS 

On-orbit performance of the batteries on 

GOES 8, 9, and 10 spacecraft indicates that 

the batteries are performing within 

specifications and results behave as a family. 

In addition, the battery capacities to date 

have shown improvement and indicate a 

leveling off at approximately 16 Ah under 

the Cl48 discharge conditions. Ground test 

data indicate that the batteries will meet 

power requirements for the spacecraft 

mission life of 5-7 years. 

ACKNOWLEDGMENTS 

The authors wish to thank NASA for 

support of this work under contracts 

NAS 5-32590 and NAS 5-32600. They also 

wish to thank the following individuals for 

their help during preparation of this paper: 

Mr. Steve Hall (Naval Surface Warfare 

Center, Crane) for life cycle test data, Mr. 

Ron Hudak (SSL) for activation and 

acceptance test data, Dr. Milt Phenneger 

(CSC), and Mrs. Cheryl Singhal for review 

of the paper. 

REFERENCES 

1. GOES I-M Databook, DRL 101-08, 

GSFC Specification S-480-2 14 

SSL, Palo Alto, CA. 

2. GOES Program, Spacecraft Operations 

Handbook (SOH), Volume I11 (Rev G, 

August 1995), Spacecraft Description, SSL, 

Palo Alto, CA. 

3. GOES I-M Electrical Power Subsystem 

CDR, WDL-TR 11088, NAS5-29500. 

4. Nickel Cadmium Battery Cell Lot 5 

500 Cycle Test Final Report, 

GOES PCC TM PWR 3336 

1997 NASA Aerospace Battery Workshop -1 12- 

5. Harry Brown, Steve Hall, and Gopal Rao, 

Review of Ni-Cd Tests at Crane, presented 

at the 1997 NASA Aerospace Battery 

Workshop. 

ACRONYMS 

Ah 

AMF 

AOCS 

BOL 

DOD 

DOY 

EED 

EOD 

EOL 

GOES 

NASA 

Ni-Cd 

NOAA 

NWS 

PCE 

PCU 

SA 

SSL 

Ampere hour 

Apogee Maneuver Firing 

Attitude and Orbit Control 

Subsystem 

Beginning of Life 

Depth of Discharge 

Day of Year 

Electro-Explosive Device 

End of Discharge 

End of Life 

Geostationary Operational 

Environmental Satellite 

National Aeronautics and 

Space Administration 

Nickel-Cadmium 

National Oceanic and 

Atmospheric Administration 

National Weather Service 

Power Control Electronics 

Power Control Unit 

Solar Array 

Space Systems Loral 

General / Ni-Cd/ Ni-MH Battery Session 
Near Earth 

I 

w 

C-, 

Cf' 

ance 

1997 NASA Aerospace Battery Workshop 

Jason E. Jenkins - JHU Applied Physics Laboratory 

Jason. Jenkins@jhuapl.edu 

Jeff We Hayden - Eagle Picher Industries, Inc. 

JefF. Hayden@kktv.com 

David F. Pickett - Eag e Picher Industries, Inca 

aaaa-energy-enterprises@msn.com 
'r 

'0 

'0 

U 

2 x 1997 NASA Battery Workshop 52 

A % NEAR Flight Battery Performance 4 

R 

m 

g 

d 

T 2 

b 

Presentation Outline 

P 

4 

+ Background & Flight Observation 

- Jason Jenkins - JHUlApplied Physics Laboratory I 

w 

w 

e + Mission Simulation Tests 

- Jeff Hayden - Eagle-Picher Industries, Inc. 

+ Explanation of Events 

9 

8 

k . - David Pickett - Eagle Picher Industries, Inc. 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
Jason. JenkinsQj jhuapl.edu 
4 

'0 

'0 

'4 

s 1997 NASA Battery Workshop 2 

b 

2 NEAR Flight Battery Performance % 

8 

(D 

i? 

d 

q System Description B 

5 ,-. 

'5 

+ System / Mission Design 

.- DET power system architecture 

- Battery isolated from main bus by redundant linear regulated 

chargers and discharge diodes 

- Mission design allows solar-only baseline operation after launch 

+ Battery Charging 

- Ground commanded 2-level constant current charge control (el20 

& CI75) with concurrent &level VT limit (Modified NASA levels 0-7) 

- Over-temperature, overlunder-voltage protection in autonomy 

software 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
1997 NASA Battery Workshop 

Flight Battery Performance 

+ Battery Design 

- 22-Cell, 9-Ah TOMS-EP 

heritage Super NiCdTM 

- Radiative thermal interface 

with thermostatically 

controlled heaters 

- Flight telemetry: stacklhalfstack voltage, 

chargeldischarge current, 

temperature 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
4 

1997 NASA Battery Workshop 

NEAR Flight Battery Performance 

Battery Activity Overview 

P 

G 

+ NEAR launched February 18,1996 

- 15.5% Launch DOD I 

C-l 

C-L 

F - Recharged at Cl20 - VT5 

- Maintained at Cl75 - VT5 trickle 

+ Four Thermal Bias modes in preparation for Mathilde 

9 

8 observation Apr 25 - June 27, 1997 ri 

\ 5 

h - Unexpected I-V performance observed I ? 

% - Ground tests initiated on NEAR life-test cells at EPI 

% 

2 

2 

9 

+ Ground tests confirm no loss of capacity 

a 

k? 

2 

5' 

3 

performance after extended trickle charge storage 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1 997 
- 

'0 

'0 

'4 

5 1997 NASA Battery Workshop 2 

b 

3 

4 . NEAR Flight Battery Performance 

R 

m 

Thermal Bias Regime 

+ Thermal bias intended to top-off battery and raise 

temperature to survive heater disablement during Mathilde 

observation 

- Start from C175 - VT5 trickle charge 

- Raise VT limit to VT7 (highest available per system design) 

- Raise charge rate to Cl20 

- Hold for 4.3 Hours 

- Lower charge rate to C175 

- Lower VT limit to VT5 

+ Thermal bias regime performed 4 times on flight battery 

NASA Battety Workshop, Huntsville,AL Nov 18-20, 1997 
4 

'0 

'0 

U 

2 1 997 NASA Battery Workshop 

2 

k 

a NEAR Flight Battery Performance 

% 

8 

Q 

2 d 

P 

q Flight Thermal Bias Observations E 

b0 

+ First Thermal Bias (Shamtilly 3A) performed in flight April 

25, 1997 (14 months into mission) I 

)-r 

h) 

9 - Starting C175-VT5 trickle charge was VT5 limited 

o Actual trickle charge current -C/100 

- With raised limits, current spiked to .21A quickly decreasing to .15A 

9 with voltage reaching VT7 

f -i 

2 \ - Voltage stayed at VT7 with current gradually increasing to .35A 

$ 

2 over 4.3 h. \ 

2 - C175-VT5 trickle charge following thermal bias did not reach VT5 t 

ttr B limit. 

d 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
'4 

=z hl 1997 NASA Battery Workshop !2 

b 

Q 

3 . NEAR Flight Battery Performance 

% 

8 

B 

a, 

9 

F Flight Thermal Bias Observations r; E (continued) 4 

+ Ground Test Regime commenced May 15,1997 

+ Thermal bias repeated 3 additional times with similarly - I 

w 

c. 

I observed behavior 

- Shamtilly 4 - May 21, 1997 

- Shamtilly 5 - May 30, 1997 

9 

8 - Mathilde Encounter - June 27, 1997 

k 

NASA Battery Workshop, Huntsville,AL N.ov 18-20, 1997 
4 

'0 'n 

'4 

5 1997 NASA Battery Workshop 2 

1 

a NEAR Flight Battery Performance 2 

Q 

'a Ii 

$ 

F 

Thermal Bias 1 (Shamtilly 3A) 

$ April 25, 1997 

% 

14 15 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
1 997 NASA Battery Workshop 

NEAR Flight Battery Performance 

2 

B 

a 

d 

F 

Thermal Bias 1 (Shamtilly 3A) 

B April 25, 1997 (continued) 

4 

11 12 13 14 15 16 17 18 

Time (h) 

11 12 13 14 15 16 17 18 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
d 

'0 

'4 

iz 1997 NASA Battery Workshop t2 

& 

a 

% 

NEAR Flight Battery Performance 

8 

m 

B 

a 

d 

P S 

Thermal Bias 2 (Shamtilly 4) 

F May 21,1997 .no 

31.4 ! I I I I I I I 

12 13 14 15 16 17 18 19 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
1997 NASA Battery Workshop 

NEAR Flight Battery Performance 

a 

VJ 

b B 9 

3 

Thermal Bias 2 (Shamtilly 4) 

S B May 2 1, 1997 (continued) 

4 

0.0 

12 13 14 15 16 17 18 19 

Time (h) 

g' 

NASA Battery Workshop, Huntsville,AL N.ov 18-20, 1997 
4 

'0 

'0 

U 

2 L. 1997 NASA Battery Workshop 52 

k 

C( NEAR Flight Battery Performance 8" 

D 

bn 3 

.;! 

P Thermal Bias 3 (Shamtilly 5) 

% *May 30,1997 

31.4 

10 11 12 13 14 15 16 17 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
Y) 

8 

2 b 

1997 NASA Battery Workshop 

2 

a , NEAR Flight Battery Performance 

3 

1 

B s 

d 

T 

Thermal Bias 3 (Shamtilly 5) 

8 F May 30, 1997 (continued) 

4 

11 12 13 14 15 16 17 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
4 

'0 

'0 

-4 

5 1 997 NASA Battery Workshop 2 

k? y NEAR Flight Battery Performance j 

W 

tu 

a d 

=3 Thermal Bias 4 (Mathilde Encounter) S E June 27, 1997 

4 

7 8 9 10 11 12 13 14 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
1997 NASA Battery Workshop 

NEAR Flight Battery Performance 

0.00 

7 8 9 10 11 12 13 14 

Time (h) 

0.0 

7 8 9 10 11 12 13 14 

Time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
Mission ion Tes 

Jeff W. Hayden - Eagle Picher Industries 

Jeff. Wayden@kktv.com 
d 

3 

s 1 997 NASA Battery Workshop 

2 

b 

ID 

a . NEAR Flight Battery Performance 

2 

OJ 

P 

2 

9 Spacecraft Cruise Simulation 

P 3 z November 30, 1995 through May 15,1997 

h0 

+ Trickle Charge 

- Test cells were selected from the NEAR flight lot 

*:+ 9-Ah 18964EC-9 Super NiCdTM Cell Lot 2 

- Nine cells were assembled into a battery for the trickle 

charge portion of the test 

- The nine cell battery was trickle charged at +5 degrees 

Celsius with 0.140 A 

NASA Battery Workshop, Huntsville,AL N,ov 18-20, 1997 
1997 NASA Battery Workshop 

NEAR Flight Battery Performance 

Cruise Simulation at EPI 

NASA Battery Workshop, Huntsvilie,AL Nov 18-20, 1997 
1997 NASA Battery Workshop 

. NEAR Flight Battery Performance 

Test Battery Reconfiguration 

+ The original nine-cell battery was reconfigured into 2 fourcell batteries, with the gth cell set aside 

- Battery 1 consists of four cells: serial numbers 5, 46, 70, and 73 

- Battery 2 consists of four cells: serial numbers 9, 36, 37 and 50 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1 997 
'r 

3 '4 

5 1997 NASA Battery Workshop 

2 

3 

-i NEAR Flight Battery Performance 

4 

W cp 

B 

2 3 Battery 1 Capacity Measurement 

T 

z May 15,1997 

G 

+ Returned to a trickle charge at .I 2A 

- Temperature, current, and voltage stabilized I 

C-' 

W 

f + Discharged at Cl2 (4.5 A) until the first cell reached 0.5 V 

- Battery temperature at 5 degrees Celsius 

+ Battery 1 delivered 1 1.4 Ah after being on trickle charge for 

9 

D 

is. 17.5 months 

- End-of-discharge voltages were as follows: 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1 997 
4 

3 

V 

2 b 

1997 NASA Battery Workshop 

2 

b 

8 . NEAR Flight Battery Performance 8" 

2 

Battery 1 Capacity Discharge 

Discharge time (h) 

I .4 

521.2 

u 

0 I 0 

a = 0.8 

- - '0.6 

0" 0.4 

I 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1 997 

-Cell 5 1 

- Cell 46 

I 

- Cell 70, 1 
+ 

'0 

'0 

'4 

2 &, 1997 NASA Battery Workshop k? 

b 

3 NEAR Flight Battery Performance 8" 

2 

2 

d 

q Battery 2 Thermal Bias Simulation 

!z 

8 

+ Returned to a trickle charge at .I 2 A 

- Temperature, current, and voltage stabilized I 

C-L 

W 

F + Trickle charged at 0.12 A constant current with a voltage 

* clamp at VT 5 or 5.792V for 2 weeks 

- The voltage clamp was not reached 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1 997 
ki 

'0 

'4 

5 1997 NASA Battery Workshop & 

h 

a . NEAR Flight Battery Performance 2 

1 

bn 

t? 

d 

=3 Battery 2 Thermal Bias Simulation 

8 B (Continued) 4 

+ The VT 5 clamp was raised to VT 7 and the current 

clamped at .45 Amps for 4.3 Hours I 

c1 

W 

7 

- The current initially spiked to about 0.44 Amps 

- The current fell to about 0.37 Amps before it began to slowly 

increase 

Q n - The entire spike event lasted about 13 minutes 

i2 5 

\ - The current increased over a 17 minute period until the 0.45 $ 

2 

\ 

Current clamp was reached 

3 k 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
4 

V, 

V, 

'4 

s 1 997 NASA Battery Workshop 

!i2 

b 

8 NEAR Flight Battery Performance 1 

W 

& 

d 

=3 

Battery 2 Thermal Bias Simulation 

% (Continued) 

- 

+ After 4.3 Hours the VT7 level was reduced to VT5 and the 

current maintained at 0.12 A I 

w 

W 

90 + The thermal bias simulation was repeated 3 additional 

times before the battery was discharged 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
'r 

'Q 

S 

2 b 

1 997 NASA Battery Workshop 

52 

b 

m y . NEAR Flight Battery Performance 4 

R 

m 

&' 

9 

F a 

2nd Thermal Bias Simulation 

z 

4 

NEAR MISSION SIMULATION 

SECOND VT 7 CHARGE 

75 76 77 78 79 80 81 

Time (Hour4 

NEAR MISSION SIMULATION 

SECOND VT 7 CHARGE 

Vdtage At c7 Before Current Reached Llmitl 

I ' I I I I I 

I Tlme (Hours) I 

NASA Battery Workshop, Huntsville,AL N.ov 18-20, 1997 
4 

Y, 

'0 

'4 

% L. 1997 NASA Battery Workshop 5-2 

k 

y NEAR Flight Battery Performance 4 

S 

ba B 

3 

3 0 

X 

4th Thermal Bias Simulation 

NEAR MISSION SIMULATION 

FOURTH VT 7 CHARGE 

0.50 

0.45 

0.40 

0.35 

1:: 

3 0.20 

0.16 

0.10 

0.05 

0.00 

22 23 24 25 26 27 28 

TIME 

NEAR MISSION SIMULATION 

FOURTH VT 7 CHARGE 

8.00 

5.95 

5.90 

E 

t '* 

> 

5.80 

6.70 5522 5 23 24 25 28 27 28 

tlmr (Hours) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
d 

3 

5 1 997 NASA Battery Workshop 

$2 

b 

2 . NEAR Flight Battery Performance 

.a" 

a 

(D 

Battery 2 Capacity Measurement 

June 6,1997 

+ Discharged at Cl2 (4.5 A) until the first cell reached 0.5 V 

- Battery temperature at 5 degrees Celsius 

+ Battery 2 delivered 11.2 Ah after being on trickle charge for 

17.5 months and boost-charged at the C120-VT7 level four 

times 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
1997 NASA Battery Workshop 

NEAR Flight Battery Performance 

Battery 2 Capacity Discharge 

11.2 Ah 

- 

- I 

Cell 9 

- Cell 36 

- Cell 37 

-- 

- Cell 50 

I I I I I 

Discharge time (h) 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
David F. Pickett - Eagle Picher Industries 

aaaa-energy-enterprises@msn.com 
4 

+ Very little, if any, active material is being converted 

+ Cells are at their oxygen gassing potential w 

P 

-P - All current is being used to generate oxygen which is recombined 

by the cadmium electrode. 

- Generation and recombination rate vary according to trickle charge 

9 current. Z k 

\ 

s + Cells are not at same state-of-charge as they would be if 

6 

\ 

3 charged at a higher rate (say Cl20) 

t 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
'-3 

2 

LA 1997 NASA Battery Workshop E 

b 

2 . NEAR Flight Battery Performance $ 

a 

'a 

a d 

F 

% VT7 Level Imposed 

+ Cells now accept more current from charging circuit 

+ A current spike results from nonfaradaic charging 

(capacitive or whatever) 

+ Spike decays after physical charging of electrode surface 

subsides. 

P 

(D 

3. 

\ 

+ Electrodes in cell begin accepting faradaic charge 

$ 

2 - Gassing \ 

2 - Active material conversion 

$ 

'a 

a 

d 

h f $. 

3 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
4 

'0 

'0 

'4 

5 1 997 NASA Battery Workshop 

& 

k y NEAR Flight Battery Performance 4 

a 

m 

bcc 

t? 

d 

3 

VT7 Level Imposed 

i+ (continued) 

+3 

+ Electrodes are accepting charge with positive electrode 

limiting. I 

CL 

$ + Cells most likely will not reach hydrogen gassing potential 

(unless one or more cells out of balance) because of 

voltage clamp 

9 

2 

& 

+ As upper voltage clamp is reached acceptance decreases \ $ 

2 + VT5 level is imposed once the VT7 time interval has timed \ 

3 out at 4.3 hours b 

B f? 

d 

% 

2. 

2 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
t5 

S 

2 b 

1 997 NASA Battery Workshop 

2 

b 

8 . NEAR Flight Battery Performance 

4 

R 

m 

& 

d 

F VT5 Level Resumed ' B ? 

% 

+ Battery now reverts back to trickle charge mode 

+ Post VT7 voltage reduced due to removed polarization I 

c.r 

P 

Y during overcharge. 

2 2 

s 

3 

NASA Battery Workshop, Huntsville,AL Nov 18-20, 1997 
Page intentionally left blank 
BATTERY WORKSHOP 

CRANE CELL TESTING 

SUPPORT OF 

Goddard Space Flight Center 

Steve Hall and Harry Brown 

NAVSURFWARCENDIV Crane 

GopalaEaishna Rao 

NASA Goddard Space Flight Center 

ROSPACE BATTERY CE 

Verification Of Secondary Cells 

Objective and Benefit 

To venfy the quality and reliability of 

aerospace battery cells and batteries for 

NASA flight programs and to design the 

batterykell for upcoming NASA spacecraft. 

1997 NASA Aerospace Battely Workshop General / Ni-Cd / Ni-MH Battery Session 
0042P SAFT/F 40 STRESSB 40 20 

0052T SUPER 50 MISSIONIA 14.4 10 

TRMM 0053T SUPER 50 MISSIONIA 17 0 

TRMM 6151T SUPER 50 MISSIONIA 25 10 

TRMM 6152'3: SUPER 50 MISSIONIA 17 0 

3600H EPI 93 MISSIONIC 11 -5 27 

3601H EPI 93 MISSIONIC 11 -5 27 

I I 

Chart above shows LEO orbit packs cycling at Crane associated with GSFC 

flight programs. There is a total of 11 project packs currently cycling. Nine of 

these are mission simulation regimes and the remaining two are in a 40% 

DoD stress test regime. 

Pack 6016F is a variable DoD cycling regime. The DoD changes every 30 

days with a 7 day trickle charge between changes. This sequence is 5,20,10, 

25,5,10,5,20 and 10 percent depth of discharge and is repeated until the end 

of life. 

Charging techniques vary for each pack and are listed below: 

A. Constant current with a V/T to a CAI ratio then trickle charge. 

B. Constant current with a V/T. 

C. Constant current to specific voltage then trickle charge. 

I997 NASA Aerospace Battery Workshop -150- General / Ni-Cd / Ni-MH Battery Session 
GSFC GEO PROGRAMS 

GOES GOES1 SAFTIG 12 REAL 60 5 

GOES GOESK SAFTIG 12 ACCEL 60 10 

GOES G027B SAFTIG 12 ACCEL 60 5 

TDRSS 6232E SAFTF 40 ACCEL 50 5 

ALL GOES PACKS ARE SEQUENTIALLY CHARGED 

SAFTIG = CELLS FROM GAZNESVILLE 

SAFT/F = CELLS FROM FRANCE 

The GOES packs, are cycled according to a 42 day GEO, with a maximum 

DoD of 60%. Except for GOES 1 the orbits are accelerated time ( six shadows 

per year) at 5 or 10 degree centigrade. All packs are sequential recharge at .9 

amp (six minutes on, six minutes open circuit) with VIT to a specified per-cent 

of ampere-hour out of previous discharge. Then sequentially trickle charge at 

.3 amp (six minutes on, 6 minutes open circuit) with V/T for remainder of the 

twenty-four hour cycle. 

The TDRSS pack 6232E was fabricated by TRW similar to the flight battery 

configuration. This included wrapping each cell in a layer of fiberglass, 

inserting thermal shims between cells and bonding the cells to the shims with 

RTV, and compressing the pack to a mechanical pre-load similar to the flight 

pack (63 psi). The cells are cycled according to a 45 day accelerated GEO 

cycling regime ( 12 hour orbit ) with a maximum DoD 50%. Pack 

temperature is controlled by a cold plate. 

199 7 NASA Aerospace Battely Workshop -151- General / Ni-Cd/ Ni-MH Battery Session 
DISCONTUNUED 

LEO PACKS 

TOMS 6090T SUPER 9 STRESS 50 20 20 

SAMPEX 0090B SUPER 9 MISSION 12 5 18 

SWAS 6015F SUPER 21 STRESS 50 20 13 

HQ 6140s SAFT/F 40 STRESS 40 20 20 

NOAA NOAAl SAFTIG 47 VAR 5 5 

PACK COMMENTS 

6090T LOW EODV .CHARGE CROSSOVER DIVERGENCE 138 MV 

0090B CHARGE CROSSOVER DIVERGENCE 148 MV 

6015F EOC DIVERGENCE 132 MV 

6140s EOC DIVERGENCE 65 MV, LOW EODV, HI CAI, 

C-4 INTERNALLY SHORTED 

NOAAl C-2 INTERNALLY SHORTED 

SAFTIG = CELLS FROM GAINESVILLE 

SAFT/F = CELLS FROM FRANCE 

Pack 6090T at start of cycling repaired leaky pressure fitting connector on cell 

5. Cell 5 removed on cycle 4 150 because of high charge voltage and sent to 

Comsat Labs for DPA. 

Pack 0090B on cycle 41 10 repaired I* pressure fitting connector on cell 5. 

Cell 5 removed on cycle 7171 because of high charge voltage and sent to 

Comsat Labs for DPA. After completing 8000 cycles cell 1 pressure gradually 

increased over a period of 400 cycles. Several reconditioning cycles were 

performed in attempt to control pressure. Recharge trip was then changed 

from 104 to 100 percent on cycle 1 1,454 and pressure returned to nominal. 

Pack 6015F. Charge divergence began after 10000 cycles and gradually 

increased though-out life. 

Pack 6140s These cells were designated as the standard design aerospace 

nickel-cadmium cell and were from of lot 196 . 

NOAAl Cells 1 through 8 are lot 6 cells and were from packs 0648N and 

0649N. Cells failed because of charge divergence aRer completing over 2900 

cycles. Cells 9 through 17 were 42 Ah Mars Observer cells. The purpose of 

the test was to investigate the effects of divergence in a battery pack 

containing more than 10 cells. 

1997 NASA Aerospace Battery Workshop -152- General / Ni-Cd / Ni-MH Battery Session 
DISCONTINUED 

GEO PACKS 

GOES 6227B SAFTIG 12 REAL 60 5 25 

sequential 

GOES 6227C SAFTIG 12 ACCEL 60 0 34 

sequential 

PACK COMMENTS 

6227B COMPLETED MISSION 

6227C COMPLETED MISSION 

SAFTE = CELLS FROM GAINESVILLE 

SAFT/F = CELLS FROM FRANCE 

Pack 6227B was a real time test but after completing the mission requirements 

of 13 shadow periods it was changed to accelerated orbit similar to pack 

6227C. 

Packs were cycled according to a 42 day GEO regime with a maximum of 

60% DoD. The orbit was accelerated time at 0 or 5 degree centigrade. 

They were sequential recharge at -9 amp (six minutes charge, six minutes open 

circuit) with V/T of 1.482 vlc to 1 15% or 1.508 vlc to 120% of ampere-hour 

out of previous discharge. Then sequentially trickle charge at .3 amp (six 

minutes charge, 6 minutes open circuit) with V/T for remainder of the twentyfour hour cycle. 

Packs were discontinued after completing mission requirements and special 

mission unique test+ 

1997 NASA Aerospace Battery Workshop -1 53- General / Ni-Cd / Ni-MH Battery Session 
HQ PROGRAMS 

6091T SUPER 9 MISSION 31 5 

6106M MAGNUM 10 

6122M MAGNUM 10 STRESS 40 20 

0106M MAGNUM 21 

0121M MAGNUM 21 STRESS 40 20 

6506M MAGNUM 50 

6522M MAGNUM 50 STRESS 40 20 

5009M HUGHES 48 STRESS 60 10 

Packs with DoD identified as variable (VAR) are being cycled as follows: 

The first 2000 cycles DoD was 10 per-cent. The next 2000 cycles DoD was 

increased to 25 per-cent. Then DoD increased to 50 per-cent for 2000 cycles. 

This sequence is repeated until the end of life. 

1997 NASA Aerospace Battery Workshop -154- General / Ni-Cd / Ni-MH Battery Session 
MAGNUM PACKS 

Due to the anomalous capacity and voltage performance of Super and 

Magnum cells have experienced at low temperatures (-10 to 5" C). A program 

was initiated to test a sampling of these cells at either a particular mission 

profile or at low temperature with variable DoD7s. 

The C/D ratio given in chart is the normal operating range for the cycle life of 

pack. 

The 10 and 50 Ah cells are cycled with the same parameters the only 

difference being the voltage limits for 10 Ah pack 6106M is 20 millivolts 

higher at all levels. 

Recharges were higher for the second sequence of 10% DoD 

Packs 0 106M and 0 12 1M ( 21 Ah cells) recharges are higher than other packs 

for the same voltage limit. This is true for both test regimes. 

1997 NASA Aerospace Battery Workshop -155- General / Ni-Cd / Ni-MH Battery Session 
PACK 6506M(50 AMPEREeOUIp 0.C VARIABLE DoD 50.10,25% 

CROSSOVER DlYERCENCE c 4 MIUNOLTS 

---- 

CYCUS 

I 

Plots shows each of the six cells of the pack during a MI charge and discharge 

cycle at 50, 10 and 25 depth of discharge . There is very little charge 

divergence( 4 millivolts) in pack. 

Crossover occurs when the charging regime is changing from constant current 

to a constant potential. 

The voltage limit on the 50 Ah pack 6506M was decreased by 20 millivolt 

(1.420 volts) after 900 cycles into the second sequence of the 10% DoD test. 

This was done because of recharges above 119%. This resulted in no increase 

in the end of charge divergence or significantly decreasing the end of 

discharge voltages. 

1997 NASA Aerospace Battery Workshop General / Ni-Cd/ Ni-MH Battery Session 
PACK6106M (10 AMPEREEOUR) O'C VAFtlABLE DoD SO. 10.25% 

CROSSOVER DIYERGMCE w 25 WOLTS 

Plots shows each of the six cells of the pack during a full charge and discharge 

cycle at 50, 10 and 25 depth of discharge. Charge divergence varies from 26 

to 32 millivolts. 

Crossover occurs when the charging regime is changing from constant current 

to a constant potential. 

The voltage limit for the 10 Ah pack 6106M was increased by 20 millivolts 

(1.500 volts) for approximately 1700 cycles . This occurred around 4600 

cycles and during the 50% DoD test. This increase in voltage did not improve 

charge voltage divergence or significantly improve end of discharge volts. 

The 5O%DoD was returned to the 1.480 volt level after this sequence. 

1997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Battery Session 
COMNENTS ON 

PERFORMANCE 

Divergence is greater at crossover 

Pack 6 106M (1 0 Ah) divergence six times 

higher than 6506M (50 Ah) 

6106M (10 Ah) voltage limit are 20 mv 

higher all DoD's 

Stress packs 6122M (10 Ah) and 6522M 

(50 Ah) divergence < than 5 mv 

Crossover occurs when the charging regime is changing fkom constant current 

to a constant potential. 

The 10 Ah pack 6 106M divergences is six times greater than the 50 Ah pack 

6506M. All testing parameters are the same except 6 106M voltage level is set 

20 millivolts higher than 6506M. 

Their sister packs in the 20 "C 40% stress regime testing parameters are 

identical. However their C/D7s and end of discharge voltages are very similar. 

1997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Batteiy Session 
COMMENTS ON C/D 

Recharges were 4-8% higher for the 

second sequence of 10% DoD 

The 10Ahpacks0106Mand0121MC/D's 

higher for same voltage value 

Recharges were 4 to 8 percent higher on the second sequence of 10 % DoD. 

The 10 ampere-hour packs CID's were 5-10 percent higher for the same 

voltage level when compared to the 21 and 50 ampere-hour packs. 

1997 NASA Aerospace Battery Workshop General / Ni-Cd / Ni-MH Battery Session 
HUBBLE CAPACITY CHECK 

Yearly capacity checks were performed 

only on 360 1 H for the first 3 years 

Pack Shunted With 12 Ohms To .2 Volts Any Cell 

After 46 months capacity checks were 

initiated for both packs. 

Packs were discharged at C/6 to .9 volts any cell prior to shunt period. 

Packs 3600H capacity increasing by 2.7% 

and 3601H by 1.1 and 4.3% each period. 

Flight batteries experiencing the same capacity increase. 

The ten cells used in these two packs were manufactured by Eagle-Picher 

Industries, Inc. (EP) for the Hubble Space Telescope @ST) project and are 

from the Flight Module 1 and Flight Module 2 lots. The cells from pack 

3600H are from lot FM1. The cells from pack 3601H are from lot F'M2 which 

is the same lot as the cells that have been in orbit on the HST spacecraft since 

April 24,1990. 

Both packs followed the same set of parameters except pack 3601H which 

was reconditioned after each yearly shadow period. A resistor of 1.2 ohm 

was placed across pack until any cell reached -2 volts. 

This procedure was changed after 46 months. Both packs are now 

reconditioned has described below: 

Charge 9.3 amp for 10 hours 

Charge 4 amp for 14 hours 

Discharge 15 amp to -9 volts first cell 

Shunt pack with 1.2 ohms to -2 volts first cell 

Sequence performed a total of five cycles. 

The capacity growth during capacity checks is similar to increase's experience 

by the flight battery. 

1997 NASA Aerospace Battery Workshop -160- General / Ni-Cd / Ni-MH Battery Session 
Y. Sone, H. Kusawake, K. Koga, and S. Kuwajima 

Office of Research and Development 

Electronic and Information Technology Laboratory 

National Space Development Agency of Japan 

(NASDA) 
ADEOS ADEOS-2 

NfisDA 

ALOS 

f **** * 

Ni-Cd 1 ETS-6 ETS-'7 

Development of Space Batteries at NASDA 

COMETS DRTS ETS-8 @ (1 ton class) IPV 

(mission) CPV 

MDS-l ( small satellite- 1 ton class) (mission) 

f \ 

Ni-MH w 

< J OICETS 

wl [L~+/c secondary f 

HOPE-X 

a platform, a lunar base, [~e~enerative FC] 

t J a lunar astronomical observatory 
@ NnSC>n 1 Organization for the NASDA Battery Develoornent 

I Office of Research and ~eveloDment I 

I- I Research Planning Division I 

1-1 Advanced Mission Research Center I 

t- I Small Satellite Systems Laboratory 1 

I ( Lunar Mission Laboratory 

Guidance and Propulsion Technology Laboratory 

I 1 Environmental & Structural Test center1 
@ NfisDA I Electronic and Information Technology Laboratory 

Major study activities 

-LIDAR Technology Demonstration Satellite 

(LIDAR : Light Detection and Ranging) 

-Solar Cell 

-Communication and Data Handling 

-Optical Inter-Satellite Communication 

-Remote Sensing Technology 
Development of Ni-MH Cells 

NASDA has been developing Space Ni-MH Cells on the contract 

with SANYO Electric Co. Ltd. 

1st manufactured cell showed initial capacity of over 35 Ah and 

specific energy density of about 50 Wh/kg (about 860 g of weight). 

Ni-MH cell was decided to be used for Optical Inter-orbit 

Communication Engineering Test Satellite (denoted as OICETS) 

which will be launched in 2,000. 

In this report, we are going to show the stable chargeldischarge 

performance of Ni-MH cells, and its applicability to OICETS. 
P\lnSnfi 1 Comparison of the Energy Density of SpaceBattery Cells I 

Energy Dens i ly (W h / k g ) 

ll : N i .... &Ill 

.) : Ni-Ccl 

A : Ni-Cdcformer qualified products) 

(tiof ~~rc~clucc!cl toclay) JERS- 1 

'25Ah C AUSSAT 0 

0 

BS-3. ANIK NOAA 

.a INTELSAT-IV 0 

BS-2X 0 

0 0 0 LEASAT 

SBS 

0 0 GALAXY-5 (DOD = 80%) 

0 GALAXY-6 (DOD = 80%) 

IN I NTELSAT- N 0 SMM <NASA/MMS> 

A ETS-V (QTS) 

A MOS-1 (QTS) 

O 35A11 

0 

I NTDLSAT-V. GPS 

CS-3. STATRSAT 

DSCS. SKYNET, DFS 

0 

TRMM 

(DOD = 40%) 

501111 @ 

LANDSAT 

UAllS 

GRO 

<NASA/MMS> 

NATO-3 and CS battery cells were prepared by Eagle Picher ' 

GALAXY-5.-6.TRMM battery cells were prepared by Hughes. Supe,r Ni-Cd 

Othcr bal tcry ccl Is wcrc prepared by Gates Acrospacc Bat lcry (for~trcr GB> 

..I L..- 1 I.-. 1 . . 

2 0 30 4 0 50 60 

Capacity / Ah 
@) NI\lnDA ( Requirement for the 35 Ah Ni-MH cell 

0.5 C, 30 min 0.8 C, 30 min 

Test We 1 2 

Requwment woOcyCles l0,oOO cycles 
Life Cycle Test with DOD = 25 % 

.6i EOCV 

.4-I 

0 10,000 20,000 0 10,000 20,000 

Cycle Number Cycle Number 

BBM in FY 1992 BBM in FY 1993 

The amount of electrolytic solution was increased, 

while the space between electrodes was decreased. 
EOCP 

@#$I 0"\9AsDA 

I 1. ' P =f m 

EOCP 

1. EODP PI 

Life Cycle Test with DOD = 40 % 

Cycle Number Cycle Number 

BBM in FY 1992 BBM in FY 1993 

The amount of electrolytic solution was increased, 

while the space between electrodes was decreased. 
@;NASl>&a , %[..'$ 1 

.fP3,.f The Capacity of 35 Ah Ni-MH Cells (BUM in FY 1993) 

The capacity of the calls designctl in FYI993 shows very stable performance. 

It revealed to have more than 10,000 life cycles with DOD=40%. 

1,000 10,000 

Cycle Number 

1,000 10,000 

Cycle Number 
@ NnsDA 1 13 Ah Ni-MH Cell for OICETS 

After the life cycle tests of 35 Ah cells designed in 

FY1993,13 Ah Ni-MH cell for OICETS, which is 

scheduled to be launched in 2,000, was designed and 

qualified for the space use. 

I Nominal Caoacitv 1 1 3Ah 

Required Life Cycles 20,000@DOD=25% 

10,00O@DOD=40% 

I Weight 1 less than 390 g 

I Energy Density I more than 40 Wh/kg 
OlCETS is a denotation of 'Optical 

Inter-orbit Communication Engineering 

Test Satellite.' This satellite is going to 

make an optical contact with ARTEMIS 

from ESA. 

The successful development of Ni-MH 

cells enhanced us to use it for this 

satellite. 

p&v~ 

*2$', 

'Optical Inter-orbit Communication Engineering ~est Satellite ' 
The Performance of 13 Ah Ni-MH EM Cells 

The trend data of 13 Ah Ni-MH cells (EM) are shown during the 

cycle tests with DOD = 40%. We could observe very stable 

performance of these type of cells. Now, the life cycle tests of QT 

cells are also being measured. The same type of stable performance 

of QT cells as that of EM ones is observed. 

EOCV 

EODV 

EOCP 

1 .o'- EODP i 

Cycle Number Cycle Number 
@) F\lnsm I The relationship between CID Ratio and Capacity 

The following figures shows the relationship between C/D ratio 

and capacity of 13 Ah Ni-MH cells. These results revealed the 

best C/D ratio for N-MH is 1.05 when the DOD. 

Charge Voltage / V Charge Voltage / V Charge Voltage / V 
NNAsfi V-T curve for OICETS 

Using tile reiatio~lsllip 

between C/D ratio ancl 

capacity, V-T curves for 

OICETS were prepared. 

The line No. 5 will be used in 

the case of DOD=25%. 

The line No. 7 will be used in 

the case of DOD=40%. 

Tl~e lines of No. 1-4 wot~ld be 

used if 'short-circuit' 

llappened to one of tile cells. 
1 The renewal designed 35 Ah Ni-MH Cells 

Aiming at the prolonged life and lightweight cell, we redesigned the 35 Ni-MH cell. 

The main defference from the standard cell is as follows. 

Design 

NEG/POS ratio 

Separator 

Electrolyte / g 

standard 

1.95 

Ny lone 

92.2 

prolonged life 

1.95 

Polypropylene 

106.03 

lightweight 

1.61 

Polypropylene 

106.03 (TBD) 
3 

.a" The Ni-MH cells were manufactured and the life cvcle 

b 

? 

J 

performance were measured. 

# 

3 @I NASDA 

Through the measurement using 35 Ah Ni-MH cells, we 

confirmed the stable performance of this type of cells with the life 

cycle more than 20,000 with the DOD=25%. 

Summary 

These stable performance enhanced us to prepare 13 Ah cells for 

OICETS. 

2 8' 

Now, the life cycle tests for 13 Ah cells are continued. The trend 

data suggests the satisfactory performance of 13 Ah cells. 
Page intentionally left blank 
Battery Systems . 

for 

X-38 Crew Return Vehicle (CRV) 

and 

Deorbit Propulsion Stage (DPS) 

NASA Aerospace Battery Workshop 1997 
4 

NASA-Johnson Space Center, Houston, TX '0 

3 

Outline 

4 , 

8 

tn 

i2 Introduction 3 

a 

T 

P Obj ectives/Approach 

- Common 

- 28V 

- 270V 

Design Trades/Solutions/Redundancy PlanIMargins 

- 28VDPS 

- 28VCRV 

- 270V CRV 

Interfaces 

Deviations from OPS CRV 

12/1/97 Eric Darcy/28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

, Top Level CRV Requirements 

Provide for the safe return of ISS crew of zero to 6 in case 

- emergency return of ill or injured crew person 

- ISS can not maintain critical systems, pressure, attitude, or is contaminated 

- Shuttle is not available to return crew 

Crew return without pressure suits 

* On-orbit lifetime of 3 years 

700 nautical miles of cross range 

Land lander 

Separation time from ISS < 3 minutes 

Planned return mission time is 3 hours maximum 

Contigency return mission time is 9 hour maximum 

1 21 1 197 Eric Darcy1281-483-9055 
NASA-Johnson Space Center, Houston, TX 

12/1/97 Eric DarcyI28 1-483-9055 
Ir 

NASA-Johnson Space Center, Houston, TX 

2 

u 

i? 2 

2 I CRV Advantages over Soyuz 

2 

& 

9 

? 

8 k. Returns a crew of six with wider size tolerances 

Has 700 nmi of cross range 

- Lifting body aerodynamics 

- Electromechanical Actuators for flight surface control 

- 2 flaps, 2 rudders 

Soft, precise land landing at many more sites 

- Gliding parafoil 

- Electric motors power winches to control the chute 

* 3 year on-orbit life 

* Potentially rehrbishable 

Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Unique Approach to X-38 Program 

Design, Build, and Test in small increments for rapid feedback 

- Pallet Drops (parachute weight tests) 

- Dog House Drop (parachute drop test with a vehicle-like shape) 

- V- 13 1 (X-24 aero shape with fixed surfaces dropped from a B52) 

- V- 132 (same shape with EMA controlled surfaces) 

- V- 133 (20% bigger, again B52 dropped) 

- V-20 1 (Shuttle launched, 5/00, unmanned return test) 

- V-202 (Ariane launched, 3/02, unmanned return test) 

No prime contractor (except for Deorbit Propulsion Stage) thru V-202 

Later, a prime contractor will build operational CRVs for ISS 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Battery Objectives and Approach 

Provide safe battery designs for lowest volume and cost, and within 

schedule 

Take advantage of less complex reqts for V201 vs OPS CRV to 

simplify design and reduce cost 

Use only existing commercial cell designs as building blocks for larger 

battery 

Derive battery designs from the ASTRO-SPAS design which is the 

largest lithium battery design with Shuttle flight experience 

Place maximum amount of battery energy on DPS 

DPS battery is non rechargeable 

* CRV batteries are rechargeable 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

12/1/97 Eric Darcyl28 1-483-905 5 
4 

NASA-Johnson Space Center, Houston, TX 

'4 

2 b 

k? 

3 

a 

% , Requirements and Groundrules 

8 

bn 

9. 

3 

F 

% B Common to all batteries 

G 

- 1 failure tolerant for mission success and to a critical hazard 

- 2 failure tolerant to a catastrophic hazard 

- Compliant to "Manned Space Vehicle Battery Safety Handbook" 

JSC-20793 

- 9 year useful life (5 for storage, 1 in transit, 3 on-orbit) 

- Capacity gauge circuit to track %Ah remaining with 1 % resolution 

- Gauge draws < 50 microA (<2 Ah over 4 years) from the battery 

- Refurbishable by replacing battery strings 

- Capable after vibration to 14 grms 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Requirements and Groundrules (cont.) 

28 V High Energy Density Battery (WhIL) 

- 28.9 kwh over a continuous 9 hour period for Crew Return Vehicle 

Avionics 2.2 kW for 9 hours 

ECLSS 0.9 kW for 9 hours 

RCS 1.0 kW for 1 hour 

- 7.0 kwh over a continuous 7 hour period for Propulsion Module 

270 V High Power Density Battery (WIL) 

- 4 kwh over a 25 minute period for EMAs, then Winches 

EMAs peak at 100 kW (400A) for 80ms every 2s for 15 minutes 

Baseline load is 8.91 kW (33A) for 15 minutes 

Winch peak is 50 kW (200A) for 5s at end of 10 minutes 

Baseline load is 4.05 kW (1 5A) for 10 minutes 

1 21 1 197 Eric DarcyI28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

DPS Battery Requirements 

Performance 

- 29.4 kwh at 28 +5V7 -4V at 7 hour rate after 

- 3 years on-orbit while at 0 to 30C (14 days for V201) 

- Minimum capacity = 1050 Ah 

- Average Power = 4.2 kW 

- Peak Power = 5.5 kW for <1 minute 

- Vacuum exposure for 3 years (14 days for V201) 

- Non-operating exposure range -3 1C to 72C 

- Use high energy density commercial lithium primary cell design 

- Compliant with NASA RP 1353 "Primary Battery Design and Safety 

Guidelines Handbook" 

Operation 

- Expired batteries replaced on ground, if CRV/PM returned via Shuttle 

- A rechargeable 28V battery on CRV is used for monthly check-outs 

12/1/97 Eric DarcyI28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Li Cell Charaterization Approach 

Released an SOW to determine feasibility of candidate cell designs 

Schedule (finish by Dec 97) and cost (<$25K) restraints 

Response from vendors; 

- Friwo-Silberkrafl; test both SO, and MnO, for $24.9K 

- WGL: only willing to sell their BCX DD-cells, Crane did testing, $24.6K 

- YTP: test SOCl, DD-cell in two phases $22.4K and $22.3K 

- Eagle Picher: wanted too much ($62K) with using existing cells 

- Centaur: no formal bid yet 

- Ultralife: can not support schedule 

- Bluestar, PCI, and BEI: no bid or no response 

12/1/97 Eric Darcy/28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

SOW for Li Cell Characterization for Manned Spacecraft 

Purpose - Determine performance and safety of a candidate building block cell 

Design - Consider only existing cell designs with very minor modifications 

Manufacture - Build enough (26 minimum) cells to support verification tests 

Performance tests 

- Cell capacity at 3 rates and 2 temperatures 

- Cell Vibration of BOL and EOL cells (14.1 grms at 0.16 g2/Hz) 

- Acceptance tests (leakage, weight, dimensions, OCV, CCV, X-ray) 

Safety tests 

- Cell voltage reversal at 3 rates and 2 temperatures 

- Cell short circuit at 2 rates and max temperature 

- Cell charging with and without diode protection 

- High temperature exposure and heat-to-vent 

Cell Pressure Analysis to determine pressure vs temperature at all states of charge 

Deliverables - Test report with complete cell drawings to establish configuration control 

12/1/97 Eric Darcy/28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

' Lithium Cell Design Trades 

LiIMnO, cell has some advantages 

Company 

Friwo-Silberkraft 

Friwo-Silberkraft 

Wilson Greatbatch 

Yardney 

less hazardous than BCX or SOCI, options 

NCR may not be required for internal short hazard 

only chemistry w/o voltage delay issue 

Chemistry 

SO2 

Mn02 

BCX II 

SOC12 

acceptable volume 

Cell 

G62 

M25 

DD 

IRDD-3H 

Eric Darcy/28 1-483-9055 

Remarks 

Proven battery design, highest volume 

Same batt design, Li-MnO2 cell 

Vendor will only do cell work 

High temp DD cell, experienced vendor 

Mass, Ibs 

643.10 

629.80 

435.09 

525.20 

Volume, cu.in. 

16,618.15 

11,870.08 

8,869.92 

10,050.67 
NASA-Johnson Space Center, Houston, TX 

Design Solutions/Redundancy Plan/Margins 

28V DPS Battery Point Design 

- Using reference Cell: LiIMnO, 32Ah cell (PIN M25) from FriwoSilberkraR, Germany 

- Battery String = 12 cells in series 

- Battery Module = 12 strings in parallel 

- Flight Battery Set = 4 Battery Modules 

- Voltage: Open Circuit = 38V, Closed Circuit = 24-33V 

- Capacity starting at OC = 1536 Ah vs 1050Ah required 

- Redundancy Plan: 3 modules needed, 4 flown 

- Margin after one module fails: 9.7% 

- Capacity gauge circuit in each module 

- Battery mounting plate for each module 

12/1/97 Eric Darcyl28 1-483-9055 

NASA-Johnson Space Center, Houston, TX 

Total weight = 79 kg 

Mounting plate t=40 mm 

I I I I Plate Weight = 7 kg 

DPS will require 4 battery modules like below 

f------------- 620 mm 

1 21 1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Interfaces 

Structural 

- BMP to be designed by DPS contractor 

- Built by battery vendor 

Electrical 

- Power - Pair of 8 AWG stud post connectors for +, - 

- Data from Capacity Gauge (8 wires, RS-422) 

CHG: need +5V for wake-up from avionics, switched off during sleep mode 

TXD: output battery data to avionics every 1.7s 

CTS, RXD: inputs for GSE only to calibrate, modify, inspect parameters 

data output: 16 byte packet (%Ah remaining, voltage, current, temperature) 

Serial output data format: 9600 baud, 8-bit, no-parity, lstop bit 

Each of the 4 lines has a dedicated return line : 

Thermal 

- Starting battery temp: 0 to 30C 

- Non operating exposure: -3 1 C to 72C 

- Generates an average 840W of heat during discharge 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

CRV 28V Battery Requirements 

Performance 

- 7.2 kwh at 28 + 5V,- 4V 

- Average power = 3.6 kW 

- Peak power = 7.2 kW for seconds? 

- Minimum Capacity = 258Ah at 2 hr rate after 36 cycles 

- Self-discharge <1% per day 

- Stowed in cabin, capable of vacuum exposure non-operating 

- Temperature range at start of operation = 10C to 45C 

- Non-operating temperature = -40C to 55C 

On-Orbit Operation 

- 1 hour discharge every month over 3 years (over 1 year for V201) 

- Recharge in 4 hrs from 4.2 kW max at 120V or 28V (28V only for V201) 

- Unavailability time < 5 hourdmonth including 1 hour check-out (99.3%) 

- Top-off continuously fiom 40 W max at 120V or 28V (28V only for V201) 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 4 

'0 

'0 

'4 

z 

2 

k 

a 

G a

NiMH Cell Design Trades m 

2 $ 

- 28V MV Battery 258 Ah # of cells 

-4 P Cell Design Capacity, Ah mass, lbs volume, cu.in. 

$ 

4 

Sanyo NiMH HS-35s 37.00 184.00 493.05 78,217.30 

I + 

Sanyo NiMH HS-50s 53.00 138.00 531.30 86,940.00 

F 

Sanyo NiMH HR-413FAU 3.50 2,208.00 41 1.71 75,288.38 

Toshiba NiMH TH-3500 3.40 2,208.00 355.40 67,154.1 I 

Smaller NiMH cells are used in point design because ,, 

* more WhIL and less memory effect than any NiCd 

more quickly available and less costly than large prismatic NiMH cells 

proven on EMU Helmet Lights on STS-82 and STS-86 

1 21 1 197 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

d 

2 

2 t~ 

NiMH Cell Charaterization Approach 

f? d Awarded a SOW to determine abuse tolerance of best 2 candidate cells 

- Toshiba TH-3500 3.398 Ah at C12 over 100 cycles 

- Sanyo HR-413FAU 3.503 Ah at C12 over 100 cycles 

Above cell designs have demonstrated highest WhIL in our testing 

Schedule (finish by Jan 97) and cost <$15K 

0, 

E g. 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

NiMH Cell Abuse Tolerance Testing On-going 

Purpose - Determine safety of a candidate building block cell 

Design - Consider only existing cell designs with very minor modifications 

Manufacture - Build enough (43 minimum) cells to support verification tests 

Performance tests 

- Cell capacity at C/2 with 2C pulses at RT 

- Cell Vibration 

- Acceptance tests (leakage, weight, dimensions, OCV, CCV) 

Safety tests 

- 100% Cell voltage reversal on 3 cells 

- Cell short circuit with 23 cells in series 

- Cell overcharging tolerance 

- High temperature exposure and heat-to-vent 

12/1/97 Eric Darcy128 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Design Solutions/Redundancy PlanIMargins 

28V CRV Battery Point Design 

- Reference Cell: NiMH 3.4Ah cell (PIN TH-3500) from Toshiba, 

Japan 

- Battery String = 23 cells in series 

- Battery Module = 12 strings in parallel 

- Flight Battery Set = 8 Battery Modules 

- Voltage: Open Circuit = 3 1 V, Closed Circuit = 24-3 1V 

- Capacity starting at 10C = 3 16 Ah vs 258Ah required 

- Redundancy Plan: 7 modules needed, 8 flown 

- Margin after one module fails: 7.4% 

- Capacity gauge circuit in each module 

- Charger circuit in each module accepting 28V input 

- Battery mounting plate for each module 

12/1/97 Eric Darcy/28 1-483-9055 

NASA-Johnson Space Center, Houston, TX 

Total weight = 26 kg 

I I 

Plate Weight = 6 kg 

fi 

CRV will require 8 battery modules like below . 

................................ ........................................................................... .................................................................................................................. 

..................................................................................................................... 

....................................................................................................................... ... .................................................................................................... > .................................................................................................................... '.':y::.:~+b~f;:pg&4g:!s$;B $~~v~:;~~d~l.$;::~~;

..... 

............ 

... 

. .................................................................................... ............................................................................................................... , :. ........ .......................... ......... .............................................................................. ........................................................................................................... 

...................................................................................................................... ................ ................................................................................................................... .......................... ................... ....................................... .......................................... 

505 mm--7 

12/1/97 Eric Darcy128 1-483-9055 

Mounting plate t=40 mm 

I I 
NASA-Johnson Space Center, Houston, TX 

Interfaces 

Structural 

- BMP to be designed by JSC 

- Built by battery vendor 

Electrical 

- Power - Pair of 12 AWG stud post connectors for +, - 

- Data from Capacity Gauge (8 wires, RS-422) 

CHG: need +5V for wake-up fiom avionics, switched off during sleep mode 

TXD: output battery data to avionics every 1.7s 

CTS, RXD: inputs for GSE only to calibrate, modify, inspect parameters 

data output: 16 byte packet (%Ah remaining, voltage, current, temperature) 

Serial output data format: 9600 baud, 8-bit, no-parity, 1 stop bit 

Each of the 4 lines has a dedicated return line 

Thermal 

- Starting battery temp: 10 to 40C 

- Non operating exposure: -3 1C to 55C 

- Average heat generation during discharge = 180W, during charge = 720W 

12/1/97 Eric Darcyl28 1-483-9055 
4 

NASA-Johnson Space Center, Houston, TX 

V) 

'0 

'4 

CRV 270V Battery Requirements 

k 

Y 

4 a Performance 

- 270 +60/-20Vduring discharge, 360-367V during charge 

- Average power = 1.65 1 kW, Peak Power = 100 kW during 80 ms 

- Minimum Capacity = 14.68 Ah after 

- 36 five minute discharge cycles once a month 

- Self-Dischage < 0.2% per day 

- Outside cabin, vacuum exposed for 3 years (14 days for V201) 

- Commercial lead acid cells 

On-orbit Operations 

- 5 minute discharge every month over 3 years (over 1 year for V201) 

- Recharge in 4 hrs from 4 kW max at 120V (GSE charging only for V201) 

- Unavailability time < 5 hourdmonth, including check-out (= 99.3%) \ 

5 - Top-off continuously from 20 W max at 120V (GSE charging only for V20,l) Q '. 

3 

- GSE charging on pad no earlier than T-10 days 

t 

1 21 1 197 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

High Power Cell Design Trades 

1 C Total # Battery Battery 

Cell Desinn Cap. (A-h) mass (#) volume (cu.in.1 

Hawker Cyclon D 1.90 1920 1052.6 14744.5 

Hawker Cyclon Tall D 3.70 1280 1073.1 14852.7 

Hawker Cyclon X 3.90 1280 1427.1 19800.7 

EPS NiCd Cs EPP-1750CS 1.65 3690 613.7 7296.7 

EPS Ni-MH Cs EMP-2200CS 2.00 3690 602.4 7335.0 

Sanyo NiCd D N-4000DRL 4.00 2214 1022.9 14719.1 

Sanyo NiCd Cs N-1700SCR 1.70 3936 654.6 8581.7 

Bolder 915 Cs 1.20 2880 727.4 10147.9 

Hawker Cyclon Tall D cells used in point design because 

Vacuum exposure tolerance of Cyclon cell has been demonstrated by test 

Crimp seals of NiCd and NiMH cells are untested at vacuum 

NiCd, NiMH require too many cells in series 

12/1/97 Eric Darcy128 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Cyclon Tall D-cell Performance and Abuse Tests 

Purpose - Determine performance and safety of a candidate building block cell 

Performance tests 

- Capacity cycling - 36 cycles with CRV peak profile while at 10,25, and 40C 

- 12V battery capacity cycling 

Discharge - crude CRV profile 

Charge - CV with 1A current limit 

- Cell Vibration 

- Acceptance tests (leakage, weight, dimensions, OCV, CCV) 

2 months exposure to 80C wl and wlo vacuum 

< 0.2% mass loss in both cases 

Safety tests 

- 100% voltage reversal on 12 V battery - benign vent 

- Short circuit with 15 cells in series at 90A - no vent 

- 12V battery overcharging tolerance - benign vent 

- High temperature exposure and heat-to-vent - benign 

12/1/97 Eric DarcyI28 1-483-9055 
r, 

NASA-Johnson Space Center, Houston, TX 

U 

2 b 

2 Design Solutions/Redundancy PlanIMargins k 

Z1 $ 

2 270V CRV Battery Point Design - 

8' 

3 - Reference Cell: Lead acid 2.5 Ah cell (PIN Cyclon Tall D) from 

Hawker Energy Products, MO 

- Battery String = 32 cells in series 

- Battery Module = 5 strings in series 

- Flight Battery Set = 8 Battery Modules 

- Voltage: Open Circuit = 342V, Closed Circuit = 250-320V 

- Capacity starting at OC = 20 Ah vs 14.68Ah required 

- Redundancy Plan: 7 modules needed, 8 flown 

- Margin after one module fails: 19.2% 

- Capacity gauge circuit in each module 

- Charger circuit in each module accepting 120V input 

bn f? 

d - One battery mounting plate for two modules side by side 

12/1/97 Eric Darcyl28 1-483-9055 
4 

NASA-Johnson Space Center, Houston, TX 

s 

U 

5 

k2 

Envelope/Size/Mass 

b 

3 j 

S 

2 battery module size wl Battery Mounting Plate 

i? d - 680 mm wide 750 mm wide 

T B - 760 mm deep 830 mm deep E 

.i: - 122 mm tall 172 mm tall 

- 122 kg 135 kg 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

I I 

Plate Weight = 13 kg 

Total weight = 135 kg 

CRV will require 8 battery modules like below 

- 750 mmp 

12/1/97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Interfaces 

Structural 

- BMP to be designed by JSC 

- Built by battery vendor 

Electrical 

- Power - Pair of 8 AWG stud post connectors for +, - 

- Data from Capacity Gauge (8 wires, RS-422) 

CHG: need +5V for wake-up from avionics, switched off during sleep mode 

TXD: output battery data to avionics every 1.7s 

* CTS, RXD: inputs for GSE only to calibrate, modify, inspect parameters 

data output: 16 byte packet (%Ah remaining, voltage, current, temperature) 

Serial output data format: 9600 baud, 8-bit, no-parity, lstop bit 

Each of the 4 lines has a dedicated return line 

Thermal 

- Starting battery temp: 10 to 50C 

- Non operating exposure: -3 1 C to 60C 

1 21 1 I97 Eric Darcyl28 1-483-9055 
NASA-Johnson Space Center, Houston, TX 

Deviations from OPS CRV 

Friwo proposes M25 lithium cell design vs M62 

- unproven shutdown separator = possible NCR for internal short 

- M25 is current production line 

waive NHB5300.4 guidelines for solder, PWB, etc, use 

high quality commercial standard 

Life requirements reduced 

MIL-W-22759 wire inspection by RIFT lab waived except 

on 270V system 

Clean room requirements waived 

12/1/97 Eric Darcyl28 1-483-9055 
Lithium-Ion Focused Session 

1997 NASA Aerospace Battery Workshop 
Page intentionally left blank 
Improved Li/BCX Primary Cells for Space Applications 

D. M. Spillman, N. M. Waite, M. F. Pyszczek, E. S. Takeuchi 

Wilson Greatbatch Ltd. 

10,000 Wehrle Drive 

Clarence, New York 1403 1 

Abstract 

LBCX (bromine chloride in thionyl chloride) primary cells have been qualified for flight aboard 

the space shuttle for over fifteen years. These cells provide high energy density while maintaining 

an excellent safety record. Recently, changes to the electrolyte have resulted in an improved 

Li/BCX 11 system with a lower self-discharge rate. The use of low molarity electrolytes in 

programs unique to NASA have improved the safety hazards tolerance of these cells while 

maintaining or increasing their energy density. 

Introduction 

Wilson Greatbatch Ltd. develops and manufactures lithium cells and batteiy packs for the medical, 

space, industrial and commercial markets. Many applications still require the use of lithium 

primary batteries with a high energy density. This is especially true for space applications. 

During the past fifteen years, the Li/BCX (bromine chloride in thionyl chloride) primary battery 

system has been qualified for flight aboard the space shuttle. The excellent safety record of the 

system, high open circuit voltage of 3.9 volts, good low temperature performance, specific energy 

of 472 Whkg and energy density of 1,035 WhfL make it an attractive power supply for a variety 

of applications (1,2,3). Presently, LiIBCX C, D and DD cells are qualified power sources for the 

video cameras, cassette tape data recorders, hand warmers in gloves, helmet lights, ANPRC-112 

survival radio and extravehicular mobility unit/personal life support system used in every shuttle 

flight. 

The LBCX system exhibits excellent voltage retention when stored for several years at room 

temperature. Data will be presented to show that advances made to the LBCX 11 system have 

improved the self-discharge previously observed. In addition, the restart capability and energy 

density of the system have been maintained or improved, particularly at lower discharge rates. In 

programs unique to NASA, the use of lower molarity electrolytes has been investigated and shown 

to enhance cell safety. LBCX 11 cells are now tolerant to a fifty milliohm short-circuit at room 

temperature and can be exposed to temperatures as high as 140°C without leaking. Despite the 

reduced salt concentration used in the electrolyte, the high, moderate and low rate discharge 

capacity goals of the various programs were still achieved. 

Chemistrv and Design Features 

All cells were spirally wound and contained in stainless steel cases and lids that have corrosion 

resistant glass-to-metal seals. Battery grade lithium was laminated onto metal grids to form the 

anode current collector assembly. The cathode current collector assembly was fabricated by 

sheeting a mixture of carbon black and polytetrafluoroethylene binder onto expanded metal. The 

separator was a nonwoven glass fiber paper. The two electrodes and the separator were wound 

together and inserted into the stainless steel case. The anode lead was then welded to the case, 

resulting in a case negative configuration, and the cathode lead was welded to the positive pin 

1997 NASA Aerospace Battely Workshop -215- Lithium-Ion Focused Session 
feedthrough in the header. The BCX and BCX I1 electrolytes were prepared by forming a suitable 

electrolyte salt in-situ in the bromine chloride in thionyl chloride solution. After filling with the 

BCX or BCX I1 catholyte, cells were hermetically sealed by welding a stainless steel ball over the 

fill hole. 

Characteristics of the Li/BCX Svstem 

The Li/BCX system offers several features which make it the power system of choice for several 

space applications. In addition to having a good safety record and high proven reliability, the 

system offers a power density as high as 95 WA and has an energy density of 1035 WM. 

Additionally, cells are rated for operation at temperatures in the range -55°C to 72°C. When 

necessary, modifications to the system allow for operation at temperatures approaching -100°C. 

Historically, the Li/BCX system has been characterized by a self-discharge rate of 3-15%/year. 

To fully characterize the LiBCX system, a large number of D cells were built and stored at room 

temperature for several years. After initially testing fresh cells under loads ranging from 0.5 ohms 

to 3010 ohms, additional cells were periodically removed from storage after 6, 12, 24 or 36 

months and fully discharged under the same spectrum of loads. As shown in Figure 1, the midlife operating voltage of the Li/BCX system remained stable even after three years of room 

temperature storage, varying by less than 200 millivolts. Furthermore, -- cells discharged under a 

1 ohm load maintained a mid-life operating voltage of over 3.0 volts. 

0 12 24 36 48 60 

Time (Months) 

Figure 1: Voltage retention of LmCX D cells 

at 50% depth of discharge. 

The capacity retention of Li/BCX D cells over a three year period is shown in Figure 2. The 

2.0 volt discharge capacity was measured and recorded for cells that were discharged after 0, 6, 

12,24 or 36 months of room temperature storage. When discharged under loads ranging from 

5 ohms to 249 ohms, fresh cells generally provided 13.5 to 15.0 Ah. Over the three year storage 

period, the cell capacity declined to about 11.0 to 12.0 Ah, representing a yearly self-discharge rate 

of approximately 9%. As expected, fresh cells that were discharged under a considerably heavier 

load of 1 ohm or 0.5 ohms provided less capacity due to increased cell polarization. Cells 

discharged under a 3010 ohm load, however, also provided lower capacity. 

I997 NASA Aerospace Battery Workshop -2 16- Lithium-Ion Focused Session 
- - 3010Q 

0.0 I"~I~"I".l... 

0 12 24 36 48 60 

Time (Months) 

Figure 2: Capacity retention of Li/BCX D cells to 2.0 volts. 

.- 

Microcalorimetry, in which cell heat dissipation is measured under load or on open circuit voltage, 

has proven to be a useful tool in estimating the lifetime of LiBCX cells (4). Recently, 

modifications and improvements to the electrolyte have resulted in the development of BCX 11 cells 

which are characterized by lower self-discharge as measured via microcalorimetry and confirmed 

via discharge test. As shown in Figure 3, when stored at room temperature over a four year 

period, LaCX I1 D cells exhibited approximately forty percent of the heat dissipation that 

Li/BCX D cells exhibited. This manifests itself in a much slower degradation in open circuit 

voltage as shown in Figure 4. While the open circuit voltage of Li/BCX D cells decreased from 

3.95 volts to 3.67 volts over a period of eighteen months, the open circuit voltage of Li/BCX 11 D 

500 1000 

Days on Test 

Figure 3: Heat dissipation of LaCX D and LiBCX I1 D cells. 

1997 NASA Aerospace Battery Workshop Lithium-Ion Focused Session 
D 

IID 

500 lo00 1500 

Days on Test 

Figure 4: OCV stability of Li/BCX D and Li/BCX 11 D cells. 

-- 

cells remained at or above 3.75 volts after four years. 

Complementing the rnicrocalorimetry results are discharge results for fresh Li/BCX D and 

Li/BCX 11 D cells. In Figure 5, the 2.0 volt discharge capacity obtained for cells tested at room 

temperature under loads ranging from 0.5 ohms to 3010 ohms is shown. Discharge capacity for 

the Li/BCX system and the Li/BCX II system is comparable under loads ranging from 0.5 ohms to 

20 ohms. At 75 ohms and 249 ohms, discharge capacity is higher for the Li/BCX I1 system, 

validating the data obtained via rnicrocalorimetry. Discharge tests under a 3010 ohm load remain in 

0.0 

.1 1 10 100 1000 10000 

Discharge Load (Ohms) 

- 

- 

- - - BCXD 

BCXIID 

I ' . . '"'1 ' ' . .'"'I ' ' ' ""'I ' ' ' ' a'"' 

Figure 5: Comparison of discharge capacity to 

2.0 volts for fresh cells. 

1997 NASA Aerospace Battery Workshop -218- Lithium-Ion Focused Session 
progress; however, capacity obtained to date with the LiIBCX I1 system has already surpassed that 

previously obtained with the LaCX system. 

Voltage delay is a common occurrence associated with liquid cathode primary lithium cells. The 

phenomenon is attributed to increased thickness of the lithium passivation layer which results in 

increased cell impedance. During application of a discharge load, the cell voltage may be initially 

depressed before recovering to a steady-state value. The voltage delay phenomenon becomes more 

pronounced when cells are stored at elevated temperature or for long periods of time and 

subsequently discharged at low temperature and high current. 

Comparisons were made of the room temperature voltage delay characteristics of Li/BCX D and 

Li/BCX I1 D cells discharged under a 4 ohm load. The results are shown in Figures 6 through 9 

for fresh cells and for cells previously discharged to lo%, 30% and 60% depth of discharge, 

respectively. For fresh cells, as shown in Figure 6, little difference existed between the voltage of 

LiIBCX and Li/BCX II cells. 

10 100 1000 

Time (Seconds) 

Figure 6: Room temperature voltage delay of 

fresh cells under 4 ohm load. 

As shown in Figure 7, differences in voltage began to emerge at 10% depth of discharge. The 

LmCX II D cells exhibited a slightly higher voltage during the first five minutes of discharge; 

however, the voltage of both systems remained above 3.0 volts following application of the 

discharge load. In Figure 8, significant differences existed between the two systems at 30% depth 

of discharge. Whereas the voltage of Li/BCX 11 D cells remained above 3.0 volts, that of the 

LmCX D cells was depressed below 2.0 volts for the initial thirty seconds of the test and did not 

reach 3.0 volts during the first fifteen minutes of discharge. When restarted after 60% depth of 

discharge, LiIBCX D cells remained between 2.3 volts and 2.6 volts for the initial fifteen minutes 

of the test. Although the Li/BCX II D cell voltage declined below 2.0 volts during the initial ten 

seconds of test, it surpassed the voltage of the Li/BCX system within thirty seconds and ultimately 

recovered to above 3.0 volts within five minutes. 

1997 NASA Aerospace Battery Workshop Lithium-Ion Focused Session 
10 100 1000 

Time (Seconds) 

- - BCX D 

BCX I1 D 

Figure 7: Room temperature voltage delay under 4 ohm load 

of cells at 10% depth of discharge. -- 

10 100 1000 

Time (seconds) 

Figure 8: Room temperature voltage delay under 4 ohm load 

of cells at 30% depth of discharge. 

1997 NASA Aerospace Battery Workshop -220- Lithium-Ion Focused Session 
- 3.0 V) - Y 

0 

5 

f& 2.0 

cd 

4 * 

P 

1 .o - - BCX D 

BCXIID 

0.0 

1 10 100 1080 

Time (seconds) 

Figure 9: Room temperature voltage delay under 4 ohm load 

of cells at 60% depth of discharge. -- 

Improved Cells for S~ace Ap~lications 

Wilson Greatbatch Ltd. and NASA have embarked on a series of development efforts aimed at 

maintaining or increasing the rate capability, capacity and temperature tolerance of Li/BCX cells 

while increasing the short-circuit hazards tolerance (5, 6, 7). To achieve these objectives, low 

molarity Li/BCX I1 cells including 0.6 M Li/BCX I1 C cells, 0.4 M Li/BCX 11 D cells and 

0.6 M Li/BCX 11 DD cells were developed for present and future applications. 

A primary goal for the next generation of cells is to be tolerant of a fifty milliohm external shortcircuit while maintaining internal continuity and not having cells leak, vent or rupture. Shortcircuit tests were conducted on low molarity Li/BCX 11 cells at room temperature using the 

designated angle iron as specified in NASA document number EP5-83-025 Rev. G. The test 

results are presented in Table 1. The 0.6 M Li/BCX 11 C cell exhibited a peak current of up to 

22.5 amps and a peak temperature as high as 85.5"C. The 0.4 M LBCX 11 D cell and the 

0.6 M Li/BCX I1 DD cell exhibited peak currents of 22.0 amps and 33.3 amps, and peak 

temperatures of 86S°C and 100.O°C, respectively. None of the 0.6 M LBCX 11 C cells or the 

0.4 M Li/BCX 11 D cells leaked, vented or ruptured as a result of these tests; however, a few of the 

0.6 M LilBCX II DD cells were observed to leak or vent mildly at the glass to metal seal. 

Table 1. Fifty Milliohm Short-Circuit Tolerance of Low Molarity Li/BCX 11 Cells. 

Peak Current Peak Temp. Physical 

(Amps) ("(3 Changes 

0.6 M C Cell 19.2 - 22.5 59.0 - 85.5 None 

0.4 M D Cell 16.0 - 22.0 77.0 - 86.5 None 

0.6 M DD Cell 26.0 - 33.3 85.2 - 100.0 Mild Vent 

1997 NASA Aerospace Battely Workshop -22 1 - Lithium-Ion Focused Session 
The temperature tolerance of low molarity Li/BCX I1 C, D and DD cells was determined by 

exposing fresh cells, cells discharged to 2.0 volts or lower, and cells subjected to a fifty rnilliohm 

short- circuit test to temperatures in the range of 100°C to 160°C at 10°C exposure intervals. The 

test results are reported in Table 2. The temperature tolerance of all three types of cells was 

maintained at or above 100°C, nearly 30°C above the maximum temperature rated for use. The 0.6 

M LiIBCX 11 C cells were tolerant to 140°C without leaking at the glass-to-metal seal. 

Table 2. Temperature Tolerance of Low Molarity Li/BCX 11 Cells. 

Minimum Leakage Temperature 

0.6 M C Cell 140°C 

0.4 M D Cell 125°C 

0.6 M DD Cell 100°C 

The specific attributes of the Li/BCX and Li/BCX 11 systems make them attractive for use in a 

number of space applications. Accordingly, NASA may use the same cell for a variety of 

applications and each application may require that a specific capacity be o6tained under a particular 

test regime. When designed, each type of cell may be required to meet minimum capacity 

requirements under various discharge loads. This allows the cell to be used interchangeably 

among the various programs and applications. 

The capacity requirements for the 0.6 M LiBCX II C cell are shown in Table 3. When discharged 

at room temperature under a 6 ohm load, cells are expected to provide a mean discharge capacity of 

4.0 Ah. When discharged under a 10 ohm or 75 ohm load, cells are expected to provide a capacity 

of 5.0 Ah and 6.0 Ah, respectively. The low molarity cell provided 4.94 Ah, 5.38 Ah and 

7.24 Ah to 2.0 volts when discharged under a 6 ohm, 10 ohm or 75 ohm load, satisfying the 

various program requirements. 

Table 3. 0.6 M Li/BCX 11 C Cell Capacity Summary. 

Discharge Load 

(Ohms) 

Requirement Achieved 

(Ah) (Ah) 

The 2.0 volt capacity requirements for the 0.4 M Li/BCX II D cell and for the 0.6 M Li/BCX 11 DD 

cell are shown in Tables 4 and 5 along with the capacity achieved. The 0.4 M LmCX 11 D cells 

provided 10.0 Ah and 12.3 Ah when discharged at room temperature under a 10 ohm or 20 ohm 

load, satisfying the program requirements of 10.0 Ah. The 0.6 M Li/BCX II DD cell was required 

to provide 22.0 Ah at room temperature when discharged under a 1.5 ohm load, 25.0 Ah when 

discharged under a 2.5 ohm load and 27.0 Ah when discharged under a 10 ohm load. The data 

provided in Table 5 indicate that these objectives were achieved. The mean capacity delivered to 

2.0 volts was recorded as 28.7 Ah, 32.3 Ah and 35.6 Ah, respectively. 

1997 NASA Aerospace Battery Workshop -222- Lithium-Ion Focused Session 
Table 4. 0.4 M Li/E3CX I1 D Cell Capacity Summary. 

Discharge Load 

(Ohms) 

Requirement Achieved 

(Ah) (Ah) 

Table 5. 0.6 M Li/BCX II DD Cell Capacity Summary. 

Discharge Load 

(Ohms) 

Requirement Achieved 

(Ah) (Ah) 

Development of the LiBCX II system coupled with the use of low molarity electrolytes has led to 

the development of improved lithium primary cells for future space missions. As shown in 

Table 6, the specific energy and energy density of DD cells has been increased from 472 Wh/kg 

and 1035 Whll to 560 Whlkg and 1256 WM, an improvement of more than 18.5%. Additionally, 

the self-discharge rate of the system is projected to decrease from 3-15%/year to 1-3%/year, further 

improving the reliability and attractiveness of the system. 

Table 6. DD Cell Capacity Comparison. 

1.14 M Li/BCX DD Cell 0.6 M LiIBCX II DD Cell 

Specific Energy 472 Whkg 560 Wh/kg 

Energy Density 1035 Wh/l 1256 Whfl 

Self-Discharge 3- 15%/Year 1-3%TYear 

Conclusions 

The Li/BCX primary system offers stable voltage and high capacity after several years of storage. 

Recent improvements to the system have led to lower heat dissipation, resulting in a more stable 

open circuit voltage, a lower self-discharge rate and improved voltage delay characteristics. Low 

molarity Li/BGX I1 C, D and DD cells have been developed to meet the requirements of future 

space applications. The short-circuit hazards tolerance of these cells has been improved and the 

temperature tolerance has been maintained. The specific energy and energy density of low molarity 

LiIBCX II cells have been increased while the self-discharge rate is projected to decrease markedly. 

1997 NASA Aerospace Battely Workshop -223- Lithium-Ion Focused Session 
References 

1. C. C. Liang, P. W. Krehl and D. A. Danner, Journal of Applied Electrochemistry, 1 1,563 

(1981). 

2. R. M. Murphy, P. W. Krehl and C. C. Liang, in Proceedings of 16th Intersociety Energy 

Conversion Engineering Conference, p. 97 (198 1). 

3. P. W. Krehl and C. C. Liang, Journal of Applied Electrochemistry, 13,45 1 (1983). 

4. E. S. Takeuchi, S. M. Meyer, and C. F. Holmes, Journal of Electrochemical Society, Vol. 137, 

No. 6, p. 1665 (1990). 

5. P. J. Size, E. S. Takeuchi, Lithium D-Cell Study, Final Report, March 1993 (contract number 

NAS9- 18395). 

6. D. M. Spillrnan, N. M. Waite, E. S. Takeuchi, NASA Flight C Cell Development Program, 

Final Report, April 1996 (purchase order number T-5789T). 

7. D. M. Spillman, N. M. Waite, E. S. Takeuchi, NASA Flight DD Cell Development Program, 

Final Report, October 1997 (purchase order number T-3 193V). -- 

Acknowledmnent 

This work was partially funded by NASA, Johnson Space Center, under contract/purchase order 

numbers NAS9-18395, T-5789T and T-3 193V. The authors would like to thank Mr. Bob Bragg 

for his useful comments and discussion related to achieving the various program objectives. 

I997 NASA Aerospace Battev Workshop -224- Lithium-Ion Focused Session 
Development of Ultra Low Temperature, Impact Resistant 

Lithium Battery for the Mars Microprobe 

H. Frank. F. Deligiannis, E. Davies. B. K Ratnakumar and S. Surampudi 

Electrochemical Technologies Group. Jet Propulsion Laboratory 

4800 Oak Grove Dr.. Pasadena. California 

and 

P. G. Russel and T. B. Reddy 

Yardney Technical Products, Znc., 

82, Mechanic St., Pawcatcuck, CT 

-- 

ABSTRACT 

The requirements of the power source for the Mars Microprobe, to be backpacked on 

the Mars 98 Spacecraft, are fairly demanding, with survivability to a shock of the order of 

80,000 g combined with an operational requirement at -80°C. Development of a suitable 

power system, based on primary lithium-thionyl chloride is underway for the last eighteen 

months, together with Yardney Technical Products Inc., Pawcatuck, CT. The battery 

consists of 4 cells of 2 Ah capacity at 25"C, of which at least 25 8 would be available at - 

80"C, at a moderate rate of U20. Each probe contains two batteries and two such probes will 

be deployed. The selected cell is designed around an approximate 112 "D" cells, with flat 

pkte electrodes. Significant improvements to the conventional Li-SOC12 cell include : a) use 

of tetrachlorogallate salt instead of aluminate for improved low temperature performance and 

reduced voltage delay, b) optimization of the salt concentration, and c) modification of the 

cell design to develop shock resistance to 80,000 g. We report here results &om our several 

electrical performance tests, mission simulation tests, microcalorimetry and AC impedance 

studies and Air gun tests. The cells have successfully gone through mission-enabling 

survivability and performance tests for the Mars Microprobe penetrator. 

1997 NASA Aerospace Battery Workshop Lithium-Ion Focused Session 
B B 

3 

F 

z Development of a Ultra- Low Temperature, Impact4 Resistant Li primary Battery for Mars Microprobe 

tL 

E3 Harvey Frank, Frank Deligiannis, Evan Davies, ? 

Kumar Bugga and Rao Surampudi 

and 

Phil Russel and To Bo Reddy 

Yardney Technical Products 

Nov. 18,1997 

NASA Aerospace Battery Workshop 

November 18-20,1997, Huntsville, AL 

SCIENCE OBJECTIVES 

Acceleration data during entry and descent 

Atmospheric pressure data 

IL 

I.4 Soil temperature data SO 

* Soil water content - spectrometer and 

electromechanical drill incorporated in the forebody. 
Jm Mars Microprobe - Mission Description 

eonurn Sensor 

Soil Sunple Chamber 

TDL H20 Expdmont Soil Cdl.ctlon Drill 

)th < 15 cm 

Nioht Dav 

Figure 1-7: Landed Configuration 
Mars Microprobe Battery Objective 

Demonstrate flight like hardware that meets performance requirements 

after representative impact test in representative thermal environment 

- 6 - 12 volts 

- 2 year life 

- 2 A-hr capacity RT 

- 0.5 A-hr capacity at -80°C 

- Survive impact 200 m/sec 

- Load profile attached 
BACKGROUND 

Li-SOCl, is the most suitable system from the polarization curves 

and discharge tests at -80°C. 

Lithium tetrachlorogallate gave improved discharge and voltage 

delay characteristics compared to tetrachloroaluminate. 

Lower salt concentrations (0.5 M vs. 1.0 M) improve electrolyte 

conductivity. 

Pancake (flat-plate) design in sliced "D" cell. 
2 

h 

a 

% 

a Jm MARS MICROPROBE CELL (D 

VERTICAL CROSS-SECTION OF THE FINAL 

i? Ah CELL DESIGN 

CATI-IODE TAB BUNDLE 

(.00l0 TI-IK. X 11 TABS) 

.016' ANNE ALE D Ni200 TEFZEL SPACER APPLIES CUMPRESSlON 

TO CELL STACK 

WELD (TYP.) 

ASSEMDL Y FIX 1 UI?I: 

SUt"f'CIK'1- T A 133 

ASSC MI31: Y I- I X I CII\!t 

111.1 I I I.IM 1'1. A I I 

- PARALL EL PI-ATE DC SIGN 

10 ~UI-L ANOIJCS 

k 1.30 in. -I 9 FULL CATHODCS 

2 HALF CATHODES 
TOPICS 

*Microcalorimetry and AC Impedance 

'Battery Motor Tests 

*Voltage Delay 

.Impact Testing 

.Accelerated Storage 

.Special Topics 

-Inspections 

-Installation and Wiring 

-Safety 

-Prior Issues 
LOAD PROFILES 

REQUIRED LOADS (ORIGINALLY) 

.Descent: first 10 min @ -40 OC 

-8 mA -4 min, 13 mA -4 min, 30 mA -2 min 

.Surface Operations for next 6 hrs 

-Sleep mode: 1 mA, -60 to 80 OC 

-Science pulses: 33 mA, -60 to-80°C, each 112 hr 

Drill: 88 mA, 10 min, -40 OC, TBD time 

-H,O Heater: 605 mA, 20 min, -60 OC, TBD time 

-Transmit: 176 mA, 15 min, -80 OC, TBD time 

NEW REQUIREMENTS FOR SURFACE OPS 

-H,O Calib, 605 mA, 2 min, -60°C,TBD time 

-Transmit: Highest Power, 15+min, -60 to -80 OC 
OBJECTIVE 

Predict capacity loss of cells for 2 year stand @ RT. 

Determine effect of temperature on capacity losses. 

Obtain 'acceleration factor' for elevated temperature 

storage tests. 

APPROACH 

Measure heat from cell from cells using 

microcalorimeter, periodically during storage. 

Measure heat from cells over a range of temperatures. 

Analyze data to make predictions. 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
(D 

a 

2 I 

W 

Observed Cell Heat Output During Storage 

Storage time, days 

Exponential decay in corrosion rate during storage 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
I Predicted Cumulative Heat Output Over 2 Years 

Storage time, days 

Predicted based on extrapolation of experimental data 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
Predicted Capacity Loss During Storage at RT 

Storage time, days 

Calculated based on a thermoneutral voltage of 3.72 V 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
Arrhenius Plot of Li Corrosion (Self Discharge) 

Activation Energy : 1.64 keaVmole 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
Predicted Capacity Loss vs. Storage Temperature 

0.4 

-50 -30 -10 10 

Storage temperatures, 'C 

Lower storage temperature desired. 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
i? 

a 4 ELECTROCHEMICAL IMPEDANCE SPECTROSCOPY I 

8 

Examine variations in the cell construction and or workmanship nondestructively. 

Examine the (lithium) interfacial conditions that would impact stability 

(life) and voltage delay. 

Detect signs of cell degradation during stand. 

APPROACH 

Establish baseline impedance spectra of the cells. 

Establish correlations, if any, with the life and voltage delay 

characteristics. 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
Impedance Spectra of Fresh Cells 

- Cell #7 1 - Cell #72 - Cell #73 

- Cell #74 - Cell #75 - Cell #77 

- Cell #78 - Cell #79 

Re Z (Ohms) 

Film Resistance varies from cell to cell. 

ELECTR0CHEM.CAL TECHNOLOGIES GROUP 
0.4 

Cell (Ohmic) Resistance 

71 72 73 74 75 77 78 79 

Cell Number 

Cell Resistance 150-210 ma; consistent with the mOhm measurement 

after film voltage delay. 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
Film Resistance 

Cell Number 

Film resistance varies from 5-40 CL 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
VOLTAGE DELAY 

(EXPLORATORY DATA ON YTP 2 AH CELLS) - --.. 

Delay at 100 mA at -40 C 

time, s 

.-..- ....... _, 

Delay at 100 mA at -40 C 

300 

time, s 

Delay at -88 C 

FILE tWlE: wdny-l TltE(sec) 

4-29-97 -BBc 25 oh load on one ha1 f D 

*Delay appears significant but not excessive @ moderate currents @ 

-40 to -60 degrees C 

*Delay could pose problems at moderate currents @ -80 degrees C 
VOLTAGE DELAY 

CONCLUSIONS 

*Voltage depression from delay is significant but not excessive at 

moderate loads @ -40 to -60 degrees C 

-Can meet early descent loads with adequate voltage@ temps to -60 degrees C 

*Voltage depression from delay is appreciable at moderate loads 

@ -80 degrees C 

-Can meet Initial transmit load but with little margin 

-May incorporate short conditioning discharge prior to transmit @ -80 degrees C 

*No voltage depression from delay later in profile @ -60 degrees 

C for H20 Expt @ 605 mA, and even high new transmit to 1A. 
BATTERY = MOTOR TESTS 

3 

B 

% OBJECTIVES 

Ib 

'DETERMINE CAPABILITY OF B ATTERY @ -30 TO -80 

P 

Y degrees C TO RUN DRILL MOTOR @ - 120 degrees C 

'DETERMINE TRANSIENT MOTOR START AND 

OPERATIONAL MOTOR CURRENTS AND VOLTAGES 

*DETERMINE TRANSIENT MOTOR STALL AND STEADY 

!? 

MOTOR STALL CURRENTS AND VOLTAGES. 
DS2 BATTERYIMOTOR TEMP TEST 

Scope triggers by motor start and records transient voltage and 

current 

Dyno set for 1 inch-lb to simulate drill 

a Motor shaft fiozen for stall 

DAS records steady state voltage and current 

VOLTAGE 

DATA 

ACQUISITION 

SYSTTEM 2 

BA'ITERY 

VOLTAGE 

- 7 

FET 

POWER 

SUPPLY 

• 

CURRENT 

L 

POWER 

SUPPLY I 

1 - 

-v + -I+ 

', 

OSCILLOSCOPE 

2 

rn 

MOTOR DYNO 

LN2 CHAMBER -120 C 
BATTERY MOTOR TESTS 

(Sample - Transient Start @ -60 degrees C) 

_ .. *Scope trace @ -60 degrees 

C 

*Peak current near 1A and 

Min V near 10.0 V 

*Lasts only 25 ms , 1 Current quickly tapers to 

near 70 mA 

@Similar results from -30 to 

-70 degrees C 
BATTERY MOTOR TESTS 

Sample - Stall Run @ -60 degrees C 

Y i $ 3.3 ' -..-. .. -1. . - -. - . ..-.- . - 

'-=I O f 2 3 4 

1. - -.-..---- 

5 

- 

6 

*Voltage near 7.7 V @ 390 rnA 

C: 

E FILE CWIE: rotorli! TIIIE(.I~) .Similar results from - 30 to -70 

7.9 9 *DAQ trace @ -60 degrees C for 

. I. ..:.. 

eta1 led motor,.btt at -68c 7-18-97 ,... ':+L':... . degrees C 

4 min. 

.Start spike noted 

: 

e 7 

r 

& 

g 4.50 4 37'' 
BATTERY VOLTAGE AFTER 4 min 

STALL OF MOTOR @ -120 OC 

-80 -60 -40 -20 0 

BATTERY TEMPERATURE, oC 
BATTERY MOTOR TESTS 

STARTING CHARACTERISTICS 

-Current near 1 A for 25 ms 

-Minimum voltage decreases with temp (8V@-70 degrees C, 4V @-80 degrees C) 

*RUN CHARACTERISTICS 

-Currents near 70mA 

-Voltages relatively stable for at least 10 min 

-Voltage level declines with temp (10V @-70 degrees C, 4V @ -80 degrees C) 

*STALL CHARACTERISTICS 

-Currents near 400 rnA 

-Voltages relatively stable for 4 rnin 

-Voltage level declines with temp (7.7 V @ - 60 degrees C) 

*BATTERY CAN OPERATE DRILL AND STALL @-60 Degrees with margin 
IMPACT TESTING 

OBJECTIVE 

*Demonstrate capability of battery: 

- Withstand 200 rnlsec impact shock 

- Deliver required electrical loads at minimum temperatures 

APPROACH 

P 

@Install cells in probe assemblies I, 

@Fire assembly into target with airgun 

.Retrieve assembly and conduct discharge tests in accord with 

mission profile and temp. 
$ 

2 

El' 

3 

IMPACT TESTING 

TEST ITEMS 

*Two cell types 

- Old: GTM + Cover with fill port in center as terminal 

- New: New GTM + Thicker cover with pin in center & fill tube on side 

- Electrode stacks + electrolyte same for old and new types 

*Two cell configurations in probe 

NEW 
IMPACT TESTING 

OVERVIEW 
AIRGUN TEST # 42 

OVERVIEW 

8 cells, old design, 1 layer config 

No venting 

But all cells had small cracks & slow leaks in GTM 

1 cell developed "open" condition" 

Remaining 7 cells delivered good output @ -80 OC 

Significance I, 

* One layer configuration appears best 

* Old GTM-Seal assembly inadequate 

* Cells survived and functioned (with leaks) 

* "Open" due to faulty pin weld 
AIRGUN TEST #42 

DISCHARGE TEST RESULTS 

@Descent Load Tests @ - 40 degrees C 

Had computer problems. 

Best estimates below 

611 ohm ........................... 3.3 - 3.5 V 

61 ohm ........................... 3.2 - 3.4 V 

10 ohm ............................. 3.2 - 3.3 V 

aDischarget Load Tests @ - 80 degrees C 

Cell # - Ah 

77.. ......................................................................... 0.85 

78. .......................................................................... 0.82 

79.. ........................................................................ .0.01 (Opened) 
AIRGUN TEST # 53 

OVERVIEW 

7 cells as 4 and 3 cell batts, new seal design, 1 layer config 

Improved external wiring 

No' leaks and no venting 

4 cell batt delivered good output on profile, also delivered 

high power transmit of 10 W @ -60 *C 

a1 bad cell in 3 cell batt limited output but this batt 

operated on constant load @ -80 OC 

', 

SIGNIFICANCE 

New seal design again withstood shock 

External wiring mods appear to have eliminated shorts 

Must investigate cause for bad cell in 3 cell batt 
AIRGUN TEST # 53 

MISSION PROFILE TEST 
ACCELERATED STORAGE 

OBJECTIVE 

*Project performance after 2 years storage. 

*Support rnicrocalorimetric projections. 

APPROACH 

*Store some cells at ambient temperature. 

~STore additional cells at elevated tem~e~rature. 

*Periodically remove and test cells. 

Compare actual with projected capacities 

*Also examine trends in voltage delay. 
ACCELERATED STORAGE 

CAPACITY TESTS @ Rr 

1 month @ RT 1 month @ 50 degrees C 

No measurable loss in capacity at 50 degrees C 

-- 

ram :t~~(n 4.": mztu) 

2.58 -. 28A 

1.m --( 1 

-8.58 i 1 1 8 1 6 

2.58 - 

0.0 2.5 5.8 7.5 !B @ !? 5 1i.R !i' 5 8.9 :2.S 5.8 7.5 18.8 12.5 15.8 17.5 

FIkew A;& Cihthrj FILE NfUlE: did TltlE(hr) \ 

1.00 - 

28-h 

-8.48 1 1 I I I I 
ASSESSMENT 1 

*Favorable Project for meeting 2 year shelf life 

-78-93% capacity retention after 2 yrs @ RT 

-Improved retentin at lower temps 

-Accelerated tests consistent with projections. 

*Battery can successfully operate drill motor 

-Sustain 1 A surge starting current 

-Run motor > 10 min @ temp c-60 degrees C 

-Sustain 1 A surge stall current ', 

-Run motor > 4 rnin @ temp < -60 degrees C 
ASSESSMENT 2 

F 

B 

b" 

*VOLTAGE DELAY 

-Effects present but not excessive for planned moderate 

initial loads @ -40 to -60 degrees C 

-Limits maximum initial load especially at very low 

temperatures, - 80 degrees C 

*IMPACT SURVIVABILITY 

-New GTM-cover design eliminated cracking and leaks and 

external wiring mods eliminated external shorting 

-Several cells delivered good post impact perf and a battery 

met profile and 10 W 

-Must determine cause for one "low" cell in Airgun #53 

-Successful repeat run would help ensure meeting qua1 
Acknowledgement 

This work was carried out at the Jet Propulsion Laboratory, 

California Institute of Technology under contract with National 

Aeronautics and Space Administration and in collaboration with 

Yardney Technical Products, Inc.,. 
Development of 20 to 50-Ah Li-Ion Cells 

for Aerospace Applications 

Lynn Marcoux and Gregg Bruce 

1997 NASA Aerospace Battery Workshop 

18 November 1997 
SLUES~AR Development of 20 to 50-Ah Li-ton Cells I 

ADVANCED 

~~CHNOLOGY for Aerospace Applications s' 

U 

5 

8 

ii? 5L Topics 

4 4 a 

m 

2 o Programmatics d 

T S 5 o Component Evaluation Development 4 

o 20-Ah Prismatic Cells 

o 20-Ah ~~lindkical Cells 

o Current Activities 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 78 November 7997 
BL UESTAR Development of 20 to 50-Ah Li-Ion Cells I - 

ADVA~VCED for Aerospace Applications 4 

Ec~NoLOGY 

5 

% 

2 Program Highlights % 

a 

% I 

8 

B o Initiated in 1991 and funded at $US 3,019k, funded 50:50 USAF:DND (DIPP) 2 

3 

P 

o Major corporate support initiated in 1996 

S 

e 

o Emphasis is on 20 to 50-Ah cells with deliverables in 1997 and 1998 

o Focus on LiCoO, positive electrodes for performance for spacecraft application 

'o Parallel development of cell chemistry and cell design 

b 

0\ o Both prismatic (wound) and cylindrical designs being evaluated 7' 

o 20-Ah Cells delivered to USAFINASA May 1997 

- > I00 Whtkg 

- > 500 cycles above 75% Qinltial 

o 50-Ah cells scheduled for delivery May 1998 ,, 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
Development of 20 to 50-Ah Li-Ion Cells 

ADVANCED 

4 ECHNOLOGY for Aerospace Applications 

b 'n 

Component Development Activities , 4 I 

8 

b3 o Component development at BATC for the most part involves evaluation of 

a_ 

9 existing materials in C cells 

$ E 

4 

o Positive electrode materials evaluated to date include 

- LixMn,O, (Riedel-de Haen, BATC, Cyprus, Kerr McGee, JEC, MaxPower) 

- L~COO, (Honjo, Seimi, Cyprus, FMC, JEC) 

- LiNiO, (Seimi) 

lb u! 

? o Near term plans for positive electrode material improvement 

- LiCoxNi,02 supplied by Westaim, FMC, ~onjo 

o Negative electrode activities have included both evaluation of commercial carbons 

in C cells and developmental materials in LilC button cells 

o Commercial Carbon materials evaluated I, 

- Lonza, KS6, KS15, KS44, KS75, SFGI 5, SFG44 

- Conoco XP Petroleum Coke 

- Nippori Carbon LK-702, LK-201 

- Osaka Gas MCMB 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 7997 
BLUES~AR Development of 20 to 50-Ah Li-/on Cells ; 

ADVANCED 

TECHN~OGY for Aerospace Applications 4 

s 

Electrolyte Studies - Rate CapabilityILow Temperature 

$ 

3 o LiCoO, vs. graphite in C cells suspended unprotected in the thermal chamber 

o All tests at room temperature ended the charge step with 1.5 hr CVC 

o Low temperature tests included charge at low temperature 

o All tests at low temperature ended the charge step with 2.5 hr CVC 

tL 

0\ 

‘P o All tests carried out between 4.1 V and 3.0 V 

o Forrate study all tests carried out with symmetrical charge and discharge rates (CI10 

charge, CI10 discharge) 

o After a charge or discharge step under low temperature conditions the cell was held for 

4 hours at rest to allow for the cell to reach equilibrium temperature 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 78 November 7997 
BLUES~AR Development of 20 to 50-Ah Li-/on Cells 

ADVANCED 

TECHN~LOGY for Aerospace Applications h 

U 

Z Capacity vs. Log Rate (C) at Various Temperatures with Various Electrolytes 5.2 Electrolyte Study: Room Temp Elctrolyte Study: -20 C 

k 

Electrolyte Study: 0 C Electrolyte Study: 40 C 

Development of 20 to 50-Ah Li-/on Cells for Aerospace Applications - 78 November 7997 

18 

I 

1 0 

14 

2 " 

P 

t 

J 1 

0 8 

0 8 

04 r 

- 12 

1 

08 

* 

+ECIDEC 

+EcmEcIwc 

-6-ECIEMC 

04 

0 2 

0 

-1 .l 2 1 -08 -0 0 -0 4 -0 2 0 

Lop (RaIe) 

- 

ECEMC 

EcmEcmMc 

ECmEC 

2 1 -0 0 4 4 4 1 0 
Sf UEW Development of 20 to 50-Ah Li4on Cells 

ADVANCED for Aerospace Applications ~CHNOLOOY h 

s 

20-Ah Prismatic Cell - First Iteration 

4 t 

2 o Welded stainless steel enclosures 

tu 

9 x - Selected welded design (tooling cost, design freedom) 9 

=4 

- 0.027" 304 stainless steel 

B - 2.69 x 12.9 x 10.9 cm (1.06" x 5.08" x 4.29") z 

b0 - Packing efficiency = 90-95% 

o Electrode dimensions 

- Positive, 600 x 8.80 x 0.017 cm 

- Negative, 630 x 9.10 x 0.01 1 cm 

tL 

2 o Standard penetrations 

- 2 GTM seals (0.090" Mo pin) 

- Fill tube ' 

o Tig welded caplcan 

o Safety Features 

- Celgard 2300" - Taped tabs !z 

% - Excess carbon capacity - 235 psi rupture disk 

S - Wide separator margin (2 x 0.25") - 2 x 0.065" Pos-Neg electrode 

$ 

a margins 2 

3 

' Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUESTAR Development of 20 to 50-Ah Li-Ion Cells 

ADVAhlCED 

TECHNOLOGY for Aerospace Applications 'r 

s 

-4 

2 20-Ah Prismatic Cell Design - First Iteration 

SECT. 8-B REMOVED 

mum 

Development of 20 to 50-Ah li-/on Cells for Aerospace Applications - 78 November 7997 
BLUES~IV~ Development of 20 to 50-Ah Li-Ion Cells 

ADVANCED for Aerospace Applications TECHNOLOGY Lr, 

s 

20-Ah Prismatic Cell Design - Second Iteration 

-A 

9 o Results of mass analysis of first generation prismatic cells 

F 9 - Caplcan 229 g 26% B 

43 - Electrolyte 216 g 24% 

- Positive electrode 260 g 29% (78% active) 

- Negative electrode 143 g 16% (47% active) 

- Other 40 g 5 % 

tL 

4 

Cf' o Mass reduction strategy 

- Reduce case thickness to 0.01 7" (from 0.027") 

- Reduce electrolyte mass to 180 g 

- Increase active material loadings 

- Explore higher capacity graphites 

- Projected specific energies of 105-1 15 Whlkg ! 

- Consider cylindrical designs 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUESTAR Development of 20 to 50-Ah Li-Ion Cells 

ADVANCED 

TEC~NOLLOGY for Aerospace Applications 

2 

'4 

Cell Capacity, Mass and Cycle Life Data 

Development of 20 to 50-Ah Li-ton Cells for Aerospace Applications - 78 November 7997 

FFFFFl wh~kg 

PO02 

PO05 

PO09 

862 

836 

848 

PO10 

PO23 

PO25 

PO27 

23.4 

20.2 

22.9 ri 

97.9 

87.0 

97.4 

768 cycles, 78% Qinit. 

300 cycles to 75%Qin,t. 

351 cycles, 67% Qin, 

21.5 

23.7 

24.2 

24.8 

.- 

90.2 

105 

104 

106 

433 cycles to 84% Q,,, 

0.01 7" case, 103 cycles to 95% Qin, 

0.01 7" case, high cap. C 

169 cycles to 83% Q,,,,. 

0.017" case, high cap. C 

', 350 cycles to h3% Qinit 
BLUES~AR Development of 20 to 50-Ah Li-Ion Ce//s - 

A~VANCED for Aerospace Applications ~ZCHNOLUGY 

+ 

Cell POI0 - Capacity vs. Cycle No. @ 100% DOD, Cl5 Rate . . 

b s 

I: 

5 0 

3 

"Y 4 

H 

4 

k? 

3 

6' 3 

, Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 7997 

25 r 

5 l5 Y 

2 

U 

a 

CI. 

0" to 

5 

0 

I 

', 

I 

0 50 100 150 200 250 300 350 400 450 500 

Cycle 
BLUES~AR Development of 20 to 50-Ah Li-Ion Cells 

ADVANCED 

s 

V 

Performance - Rate Capability, 100% DOD, Room Temp 

0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17' 18 19 20 

Capacity (Ah) 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 78 November 7997 
SLUESTAR Development of 20 to 50-Ah Li-/on Cells 

ADVANCED 

TECHN~LOOY for Aerospace Applications 4 

2 

Energy and Charge Efficiency - Cell PO03 

o Assuming all energy inefficiency results from resistive losses implies a cell 

resistance of 35 mR 

- 

CycletC 

15 

25 

50 

75 

100 

150 

200 

225 t 

o Actual cell resistances (1000 Hz and d.c. polarization) yield cell resistances of 30- 

50 mn depending on dod and cycle no. 

', 

o System mass implications 

- Thermal management (radiator size) 

- Array size 

Development of 20 to 50-Ah Li-/on Cells for Aerospace Applications - 18 November 1997 

Dishcharge 1 

AH 

20.010 

20.73 

20.33 

20.001 

19.911 

19.5n 

19.386 

19.171 

% CVC 

8.517 

8.059 

7.138 

6.449 

6.401 

5.939 

6.203 

7.243 

WH 

75.304 

71.627 

73.109 

72.320 

71.670 

70.49 

69.79 

69.017 

, Charge 

Eff (Ah) % 

99.643 

09.678 

99.004 

99.826 

100.126 

99.730 

90.630 

99.974 

Elf (Wh) % 

92.212 

92.267 

92.307 

92.397 

92.552 

02.161 

91.971 

91.948 

,AH (CC) 

19.205 

19.121 

10.916 

18.828 

10.613 

10.167 

10.251 

17.787 

WWH (CC) 

71.426 

74.104 

73.357 

73.051 

72.319 

71.0 

71.027 

69.457 

AH (CV) 

1.708 

1.676 

1.154 

1.290 

1.273 

1.166 

1.207 

1.309 

WH (CV) 

7.230 

6.770 

5.063 

5.229 

5.127 

4.606 

4.056 

5.604 
BLUESEAR Development of 20 to 50-Ah Lidon Cells 

ADVANCED 

b 

5 

2 k 

k 

a 

% I Conclusions and Future Plans (Prismatic Cells) 

8 

m s' 

3 o Wall thickness of enclosure reduced from 0.027" to 0.017" 

3 ? E 

4 o Electrolyte reduced to 180g from 200g 

o Build more cells with MCMB and LK-702 positives and fully 

characterize performance t5 

4 

'7 - Various rates and charging regimes 

- Various temperatures 

o Fully characterize safety 

o Prismatic vs cylindrical cycle life? I# 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUES~AR Development of 20 to 50-Ah Li-Ion Ce//s 

ADVANCED 

TECHNOLOGY for Aerospace Applications s 

..J 

rl Bluestar 1 -& 

3 

E! 

b 

First Iteration 20-Ah Cylindrical Cell - Design Features 

9 

G I 

a 

t3 o Standard Drawn Can 

& 

d - HU-3830 

3: - .SS 304 

a 

$ - 0.019" 

.n" - 3.50" dia. 2.50" height 

o Penetrations identical to prismatic design except for GTM 

- Two different types 

- Both pins 0.125" Mo or Ta ' 

- Ta pin is flattened. drilled and threaded (2-56) 

o Electrode dimensions 

- Positive. 1300 x 4.40 x 0.01 7 cm 

- Negative, 1330 x 4.70 x 0.01 1 cm 

', 

o Project performance 97 Whlkg (LK-702) 

- Measured Performance 109 Whlkg 

Development of 20 to 50-Ah Li-ton Cells for Aerospace Applications - 78 November 7997 
BLUES~AR Development of 20 to 50-Ah Li4on Cells 

s 

E First Iteration 20-Ah Cylindrical Cell - Cell Design 

SECT. A-A wum Crul 

GENERAL VIEW NI% 

SECT. 8-8 REMOVED 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUES~AR Development of 20 to 50-Ah Li-ion Cells r - 

ADVANCED 

ECHNOLOOY for Aerospace Applications 4 

\O 

S 

Cell Capacity. Mass and Cycle Life Data 

Development of 20 to 50-Ah Li-ton Cells for Aerospace Applications - 18 November 1997 

~~Whlkg 

COO2 

COO4 

COO6 

to 

C024 

C033 560 cycles to 70% Q,,,,. 

C034 23.8 710 cycles to 72% Q,,,, -. 

C037 802 24.7 133 cycles to 92% Qinit. 

- 

C039 806 25.6 114 286 cycles to 89% Q,,,, ' 

t839 

799 

738 

-r27.3 

25.9 

24.0 

e 

117 

117 

I 10 

2 . 

High cap. C, 446 cycles to 76% Q,,, 

376 cycles to 76% Qinit. 

Delivered to USAF 
BLUESCAR Development of 20 to 50-Ah Li-lon Cells 

ADVANCED 

TECHNOLOGY for Aerospace Applications 4 

'Q 

'0 

U 

Cell COO4 Capacity vs. Cycle No. @ 100% DOD, C/5 Rate 

', I 

0 

0 50 100 150 200 250 

Cycle 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUE- Development of 20 to 50-Ah Li-/on Cells 

ADVANCED 

4 

T&CHNOLOGY for Aerospace Applications 

Y) ..I 

, Bluestar F. . . rF 8 '2. 

Cell COO2 Capacity vs. Cycle No. @ 100% DOD, C/5 Rate 

30 

25 

20 

E 

4 

P 15 0 

m 

4 

0 

10 

5 

0 

0 50 100 150 200 250 300 350 400 450 500 

Cycle 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUE- Development of 20 to 50-Ah Li-/on Cells 

ADVANCED 

TECHNOLUGY for Aerospace Applications 4 

'0 

Y3 

U 

Capacity vs. Rate for Cylindrical and Prismatic Cells 

0.01 I I 

-1.2 -1 -0.8 -0.6 -0.1 -0.2 0 

Log C-Discharge Rate 

1. 

a 0.8 . .- 

0 

m 

a 

0" 

0.6 - 

N .-. 

B 

0" 

Z 0.4 . 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 

24 "C Prismatic -- 

L 

- 
BLUES~AR Development of 20 to 50-Ah Li-Ion Cells 

ADVMED 

RCNNOLOGY for Aerospace Applications 4 

'0 

3 

5 

2 

b 

20-Ah Cylindrical Cells - Accelerated LEO Test 

8 

4 I 

f: 

m 

'a o Two cells tested at 25 OC 

a 

3: 

% 

E o 12 Min. discharge at c rate (20 A) 

.a" 

o 36 Min. Charge 

- Constant current to 4.05 V 

- Constant current, to end of charge period 

o Over 1500 cycles > 3.00 V to date 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUES~AR Development of 20 to 50-Ah Li-lon Cells 

ADVANCED 

4. 

V) 

S 

20-Ah Cylindrical Cells - Accelerated LEO Test 

3 1 > 

0 200 400 600 800 1000 1200 1400 1600 

Cycle 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUE-AR Development of 20 to 50-Ah Li-/on Cells 

ADVANCED for A erospace A pplica tims TECHNOLOGY 

% 

s - 

'4 

5 

2 Performance - SafetyIAbuse Tests 

L 

3 

z;r I x 2 o Short-circuit (0.050 ohm), safe no firelvent. Max temperature 2 

a 8!j°C. Cell was cycled after test 

9 

7 

$ 

4 

o Oven test (150°C, 2 hours) violent ventlfire 

o Overcharge 50-Ah. 4.66V safe no firelvent 

o Overcharge test to 100-Ah, 56-Ah. 4.68V firelvent 

o Forced over-discharge (50-Ah) safe no firelvent 

- Cell charged after this abuse test showed no capacity, acting as a 

resistor 

- Cell disassembly revealed copper plating through the separator 

, Development of 20 to 50-Ah Li-/on Cells for Aerospace Applications - 18 November 1997 
BL UES~AR Development of 20 to 50-~h Li-Ion Cells 

ADVANCED 

TECHNOLOGY for Aerospace Applications 5 

'4 

JPL Test Results (S. Surampudi & K. Bugga) 

Conditioning Cycles (Cl10 charge, C15 discharge) 

22 , 

1 2 3 4 5 (i 

Cycle #I 

' Development of 20 to 50-Ah Li-ton Cells for Aerospace Applications - 18 November 7997 
BLUES~AR Development of 20 to 50-Ah Li-Ion Cells' 

ADVMCED 

~&CHNOLOGY for Aerospace Applications s '4 

2 L E 

b 

a JPL Test Results (S. Surampudi & K. Bugga) G I 

8 60 A, 120 msec Pulses at Room Temp. & 100% SOC !$' 

3 4.1 I I 

time,-s 

Development of 20 to 50-Ah Won Cells for Aerospace Applications - 18 November 1997 
BLUES~AR Development of 20 to 50-Ah Li-/on Cells 

ADV~ED 

TECHNOLOGY for Aerospace Applications 4 

'0 

'0 

'4 

JPL Test Results (S. Surampudi & K. Bugga) 

60 A, 120 msec Pulses at Room Temp. & 20% SOC 

time, s 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 78 November 1997 
BLUE- Development of 20 to 50-Ah Li-Ion Cells 

ADV~CED 

s 

-4 

2 

JPL Test Results (S. Surampudi & K. Bugga) 

40 A, 60 sec Pulses at Room Temp. & 100% SOC 

- RS24(97.1 %,SOC) fp 

- RS24(94.2%~ SOC) 

2.6 - 

', 

cell# ns 24 

2.2 - -A I 

I) 20 40 ii0 80 

time, s 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUES~AR Development of 20 to 50-Ah Li-ion Cells 

ADVANCED 

TECNNOLLOGY for Aerospace Applications LC 

'0 

S 

JPL Test Results (S. Surampudi & K. Bugga) 

40 A, 60 sec Pulses at Room Temp. & 40% SOC 

- DS24(45.5"/;, SOC) 

- l?S24(42.6% SOC) 

- DS24(39.7%, SOC) 

After 2nd tliscl~nrgc n t 25 C , 

I -..I. 

, Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 78 November 7997 
BLU~AR Development of 20 to 50-Ah L ;-/on Cells 

ADVANCED for Aerospace Applications La 

8 

First Iteration 20-Ah Cylindrical Cells - Conclusions 4 I 

1 

l? o Overall high rate capability of present cylindrical design is poor with respect 

d 

s, to power $ E 

.a0 o As expected the low temperature performance of the design is also poor 

o To date the safety of this design does not appear to be an overriding issue 

with possible exception of the high temperature oven test and extended 

overcharge 

o Larger diameter lower pressure rupture disc will be added safety feature 

o Change length to diameter aspect ratio to improve high rate and low 

temperature performance '< 

Development of 20 to 50-Ah Li-/on Cells for Aerospace Applications - 18 November 7997 
BLUESTAR Development of 20 to 50-~h' Li-Ion Cells 

ADVMCED 

TECHNOLOGY for Aerospace Applications 4 

s 

-;I 

5 

2 Near Term Plans 

P 

a 

2 I 

2 o Continue present activities 

2 - Continue component developmentlimprovement (carbon) 

9 - Continue design developmentlimprovement (safety features, current collection) 

F - Fabricate and test next generation 20-Ah prismatic and cylindrical cells (Lot 1 in formation) 

$ - Complete fabrication of Lot 1 50-Ah cells 

% Safety and abuse tests 

- Cycle life tests (including accelerated GEO testing) 

o Develop pre-pilot scale (40, 50-Ah cellslweek) production capability-(May '98) tL 

Q - Present capability is 10, 20-Ah cells /week f 

- Capital expenditure of $CDN 2.5 million ' 

o Initiate scale-up (to 100-Ah) design analysis 

- Cell Model (current density, aspect ratio, thermal, resistance) 

- Compare cylindrical vs. prismatic 

- Determine scale up limits ', 

' Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
LI, 

'0 

SLUES~AR Development of 20 to 50-Ah Li-/on Cells I

S 

5 Next Generation of 20-Ah Cylindrical Cells 

2 

b 

3 

4 I 

o Pursue new design with better ratellow temperature capability 

R 

Q. 

ADVANCED 

o Standard Hudson Tool and Die can 

- HU-3570. 0.019" 304SS. 2.625" diameter, 4.5" length 

for Aerospace Applications 

k 

s ' o Penetrations similar to previous designs except for GTM 

1 'V 

- GTM is larger and Ta pin (0.187") is flattened and treaded (4-40) 

TECHNOL UGY Bluestar/) 

o Electrode dimensions tL 

\O 

CT - Positive, 660 x 8.80 x 0.017' cm 

- Negative. 680 x 9.10 x 0.01 1 cm 

o Projected performance 

- 1 10 Whlkg (LK-702). 127 Whlkg (MCMB) 

- 220 Whll (LK-702). 240 Whll (MCMB) 

b 

P 

2 

r 3 o First lot of cells presently undergoing formation 

r/, 

8, 

6' 3 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
BLUE- Development of 20 to 50-Ah Li-Ion Cells 

ADVANCED 

TECHN~LOGY for Aerospace Applications 

+ 

3 

'4 

2 b 

Q Next Generation of 20-Ah Cylindrical Cells - Cell Design b 

b 

0.m ar. 

d 

SECT. 8-0 REMOVED 

I)~~smW.MmUU an 

HIW. tns tw 3 to. rm. e.c,uc. or I 

I d d I 

SECT. A-A IUCULP CIU) 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 78 November 7997 
BLUESAR Development of 20 to 50-Ah Li-Ion Cells. 

ADVANCED for Aerospace Applications TECHN~LOGY 

+ 

\O 

50-Ah Cylindrical Cell Design Features 

o Standard Hudson Tool and Die drawn can 

- SS 304 

- '0.019" 

F s - 2.56" dia, 7.09" height !z 

4 

o Penetrations similar to previous designs except for GTM 

- GTM is larger and Ta pin (0.187") is flattened, drilled and treaded (4- 

40) b 

\O 

Y 

- Investigate lower pressure rupture disc 

o Electrode dimensions 

- Positive, 675 x 16.1 x 0.017 cm 

- Negative, 690 x 16.4 x 0.01 1 em 

'# 

o Projected performance 

- 11 7.7 Whlkg (LK-702), 125.6 Whlkg (MCMB) 

- 278.1 Whll (LK-702), 319.6 Whll (MCMB) 

Development of 20 to 50-Ah Li-Ion Cells for Aerospace Applications - 18 November 1997 
Page intentionally left blank 
tL 

\O 

Rechargeable Lithium-Ion Battery 

'p Development at Eagle-Picher 

Industries 

Chad Kelly I, 

Advanced Electrochemical Systems Group 

Federal Systems Department 

Eagle-Picher Industries, Inc. 

Joplin, Missouri 64802 


B 

% Three configurations of electrodes 

+ stacked - prismatic 

+ spiral wound - cylindrical 

+ pseudo spiral - prismatic 

Cell case materials 

', 

+ stainless steel 

+ Surlyn coated aluminum bags 

+ Aluminum 
Spiral-Wound Cell (7OOmAh) at 42% DOD 

-4 Celcius at C15 

1 4.5 

0.9 4 

4 

'0 

'0 

g $ 

2 

Cycle Number 

Technologies Division i 

3.5 0, P- 3 al!! 

5 O 2.5 .I aL 

2 :8 

08 1.5 , 0 

1 3', 
9 FEDERAL a- - I 

4 

!------ Spiral-Wound Cell (700mAh) at 14% DOD 

-4 Celcius at 615 

L Cycle Number 
Cycle Life Characteristics of June 96' 

Li-ion Technology (7OOmAh cell) 

Cycle Number 

- 

- 

Tests in Progress, 

Projected to EOL 

EOL 

I I I I I 
Cycle Num ber 
28 Volt-1 5 Ah Prototype Eagle-Picher Battery 

Preliminary Cell Capacity Performance 

Mean Capacity 14.275 Ah 

Standard Deviation 22.8 mAh 
1 5 Ah Cell Weight Attributes 

% Header 

% Electrolyte. 3% 1 
>70 @ C PI00 @ CIIO 

Test in Progress 

20 mAh - 100 Ah 
Page intentionally left blank 
- Evaluation of 20 Ah Li Ion Cells 

Marshall Smart, B. Y. Ratnakumar, C. K. Huang and S. Surampzldi 

Electrochemical Technologies Group, Jet Propulsion L-a-borato~ 

4800 Oak Grove Dr., Pasadena, CA 

Carole Hill and Dan Ra&ykewycz 

Air Force Research Lab, Albuquerque, h4.l 

R A. Marsh 

Air Force Research Lab, Dayton, OH 

ABSTRACT 

Lithium ion cells of 20 Ah capacity were fabricated by Bluestar Advanced 

Technology Corporation, Canada under a developmental contract fiom US Air Force. In this 

paper, we report our studies on the evaluation of these cells under various test conditions. 

These include generic test conditions such as discharges and charges at different 

temperatures to understand the rate-limiting processes in the dischargelcharge processes as a 

hnction of temperature, and cycle life under standard cycling conditions (100 % DOD) at 

ambient temperature. In addition, tests are being done to ascertain the performance of the 

cells in the Mars 2001 Lander application, which includes pulse testing of the cells at 60 A 

and 40 A loads for 100 mS and 1 min., respectively at different states of charge and 

temperatures, and cycling at low temperature at partial depths of discharge. , 

1997 NASA Aerospace Battery Workshop Lithium-Ion Focused Session 
EVALUATION OF 20 Ah LITHIUM ION 

CELLS (BLUESTAR) 

Marshall Smart, Kumar Bugga 

Chen-Kuo Huang and Rao Surampudi 

Electrochemical Technologies Group, JPL 

and 

CarolHiIl, Dan ~adzykewyci and R. A. Marsh 

Air ForceResearchLaboratory 

November 18,1997 . Y'. ,.: 

NASA Aerospace Battery W orbs ho p 

November 18-20, 1997, Hunlsville, AL 
B lue Star Lithium -Ion Testing Flow Chart' 

L Blue Star Cell8 . 1 

I lnoom ing Phyrlcal lnrpeotlon 

(A II C a 11.1 I Phyrlca'l~in8p~~flon 

Electrical lnrpeotlon I 

I rn 

Condittionlng Cycler 4 

(A I1 Cellr) 10 oycles 

I 1 I 

Electrical C haracterization 

(3 Cells) 

I 

Cycle Life Characterlzation 

(5 Cells) I 

Discharge Characterization 100 % DOD Cycle Life Test Mars 2001 Lander Life 

Simulatlon Tert 

1 Charge Characterization 

I 

Tem perature Effects 1 

I I 

Self-Dlscharge Tert8. 1 

Mars 2001 Lander Life 

Simulatlon Tert 

(50 W DOD Cycle Llfe Teat) I 
Characterization Tests 
Weight Distribution 

Weights 

0 

of 20 Ah (Blue Star) Li Ion Cells 

BS06 BS08 BS09 BS15 BS18 BS19 BS24 BS23 

Cell # 

Cell weight more consistent among the cells compared to the cathode weight. 
Weight vs. Capacity 

Cell weight 

L3 - r\ 3 

- / - 

A 

Cathode weight 

- - 

1, 

I I 

20 Ah (Blue Star) Li Ion Cells 

22 23 24 25 

Capacity, Ah 

Capacity increase with increasing cathode weight. 

850 250 
Conditioning Cycles 

Conditioning Cycles of 20 Ah (Blue Star) Li Ion Cells 

* Cell capacity around 23 Ah 
AC IMPEDANCE OF 20 Ah CELL 

Impedance reflective of a slow charge transfer process followed by 

diffusion. 
2 

3 

2 b 

2 

b 

s 

EIS of JPL Experimental Li Ion Cells 

Re Z (Ohms) ! 

Cathode impedance dominant portion of the total cell Impedance. 
Pulses at 60 A at 25OC 

Pulsing of Blue Star 20 Ah Cell at 60 A (120 mS) 

- BS24(19.7%SOC, pl) 

- BS24(19.7%SOC, p3) 

u u uu 

- 

1, 

After 3rd discharge at 25 C 

I I 

time, s 
Pulses at 60 A at 25OC 

Pulsing of Blue Star 20 Ah Cell at 60 A (120 mS) 

4.1 

time, s 

3.7 

3.6 

3.5 

- ! 

100 % SOC at 25 C 

- 

Cell# BS 24 

I I I 

0 1 2 3 4 
Pulses at 40 A at 25OC 

Pulsing of Blue Star 20Ah Cell at 40 A (60 S) 

-BS24(100%SOC) 

- BS24(97.1% SOC) - - BS24(94.2% SOC) 

- 

- 

25 C 

- 

Cell# BS 24 

I I I 

time, s 
Pulses at 40 A at 25OC 

Pulsing of Blue Star 20 Ah Cell at 40 A (60 S) 

rn - BSZ~(~Z.~% socj 

- BS24(39.7% SOC) 

I After 2nd discharge at 25 C 

0 2 0 4 0 6 0 80 

time, s 
Pulses at 60 Aat -20°C 

Pulsing of Blue Star 20 Ah Cell at 60 A (120 mS) 
Pulses at 40 Aat -20°C 

Pulsing of Blue Star 20 Ah Cell at 40 A (60 S) 

20 40 68 80 

time, s 
Discharge at 25OC 

Cell# 18 at 20 C 

- - 7.5 A 

-+- 2.3 A 

+ 4.6 A - 

Charge at C110 to 4.1 V 

followed by taper to CIS0 

', I I I I 

Capacity, Ah 

Nearly 100 % utilization at C/3 discharge rate. 
Charge at 25OC 

+ 2.3 A Charge cvmnt (CAO) I1 

I + 7.5 A chaqe cvmnt (CB) 

Charge current to 4.1 V 

Constant voltage charge taper to C/SO 

Temperature = 23'~ 

Cell BS 019 - 

', - 

I I I I 

Time (Hours) 
Charge at 25OC 

- 23 A Charge cumnt (C110) 

- 4.6 A Charge current (C15) 

- 7.5 A charge current (CB) 1, 

I I I I I I I I I I I I 

0 2 4 6 8 10 12 14 16 18 20 22 24 26 

Charge capacity (Ah) 

Above 213 capacity in constant current mode at highest rate tested, i.e., C/3/ 
Charge at 25OC 

4.2 

4.0 

Temperature = 23'~ 

3.8 

E 

a 3.6 w 

ed - .c, 

0 

5 

U 

3.2 - 

3.0 - 1% 

I 2.8 I I I I I I I I I I I 

0 2 4 6 8 10 12 14 16 18 20 22 24 26 

Discharge capacity (Ah) 

Nearly 100 % charge input in 3 hours of charging. 
Discharge at O°C 

Charge current 2.3 A (C1lO)to 4.1 V 

Constant voltage charge taper to CIS0 

4.00 

Cell BS 019 

- 

+ 2.3A Discharge (C110) 

-~c- 4.6A Discharge (CIS) - 

+- 7.5 A Discharge (C13) 

2.80 I I t I I 

0 5 10 15 20 25 

Discharge capacity (Ah) 

Nearly 90 %capacity realized at C/3 discharge rate at 0 OC 
Charge Profiles at O°C 

3 

2 

1 

0 

5 10 15 20 25 

Charge Capacity (Ah) 
Discharge at -20°C 

-2.3 A Dis charge current (dn0) 

0.0 0.5 1.0 1.5 2.0 ' 

Discharge Capacity (Ah) 
Capacity vs. Temperature 

-40 -30 -20 -10 0 10 20 30 40 

Temp., OC 

Poor performance at -20°C 
Ragone Plot 

70 80 

Energy, Wh 
Cycling (100% DOD) at 25OC 

Constant cumnt charge = 4.6 A (C15) 

Constant voltage (4.1 V) taper to 0.46 A 

Discharge current = 4.6 A (Cl5) 

Discharge cut-off voltage = 3.0V 

Cycle Number 

Capacity fade 0.025 % per cycle. 
EIS During Cycling 

Marginal increase in the charge transfer (Faradaic) impedance 

* Additional charge transfer polarization. 

Decrease in the series (Ohmic) resistance. 
Cycle Number 

Performance at -20°C inadequate. 
Cycling at 50% DOD at OoC- Mars 2001 Lander 

3.8 

Constant current charge = 4.6 A (C15) 

Constant voltage (4.1V) taper to 0.46 A 

Discharge current = 4.6 A (Cl5) 

Discharge cut-off = 10 Ah 

3.2 

0 ' 50 100 150 200 250 

Cycle Number 

Depression in the EOD Voltage : 0.2 mV1cycle. 
SUMMARY 

& 

s B 2 

i$' Bluestar 20 Ah cells have so far shown good cycle life characteristics 5 9 

s! both at 100 % and 50 % DOD, which encouraging from the Mars 

$ 

E 

+% 

2001 mission point of view. 

Charge and discharge characteristics are fairly good at moderate 

G 

W 

\P rates. 

Low temperature performance needs to be improved to enable the 

Mars Exploratory missions. la 
Acknowledgement 

This work was carried out at the Jet Propulsion Laboratory, 

California Institute of Technology under contract with National 

Aeronautics and Space Administration and in collaboration with the 

Air Force Research Laboratory. The cells were fabricated by 

Bluestar under Air Force contract. 
The Li-ion Battery Qualification 

Programme for STRV-I d 

Vijay V. Thakur 

Space Power Laboratory 

Defence Evaluation and Research Agency, UK. 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Overview 

+ Mission background 

+ Battery requirements 

+ Qualification programme details 

+ Test results 

+ Conclusion 

@I AFA 1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
STRV-1 c&d mission background 

+ 100kg microsatellites in GTO as Ariane V piggyback 

+ Technology demonstration mission in high radiation orbit 

exposed to space environment from LEO to GEO 

+ Approximately 750 eclipses per year 

+ Fully regulated 28V bus, GaAs and InP solar arrays, NiCd 

primary battery I* 

+ Li-ion battery experiment (LIBE) on STRV-Id 

1997 NASA Battery Wosksliop, Huntsville, 18th - 20th Nove~iiber 1997 
LIBE mission obiectives 

- 

+ First flight of Li-ion battery in space 

+ Operate entire mission following LEOP with Li-ion battery 

+ Primary NiCd battery in 'standby' mode, autonomous 

switchover to NiCd in the event of a Li-ion battery 

performance anomaly 

+ Monitor battery voltage, individual oell voltages, batte.ry 

temperature 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Li-ion battery requirements 

+ Battery voltage : 17.OV I V,,, I 24.OV 

+ Battery delivered energy : 46Wh @ 60% DoD 

+ Battery operating temperature : -1 5°C I T,,, 5 45°C 

+ Battery total mass < 4.5kg, individual module dimensions 

< 175mm x 155mm x 40mm 

+ No effect on battery delivered energy due to 'sinusoidal' 

charge current I, 

+ Pulse discharge performance to be demonstrated 

+ 2 1000 life cycles at 46Wh discharge each with 

V,,, > 17.0V n 

fa AEA 1997 NASA Battery Workshop, Huntsville, 18th - 20111 November 1997 DERA 
b 

a I 

B 

a 

R 

tsr 

a 

Qualification programme details 

3- 

!$ 

4 

+ Discharge performance (delivered energy) for Cl5, Cl2 

and 2C discharge to 17.OV (ambient temp.) 

C 

P 

?' 

+ Discharge performance for C/5 discharge to 17.OV over 

temperature range -1 5°C I T,,,, 5 45°C (I 5°C intervals) 

+ Discharge performance for sinusoidal charge current 

profile between OmA and 600mA 

P + Pulse discharge performance for 2C pulses, lsec 'ON', 3 

si 

S 9sec 'OFF' 

3 

9 9 + Life cycle test : charge at 1.08C to 24.OV, discharge at 

B 

k? 

2. 0.6C to deliver 46WhIcycle S /, 

1997 NASA Battery Workshop, Huntsville, 18d - 20111 November 1997 DERA 
g I 

R 

(D 

2 

Li-ion battery physical parameters 

P 

9 

3 a D 

4 

+ 12 x SONY 26650 cells, 2.7Ah nameplate capacity 

+ Recommended cell operating values : V,,, = 4.2V, V,,, = 2.5V I 

W 

- 

I 

+ Two battery modules in parallel, 6 series cellslpack 

+ Battery mass (total) : 1.74kg 

+ 62.3Whlkg (C/5 discharge, 100% DoD, 20°C ) 

+ Battery volume (per module) : 183mm x 94mm x 31 mm 

+ 61 .OWhll (Cl5 discharge, 100% DoD, 20°C ) 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 
ery modu 

1997 NASA Battery Workshop, I-iuntsville, 18111 - 20th November 1997 
Discharge performance tests - delivered energy 

0.5 

Discharge rate (C) 

- - -. Operating temperature = 20 deg.C I j 

Discharge to 17.OV ' I 

0.6A charge to 24.OV - - - - - - - - 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 
Discharge performance tests - battery voltages (1 ) 

C/5 discharge 

0.00 1.00 2.00 3.00 4.00 5.00 6.00 7.00 8.00 

Time (hours) 

C/2 discharge 

Time (hours) 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 
Discharge performance tests - battery voltages (2) 

2C discharge - pack A 2C discharge - pack B 

0.00 1.00 2.00 3.00 4.00 5.00 

Time (hours) 

0.00 1.00 2.00 3.00 4.00 5.00 

Time (hours) 

- - -- . 

1997 NASA Battery Workshop, Huntsville, 18111 - 20th November 1997 
Discharge performance tests - 

Individual cell voltages during discharge 

Time (hours) 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Temperature performance tests - 

6 (D 

B 

a, delivered energy 9 

F 8 S 

Test con&-- 

'0.2C discharge to 17.OV 

10.6A charge to 24.OV 

I 

I 

15 

Cell temperature (deg. C) 

1997 NASA Battery Worksl~op, Huntsville, 18th - 20111 November 1997 
Temperature performance tests - 

battery discharge voltage 

w 

increasing temperature 

a AFA 

Time (hours) 

- .. - - - 45 degC 

- - - - - - - 30 degC 

---.,I5 degC 

- - -, 0 degC 

-15 degC .- . . -. . - . . . . - - . . . . . - - 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Sinusoidal charge current tests - 

current profiles 

Time (seconds) n 

1997 NASA Battery Workshop, Hu~~tsville, 18th - 20th November 1997 DERA 
Sinusoidal charge current tests - 

energy delivered 

AFA 1997 NASA Battery Workshop, Huntsville, 18tl1 - 20th November 1997 DERA 
Sinusoidal charge current tests - 

battery voltages 

1997 NASA Battery Worksliop, Huntsville, 18111 - 20th November 1997 DERA 
Pulse discharge performance tests - 

battery voltages 

Time (seconds) 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Life cycle test - battery voltages 

0 100 200 300 400 500 600 700 800 900 1000 

Cycle number 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Life cycle test - 

individual cell voltage spread 

I 

Cycle N uin ber I 

. . . .. - . . - .- . . ... . .. . .. .. . .. . .- -. . . 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 
Test results summary 

+ Energy requirements met for Cl5, Cl2 rate, not for 2C rate 

+ Energy requirements met for 0°C I T,,, 5 45"C, not for - 

+ No adverse affect of sinusoidal charge current profile on 

delivered energy 

+ > 1300 2C pulse discharges achievled 

+ 1000 cycles under life test conditions achieved 

*r, 0 

0 

E 

% 

+ Mass and volume requirements met except for x-axis 

t/, 

2 

g dimension (acceptable in current spacecraft configuration) #, 

1997 NASA Battery Workshop, Huntsville, 18th - 20th Noveniber 1997 DERA 
Conclusions 

+ Required performance at 2C rate and -15°C can be met if 

V,,, operating limit reduced 

+ Veo, = 17.OV (2.83VIcell) is above manufacturer's limit of Veo, = 15.OV 

(2.50VIcell) 

+ Pulse discharge and life performance meets STRV-1 d 

requirements 

', 

+ life cycling to continue at new test facility, December 1997 ' 

+ Capacity reduction vs. cycle life data still to be analysed 

+ increasing recovery time result not yet explained 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 
Acknowledgments 

+ Funding 

+ British National Space Centre - ATS1313 programme 

+ AEA Technology 

+ Engineering team (AEA Technology - Space and Defence 

Systems Group) 

+ Mike Cooke (electrical design1 test programme manager) 

+ Charlie McCarthy (mechanical design I modeling) 

+ Carl Thwaite (test data analysis) 

1997 NASA Battery Workshop, Huntsville, 18th - 20th November 1997 DERA 
Page intentionally left blank 
Rate Capability Analysis 

;rl 

2 3 

P a iz 

of Li-ion Cells 

% 

d> 

a 

Y 

Esther Sans Takeuchi and 

Randolph A. Leising 

Wilson Greatbatch Ltd. 1, 

!? s 

E' 

3 

f 

10,000 Wehrle Drive, Clarence NY i4031 

"rl 8 
Outline 

2 'a 

R 

(D 

bJ 

E, 

Introduction 

3 

a 

q Results 

% Specifications for WGL Li-ion Cells 

Safety Test Results 

Cycle Test Results 

High Rate Testing 

Continuous Discharge 

Pulse Testing 

Summary 
Energy Density Comparison for, I 

Primary and Secondary Batteries 

Lead NiCd NiMH Li-ion LVSVO LVCSC 

Acid 

Cell Type 
WGL Li-ion Specifications 

Energy Density 

Cell Size 

Weight (g) 

Volume (cc) 

Capacity (Ah) 

Whlkg 

WhIL 

19 38 62 142 

7.2 14.8 24.9 53.7 

0.55 0.75 1.95 4.40 

107 73 It 116 115 

283 188 290 ' 303 
WGL Li-ion Specifications 

Voltage Profile 

WGL D-Size Li-ion (750 mA Discharge) 
WGL Li-ion Safety & Abuse Testing 

Short circuit testing (BOL) - no vent 

Short circuit testing after 

extended cycling - no vent 

Force overcharge testing - no vent 

Force overdischarge testing - no vent 

Shock test - no vent 

Max currentlmax temp I, - no vent 

Initiated DOT crush test - passed test 
WGL Li-ion Performance Testing 

Cycle testing (25"C, 37"C, 50°C) 

High cycle number testing (%DOD) 

Rate capability (continuous discharge) 

High rate pulse discharge 

Storagelself discharge characterization 
WGL Li-ion Cell Performance 

WGL C-Size Li-ion (750 mA Discharge) 

100% DOD 

', 

0.0 1 , 

I I I I 

I 

0 100 200 300 400 500 

Cycle Number 
vv UL LI-IWII Lell rerrormance 

WGL C-Size Li-ion (1.0 A Discharge) 
Li-ion Rate Capability Testing 

Accelerated Rate Test Procedure 

Fully charge cell to +4.1 V 

Discharge cell at current densities of 8,4, 2, 1, 

0.5, 0.25 and 0.125 mAlcm2 to +2.75 V without 

charge steps between discharge steps 

Allow cell to rest for 5 min between discharge 

steps 

Doyle, Me; Newman, J.; Reimers, J. J. Power 

Sources, 1994,52,211-216. 
Li-ion Rate Capability Testing 

Accelerated Discharge Rate Test Procedure 

WGL C-Cell(8,4,2,1,0.5,0.25,0.125 mAlcm2) 

Time (h) I 
+ 

'a 

'0 

'-I Li-ion Rate Capability Testing s 

52 

b 

(D 

a 

WGL C-Size Li-ion Accelerated Rate Data 

2 I 

D 

tu 

9 % Discharge Currrent Capacity (% of Total) 

Current Density 
Li-ion Rate Capability Testing 

WGL C-Cell Discharge Rate Test 

- 4.0 Amp 

1 .O0.5 - 

- - 2.0 Amp 

- 1.0 Amp 0.5 Amp 

0.0 I l'l'l'~'~g~v~-~ I*, ' 

0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0 ' 

Capacity (Ah) 
WGL G-Ge Discharge Rate Test Comparison 

Accelerated Full Cycle 

w 2.0 Amp 

m 1.0 Amp 
Li-ion Rate Capability Testing 

WGL D-Size Li-ion Accelerated Rate Data 

ki? 3 

P r: Discharge Currren t Capacity (% of Total) !z 

4 Current Density 
Li-ion Rate Capability Testing 

WGL D-Cell Discharge Rate Test 

4.5 

4.0 

n 3.5 

> - 3.0 

& 2.5 

a - 2.0 

0 

> 1.5 - 

1 .o - - 

0.5 - - 0.75 Amp 

0.0 1 

I I I 1 I I I I I I 

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 

Capacity (Ah) 
Li-ion Rate Capability Testing 

WGL D-Size Accelerated Charge Rate Data 

Charge Currrent Capacity (% of Total) 

Current Density 
WGL 
Li-ion High Rate Pulse Testing 

WGL C and D-Size Li-ion Cells 

Cell Voltage at 50% DOD 

Prepulse V Pulse V RDC 
Li-ion High Rate Pulse Testing 

WGL D-Size Li-ion (10 Amp12 sec Pulses) 

4.5 

4.0 

3.5 

3.0 V 

a, 2.5 

cn 

Q 

- w 2.0 

0 

> 1.5 

1 .o 

0.5 

0.0 

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 

Capacity (Ah) 
Summary 

* WGL has developed cylindrical, hermetically- I 

sealed Li-ion batteries in the AA, 1/2C, C & D sizes 

WGL Li-ion cells display high cycle life under a 

variety of test regimes 

An accelerated rate test was used to evaluate the 

rate capability of Li-ion cells 

* WGL Li-ion cells provide high rate capability 

compared to commercially available Li-ion cells 
Page intentionally left blank 
Nickel-Hydrogen Session 

1997 NASA Aerospace Battery Workshop 
Page intentionally left blank 
MARS PATHFINDER BATTERY : 

PERFORMANCE 

B. OTZINGER, D. PERRONE, S. DAWSON, T. VALDEZ 

C 

00 

Y S. SURAMPUDI , R. EWELL, M. SHIRBACHEH, 

I 

8 NASA BATTERY WORKSHOP 

B 

& Y 

R. a 

HUNTSVILLE, ALABAMA 

% 

a NOVEMBER 18-20,1997 ? $. 

3 ELECTROCHEMICAL TECHNOLOGIES GROUP 
OUTLINE 

MISSION REQUIREMENTS 

BATTERY DESIGN FEATURES 

BATTERY OPERATIONAL OVERVIEW 

BATTERY PERFORMANCE 

- PRELAUNCH 

- CRUISE 

- EDL 

- MARS OPERATIONS 1, 

SUMMARY AND CONCLUSIONS 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
ARS PAIHF NDER LANDER 
MISSION REQUIREMENTS 

ti 

a_ ..+ PRELAUNCH 4 MONTH ACTIVE INVERTED STORAGE 9 

P 8 E 

.a0 LAUNCH INVERTED BATTERY LAUNCH 

* CRUISE 7 MONTH ACTIVE STORAGE 

I 

W 

\O 

Y EDL 40 AH, 1080 WH 

MARS OPR. CYCLES 30 CYCLES(lCYCLEIDAY, 8 

HOUR CHARGE 16 HOUR 

DISCHARGE) 50% DOD 

WEIGHT 15 KG 8, I 

g 

s k k 

DIMENSIONS 

B 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
PF Ag-Zn BATTERY 

ELECTROCHEMICAL TECHNOLOGIES GRC 
- 

% JPL 

5 

MPF BATTERY SPECIFICATIONS 

VOLTAGE 

CAPACITY 

RATE CAPABILITY 1-5 A 

G 

\O 

P PULSE CAPABILITY 40 A FOR 40 MSEC 

CYCLE LIFE 40 

WET LlFE 

WEIGHT 

DIMENSIONS 

14 MONTHS 1. 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
JPL - 

7 

V) 

'0 

U 

5 r4 KEY CELL DESIGN FEATURES 

6 Robust separator system - 5 layers of cellophane and 2 layers of polymer membrane. 

3 

s 

To achieve long calendar and cycle life 

G Triple redundant case-to-cover seal including basic ultrasonic seal. 

To prevent electrolyte leakage 

To improve safety 

Large cell plate area - approximately 200 square inches. Y Low temperature operation 

For enhanced pyro-firing 

Unique leak-free cell vent valves. 

For inverted battery operation 

allow gas venting under off-limit operation 

3 

- Minimal free electrolyte 

V Q 

a For inverted battery operation a 

"", 

3 

2 

r/, 2. 

a ELECTROCHEMICAL TECHNOLOGIES GROUP 
- 

'r 

'a 

-JPL 

'a 

'4 

5 

e KEY BATTERY DESIGN & PROCESS 

FEATURES 

Isj* 

e Titanium fabricated battery case with sealable cover and gasket. 3 . 

F Light weight construction 

B 

+8 gas and electrolyte containment 

Battery vent valve and pressurization port. 

Redundant valve to protect cells from electrolyte loss 

C 

\D 

9' . Battery heater and two temperature sensors. 

Thermal management for charge control 

Over-pot of cells, surface conformal coating, connector back side potting and 

electrolyte absorption system. 

Prevent ionic conductive paths \ 

Improve safety 3 

R 

E .Cell Matching and selection k 

a, a 

% a Extend cycle life 

8 

8. 

0 

3 ELECTROCHEMICAL TECHNOLOGIES GROUP 
KEY BATTERY OPERATING STRATEGIES * 

2 

d 

R 

m 

B 

B Battery was mounted inverted and maintained at 12 C during pre-launch and d 

P launch Phases 

0, 

$ 

% .Battery was partially discharged to 80% SOC and on open circuit stand, at -5 to 

0 degrees C, during 7 Month cruise period. 

(k, Battery was charged at end of cruise through 1.2 Ohm resistor to 0.2 A cut-off 

\O 

Y and 1.95 V I cell Ave. 

Battery was heated to 15 to 20 degrees C prior to all charges. 

Battery was charged during Mars operations without resistor to a selective 

shunt limiter controlled maximum voltage. Six voltage settings were available , 

with 1.95 V I cell Ave.., the nominal full charge selection. 

3 

P k 't 

Battery was taper charged to a constant shunt limiter voltage, 1.95 V I cell 

2 Ave.. well below the 2 V. limit employed in constant current charging with same 

2 

3 

X 

full charge result. 

z 

3. 

3 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
BATTERY CAPACITY VS CHARGE VOLTAGEc 
MARS Pathfinder Power System Configuration Relays b j 

2 

Main Bus 

GSEICruise 

7 

Solar Array 

)~~spE~~ Relay 

I Battery 

Bus Return 

Battery { lso~ation 1 Relay 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
MPF BATTERY PRELAUNCH 

DISCHARGE 

Mars -- BST 18 Cell 40 Ampere-Hour, Ag-Zn, Flight Battery 

JPL Discharge #I AT 3.OA lntil 1st Cell Reaches 1.200 Volts - At 25'~ . . . . - . - . . . .. . . . .. -- - . -. - . - - . . . . . . - - - -- --- - - -- - - - -- -. - -- . - - . - - -. - -. .

- 

-. 

--.l.--l-lJ--.--l-.- 1111 

30 

Ampere-Hours 

8 

6' 

3 ELECTROCHEMICAL TECHNOLOGIES GROUP 
MPF BATTERY PRELAUNCH CHARGE 

mr -- BST 18 Cell, 40 Ampere-Hour, Ag-h, Flight Battery 

JPL Charge #2 AT 1.960 Volts Per Cell To A 0.2A Cut Off At 25'C - With 1.14 Ohm Resistor 

60 

Hours 

2 

0' 

3 ELECTROCHEMICAL TECHNOLOGIES GROUP 
- 

+ -JPL 

3 

'4 

% MPF BATTERY VOLTAGE AND 

TEMPERATURE DURING CRUISE 

AgZn Battery Performance during Day 37 to 60 of Cruise 

-Battery Current 

PRT #2 

Day of Mission 

ELECTROCHEMIGAL TECHNOLOGIES GROUP 
MPF BATTERY PERFORMANCE-EDLISOLj 

Mars Pathfinder Flight Battery - EDL and SOL 1 

0 5 10 15 20 25 30 35 

Local Solar Time, Hours 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
MPF BATTERY PERFORMANCE- SOL25 

g 

a Flight Battery Sol 25 data (D 

2 

4 - 

3 . -. - -- - -. - 

-.- 

1 . -- - -- - 

- 27 

-- 24 

.- 21 

m 

8 - 18 Q2 

;;;' 15 3 

u 

--- . 

.. . . 

-. 

-2 - , 1 

$ -- 12 5 

(D 

0 

9 E. 

E 

1A 

- 6 

3 

- 0 

12 18 0 6 12 18 0 

SCET, UTC 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
MPF BATTERY PERFORMANCE SOL-25 

SOL #25 CHARGE 

.- ELECTROCHEMICAL TECHNOLOGIES GROUP 
MPF BATTERY PERFORMANCE-SOL 25 

SOL #25 DISCHARGE 

TIME 

5 ELECTROCHEMCAL TECHNOLOGIES GROUP 
- 

4 

'0 

-JPL 

V, 

-4 

2 L 

h 

MPF BATTERY PERFORMANCE-SOL 68 

SOL 68 DISCHARGE 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
SUMMARY AND CONCLUSIONS A + 1 

3 1 The first use of a silver-zinc battery in a spacecraft application that 

(D called for extensive rechargeable operation after 12 months of active 2 

4 

9 stand hasproven to be very successful. 

c 

B BST has developed a silver-zinc battery with unique design features 

4 for the Mars Pathfinder mission. 

. JPL has developed battery management strategies to meet the Mars 

b Pathfinder mission requirements. 0 

?=' - Partial SOC at low temperature and open circuit stand was found to be 

the most effective method for insuring extensive cycle life following a 

long period of active storage 

- Silver-zinc battery charging at a reduced constant voltage was shown 

to provide full charge and a capability of supporting a large number of 

cycles. I, I 

- Launch of an inverted silver-zinc battery was shown to be possible $ 

E when a leak free cell vent valve is employed. 

'-i Q 

a - The use of a silver-zinc battery and shunt limiter in a direct energy 

a 

9% 

3 

transfer power system resulted in a very energy efficient design 

2 approach. z 

8' 

3 ELECTROCHEMICAL TECHNOLOGIES GROUP 
ACKNOWLEDGMENT 

The work described here was carried out at the Jet Propulsion Laboratory, 

California Institute of Technology, under contract with the 

National Aeronautics and Space Administration. 

1 

6' 

3 

ELECTROCHEMICAL TECHNOLOGIES GROUP 
Page intentionally left blank 
CYCLE LIFE EXPECTATIONS 

FOR 

IPV NICKEL-HYDROGEN CELLS 

Lawrence H. Thaller 

The Aerospace Corporation 

Presented at 

The NASA Aerospace Battery Workshop I 

3 Huntsville, AL 

kf 

G v e 18-20 November, 1997 d 

O"m 3 Electronics Technology Center 

2 Energy Technology Department CORPORATION 

THE AEROSPACE 

z 

6' 3 
OUTLINE 

Reasons For The Study 

Sources Of The Information Used In Study 

Types Of Information Used In Study 

Types Of Models Available To Interpret Data 

Focus Of The Study 

Initial Findings 

Classification Of Life Limiting Factors 

Assessing Significance Of Degradation Modes 

General Statements Concerning Expected Cycle Life 

Electronics Technology Center 

Energy Technology Department 
REASONS FOR THE STUDY 

IPV Nickel-Hydrogen Technology About 25 Years Old 

Several Significant Data Bases Have Been Developed 

Some Programs lnterested In Longer Life Missions 

Some Programs lnterested In Higher Energy Density 

Operation Of Nickel Electrode Better Understood 

DPA Studies Have Yielded Much Useful Information 

Results To Date Suggest Higher Useable Energy Densities 

Should Be Available Within Certain Restrictions 

A Valid Question Would Be - How Long Should IPV Nickel 

Hydrogen Cells And Batteries Last? 1, 

Electronics Technology Center rn THE AEROSPACE 

Energy Technology Department CORPORATION 
DATA BASE OF TESTS AT 60% DOD 

Sources - AF, MM, NASA LeRC, NASA SSA 

Cyles to Failure in 1000's 

Electronics Technology Center 

Energy Technology Department a THE AEROSPACE 

CORPORATION 
SOURCES OF INFORMATION USED IN STUDY 

Cycling Studies Conducted At LMA Denver Supported By 

Several Customers 

Cycling Studies Conducted At The Navy Facility At Crane 

Supported By Several Different Air Force Elements 

Cycling Studies Conducted At the Navy Facility At Crane 

. And At The NASA Lewis Research Center Supported By 

The NASA Space Station Activity 

Cycling Studies Conducted At Several Locals Supported 

By The Battery Group At NASA Lewis 

The Open Literature and Private Communications 

Addressing Pertinent Issues Affecting Cycle Life And 

Degradation Modes Is I 

Electronics Technology Center la THE AEROSPACE 

Energy Technology Department CORPORATION 
TYPES OF INFORMATION USED 

Capacity Retention As Cells Are Cycled 

Pressure Rise As Cells Are Cycled 

End Of Charge Voltage As Cells Are Cycled 

Temperature Effects On Cycle Life 

Electrolyte Concentration Effects On Cycle Life 

Cell Design Factors And Their Effect On Cycle Life 

Electrolyte Fill Amounts And Their Effect On Cycle Life 

A Review Of Manufacturing Difficulties That Are Common 

To This Technology 

The DPA Results Following Premature ,,Cell Failure , 

The Static And Dynamic Modeling Work At Aerospace 

Electronics Technology Center THE AEROSPACE 

Energy Technology Department CORPORATION 
MODELS AVAILABLE TO HELP 

INTERPRETATION OF DATA L 

Cycle Life Versus Depth Of Discharge Model 

- Suggests the relationship between cycle life and depth of 

discharge when the degradation mode is due to a slow 

wearout of the cell 

Static IPV Cell Model 

- Follows the affect of plate expansion, plaque corrosion, and 

gamma phase formation on the electrolyte requirements of 

the particular cell design 

Dynamic IPV Cell Model 

- Follows the mass transport requirements of a cell design as a 

function of electrolyte concentration and state of charge 

Electronics Technology Center THE AEROSPACE 

Energy Technology Department CQRPORATION 
I MODEL PREDICTIONS: I 

DEPTH OF DISCHARGE, 0, % 

Electronics Tecl~nology Center 

Energy Technology Department 

THE AEROSPACE 

CORPORATION 
FOCUS OF THE STUDY 

Information That Could Be Gleaned From Cell Tests That 

Cycled A Long Time 

lnformation That Could Be Gleaned From Cell Tests That 

Did Not Cycle For A Long Time 

Cell Tests That Were Conducted At High Stress Levels 

- Deep depths of discharge 

- High recharge ratios 

- High temperatures 

- High end of charge voltages 

Electronics Technology Center m THE AEROSPACE 

Energy Technology Department CORPORATION 
INITIAL FINDINGS 

ii 

s In Reviewing The 60% DOD Data, The Cycle Life Was 

g, 

3 Found To Span The Range Of Less Than 1000 Cycles To 

F 

$ Over 50,000 Cycle and Still Cycling 

s. 

4 The Wearout Model Could Not Be Used, One Had To 

Assume Inappropriate Cell Designs, Manufacturing, Or 

Charge Management Problems 

A. 

N 

? 

Cycling Problems Were Divided Between: 

- Cell design related issues 

- M'anufacturing related issues 

- Charge management related issues 

Electronics Technology Center a THE AEROSPACE 

Energy Technology Department CORPORATION 
CELL DESIGN ISSUES 

1997 IECEC Paper By Thaller Deals With This Issue 

2 

d - Cell Designs must take into account the changes taking place 7 

i?Y within a cell as it is cycled 

5 

4 - In particular, changes that impact the electrolyte requirements 

are very important 

- These are plate expansion, plaque corrosion, gamma phase 

b 

h, 

formation, and temperature gradients 

Electronics Technology Center 

Energy Technology Department 

THE AEROSPACE 

CORPORATION 
4 

'0 

'0 

'4 

s 

k2 

L 

2 

MANUFACTURING ISSUES 

U 

B Most Of The Manufacturing Related Issues Are Associated !$' 

d With The Nickel Electrode 

- Sinter pore size and pore size distribution 

- Active material loading levels and distribution within the 

plaque 

- Activation procedures, storage procedures, and contaminants 

B 

tu 

Y Most Of These Issues Are Reported In The Open Literature And Can Be Kept Under Control 

d 

'% 

3 

3 Electronics Technology Center H THE AEROSPACE 

8 Energy Technology Department s 

CORPORATlON D 
RECENT FINDINGS REPORTED BY CRANE 

Cells cycled at 10°C always cycled longer than equivalent cells cycled at 

-5OC 

e 

F 2 A recharge protocol that included a taper charge as the cells approached 5 

+s full charge always resulted in longer cycle lives compared to equivalent 

cells that were charged at a constant current to the same charge return 

ratio. 

Cells that were filled with 26% potassium hydroxide as the electrolyte I 

8 

Y always cycled longer than equivalent cells filled with 3 1 % potassium 

hydroxide. 

The cause of failure of cells cycled to 60% DOD in all cases where 

failed cells were disassembled was always attributed to degradation of 

the nickel electrodes. 

5 1997 IECEC Papers presented at the Hawaii meeting 

ff 

s Electronics Technology Center 

Fa 

M Energy Technology Department 6' 3 
INFORMATION GLEANED FROM DATA BASES 

Of the 23 packs where a failure was reported in the Air 

Force and NASA supported testing at the Navy facility at 

Crane, 22 were associated with either recharge ratios that 

were 1.05 or higher, or the end-of-charge voltage was 

above 1.55 Volts. 

Of the 21 failures reported in the testing carried out at 

NASA Lewis under Space Station sponsorship, 19 

occurred in tests where the recharge ratio was 1.05 or 

above, while for the 18 tests where there has been no 

failure, the recharge ratio has been 1.04 or below. The 

tests that are still running have accumulated between 

44,000 and 50,000 LEO cycles at 35% DOD. I 

Electronics Technology Center 

Energy Technology Department 

THE AEROSPACE 

CORPORATION 
4 

\O 'a 

'4 

5 

k? b CHARGE MANAGEMENT ISSUES m t 

a 2 2 

bo 

a, 

Charge Management Was Found To Play An Important 

3 Part In The Cycle Life Of Test Carried Out At 60% DOD 

- High end of charge voltages can result in high rates of plaque 

corrosion and excessive oxygen formation within the pores of 

the nickel electrode. 

- NASA Lewis is testing cells with plaque material that has 

undergone a passivation process 

t3 

cf -- Cycling to date has indicated a significant reduction in the 

corrosion rate estimated by the pressure rise with cycling 

Electronics Technology Center 

Energy Technology Department 

THE AEROSPACE 

CORPORATION 
$ 

U 

2 b 

52 A 

SIGNIFICANCE OF DEGRADATION MODES 

2 1 

.n" 

3 

m 

5 Plate Expansion c 

d 

P 

- NASA supported work at Hughes (1980s) reports up to 80% 

B z plate expansion 

% - Air Force funded testing on Pack 5402H showed an average 

of 6% plate expansion when DPAed after 42,000 40% DOD 

cycles (cell was removed from test prior to failure) 

b 

h, Plaque Corrosion 

?' 

- NASA supported work at Hughes reported up to 50% 

corrosion of the nickel after about 40,000 at 80% DOD in an 

accelerated 45 minute LEO cycling test 

- Cell from Air Force Pack 5402H showed an average corrosion 

of about 22% when analyzed during the DPA studies I 

$ 

E k V 

DPA studies on Air Force cell were carried out at Aerospace 

% by G.To and M. Quinzio under the direction of Dr. Zimmerman 

OQm 3 

tl Electronics Technology Center 2 g. Energy Technology Department CORPORATION THE AEROSPACE 

3 
CONCLUDING REMARKS 

IPV Nickel-Hydrogen Cells Have The Potential For 

Extremely Long Cycle Lives 

- Current champion in terms of DOD and cycle life are cells 

manufactured by Eagle Picher under the direction of the 

design team at NASA Lewis 

- 50,000 LEO cycles at 60% DOD 

Deeper Depths Of Discharge Are Available If The Four 

Main Factors Resulting In Abbreviated Cycling Are Properly 

Addressed 

- Proper Cell Design 

- Proper Nickel Electrode Manufacture 

- Proper Charge Control i , 

I 

- Proper Thermal Environment 

Electronics Technology Center TEnergy Technology Department CORPORATION 
Page intentionally left blank 
h Technologies Division 

Life Cycle Testing 

Spaceflight Qualified 

Nickel-Hydrogen Battery Cells 

Chris Fox, Jennifer Francisco, Kevin Ames, Ron Repplinger 

Power Subsystems ~epadrnent I 

Joplin, MO 64802 

(41 7) 623-8000 

http:Ilwww.epi-tech.com 

, NASA Aerospace Battery Workshop, November 18 - 20,1997 I 
Cell Design Variable 3 

.a 8 

3 + Plaque Fl 

q k Thickness 2 

@ P Porosity 

k Type 

Dry Sinter vs Slurry 

b + Electrolyte (KOH) Concentration 

W 

e h 31% vs 26% 

+ Separator Type 

k Asbestos vs ZIRCARB vs 2 ZIRCAROD 

+ Electrolyte Management 

b Wal! Wick vs No Wall Wick vs Catalyzed Wall Wick vs PTFE , 

L 

2 + Cell Configuration 

B 2- COMSAT Design vs Mantech (Air Force) k '-4 R a 

2: IPV, CPV, DPV, SPV (5" and 10" diameter, batteries only) 

% 

z P 2.5", 3.5", 4.5", 5.5" diameter 

e g . 

z 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
m , IN-HOUSE NiH, LEO LIFE TEST SUMMARY 

3 

EAGLE @PICHER 

Tec hnoloaies Division 

-2 

m 

bcr 

s 

9 

F Cell 

B Z RNH 50-15 

3 RNH-80 

RNH-50-15 

RNH-76-3 

b 

W 

C 

RNH-76-9 I RNH 30-1 

RNH 50-15 

RNH-76-3 

RNH-76-3 

RNH-76-3 

3 

RNHC 10-1 

E RNHD 40-1 k Y 9 RNHD 60-1 

Cell 

DE 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

CPV 

DPV 

DPV 

# of 

Cells 

19 

1 

1 

5 

1 

10 

7 

7 

10 

4 

I I 

3 

3 

Fa g . 

a RT = Real Time ACC = Accelerated 

Re~ime 

ACC LEO 

ACC LEO 

ACC LEO 

ACC LEO 

ACC LEO 

RT LEO 

RT LEO 

RT LEO 

ACC LEO 

RT LEO 

RT LEO 

RT LEO 

RT LEO 

DOD 

15% 

15% 

# of 

Cycles 

127,770 

1 17,908 

112,903 

108,015 

96,866 

69,806 

56,880 

55,879 

50,170 

35,179 

17,897 

8,004 

5,225 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
: IN-HOUSE NiH, LEO LIFE TEST SUMMARY 2 

Cell 

RNH-56-1 

RNH-56-1 

RNH-56-1 

RNH-56-1 

RNH-56-1 

RNH-56-1 

RNH-56-1 

Cell 

IYm 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

I PV 

# of 

Cells 

3 

3 

1 

4 

3 

4 

3 

Regime 

RT LEO 

RT LEO 

RT LEO 

RT LEO 

RT LEO 

RT LEO 

RT LEO 

# of 

DOD Cycles 

70 *I 5,100 

70 *15,100 

70 *15,100 

70 *15,100 

70 *I 2,900 

70 *5,200 

70 *800 

Cell # of # of 

I 

Battery IUE Cells Regime ' DOD Cycles $ 

9 SAR-10013 I PV 27 RT LEO 12% 31,443 k k 

Q a RT = Real Time ACC = Accelerated * = RLF Plaque B 

NASA Aerospace Battery Workshop, November 78 - 20, 7997 
"IN-HOUSE b NiH, GEO LIFE TEST SUMMARY 

2 

EAGLE E$ MCHER 

Technoloaies ~ivision 

% 

d 

P 

Cell 

Cell # of # of 

DIE Cells Reaime DOD Seasons 

2. 

s 

4 

RNH 120-H I PV 4 ACC GEO 75% 44 

RNH-65-15 I PV 6 ACC GEO 56% 42 

RNH-120-1 I PV 6 ACC GEO 100% 23 

b 

W r RNH-76-11 I PV 6 ACC GEO 57% 21 

I RNH-100-7 I PV 6 ACC GEO 57% 20 

I RNH 65-33 I PV 4 ACC GEO 70% 17 

Cell # of +. # of 

5 Battery QE! Cells Regime 

iT 

'4 k SAR-I0017 * I PV 27 RT GEO 

DOD 

75% 

SAR-10017 I PV 27 ACC GEO 75% 

$ RT = Real Time ACC = Accelerated 5. * = Complete 

3 

Cycles 

NASA Aerospace Battery Workshop, November 78 - 20,7997 
G WLE L~PKHER 

OFF-SITE NIH, LIFE TEST SUMMARY F Technologies ~ivision 

a 

t3 

:: 

Q 

$ 

d Cell Tvoe 

% 

q- ' RNH 30-1 

!z 

4 RNH 100-1 

RNH 90-3 

RNH 90-3 

RNH 90-3 

b 

W 

f 

RNH 90-3 

RNH 90-3 

EP-26 

RNH 90-13 

RNH 48-1 

RNH 48-1 

RNH 65-15 

$ 

i3 

RNH 100-7 

& rl RNH-50-59 

2 

Z 

3 

' RNH 98-1 

tn 8. RNH 98-3 

3 

# Cells 

14 

8 

4 

DE 

LEO 

LEO 

LEO 

LEO 

LEO 

LEO 

LEO 

LEO 

LEO 

LEO 

LEO 

GEO 

GEO 

GEO 

GEO 

GEO 

Cell 

% DO9 

13% 

40% 

7% 

7% 

7% 

7% 

7% 

60% 

30% 

35% 

45% 

55% 

65% 

65% 

65% 

65% 

# of 

Location Cycles 

MSFC 53,200 

LMC 44,488 

MSFC 42,700 

MSFC 41,600 

MSFC 41,500 

MSFC 40,900 

MSFC 40,860 

LMC 33,019 

MSFC 27,100 

MSFC 4,700 , 

Is MSFC 4,700 

TRW 42 Seasois 

TRW 20 Seasons 

LMC 19 Seasons 

TRW 02 Seasons 

TRW 02 Seasons 

NASA Aerospace Battery Workshop, November 78 - 20, 7997 
I EAGLE L~PKHER " i TERMINATED NiH, LIFE TESTS Technologies ~iv,ision a I 

Cell 

% DOD 

13% 

# of 

Cycles , 

49,815 

Cell Type 

RNH 35-3 

# Cells 

14 

TviDe 

LEO 

Location 

MSFC 

LEO LMC 

LEO LMC 

LEO LMC 

LEO MSFC 

LEO LMC 

LEO EPI 

LEO \ 

EPI 

LEO EPI 

LEO LMC 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
cl 

U 

s 

ON-ORBIT NiH, IPV CELL HERITAGE i? Technologies Division 

CELL/ PROGRAM TOTAL d 

Y BATTERY NAME CYCLES ;-" - DOD 

0, 

USAF 

NTS2 

IV FM6 

IV FM7 

IV FM8 

SPACENET FM1 

SPACENET FM2 

IV FMIO 

G STAR FMI 

IV FMI 1 

ASC FMI 

IV FM12 

SATCOM FMI 

SATCOM FM2 

G STAR FM2 

SPACENET FM4 

IV FM13 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
s k2 EAOLE @PICHER 

Technologies Division 

k 

a 

2 2 CELL/ 

BATTERY 

RNH-35-1 

RNH-30-1 

RNH-50-15 

RNH-30-1 

RNH-30-3 

RNH-30-3 

RNH-50-15 

RNH-50-27 

RNH-40-7 

RNH-40-7 

RNH-50-27 

RNH-78-1 

RNH-78-1 

RNH-50-15 

PROGRAM 

NAME 

PAN AM SAT FM#I 

G STAR FM3 

ASTRA 1A 

IV FM15 

G STAR FM4 

ITALSAT 

ASTRA I B 

ANIK E F2 

ASC-2 

AURORA-2 

ANIK E F1 

TELECOM 2-A 

TELECOM 2-8 

INTELSAT K 

TOTAL 

CYCLES 

NASA Aerospace Battery Workshop, November 18 - 20,1997 I 

. . 
4 

\O 

3 

s 

ON-ORBIT NiH, IPV CELL HERITAGE EAOLE @PICHER 

b 

3 Technologies Division 

g 

3 CELL1 PROGRAM TOTAL 

8 BATTERY 

3' - NAME CYCLES - DOD 

SATCOM C4 

SATCOM C3 

HISPASAT IA 

HISPASAT 1 B 

TELSTAR 401 (a) 

TELSTAR 401 (b) 

MILSTAR DFSI 

BS3N 

MILITARY (Classified) 

KOREASAT - 1 

TELSTAR 402R '1 92 

MILSTAR DFS2 182 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
ON-ORBIT NiH, IPV CELL HERITAGE - 

CELL/ PROGRAM 

BATTERY 

TOTAL 

- NAME CYCLES 

KOREASAT - 2 

INMARSAT- 3 Fl 

INMARSAT- 3 F3 

TELECOM 2C 

TELECOM 2D 

INMARSAT- 3 F2 

GE- 1 

G E-2 

TEMPO-2 

TELSTAR 501 

INMARSAT- 3 F4 

GPS-IIR 

GE-3 

ECHOSTAR-3 

APSTAR 2R 

GPS-2-28 

EAGLE @PKHER 

Technologies ~iv,ision 

- DOD 

70% 

70% 

70% 

80% 

80% , 

70% 

80% 

70% 

70% 

100% 

70% 

60%l I 

70% 

70% 

100% 

60% 

4 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
5 i EAGLE L? PICHER " ! ON-ORBIT NiH, IPV BATTERY HERITAGE Technologies Division 

4 

2 

e 

F B 

BATTERY 

B 

4 SAR 10007 

(RN H-35-3) 

SAR 1001 I 

(RNH-30-3) 

b 

P 

? SAR-10015 

(RN H-90-3) 

SAR 10009 

(RNH-65-1-3) 

SAR 10009 

(RNH-65-1-3) 

3 0' SAR 10009 

5 

& Y 

(RNH-65-1-3) 

a a 

2 

3 

SAR 10009 

GI 

g 8 . (RNH-65-1-3) 

a 

PROGRAM 

NAME 

OLYMPUS 

HST 

EUTELSAT FMI 

EUTELSAT FM2 

EUTELSAT FM3 

EUTELSAT FM4 

TOTAL 

CYCLES - DOD 

60% 

60% 

9% 

80% 

80%: 

80% 

80% 

NASA Aerospace Battery Workshop, November 18 - 20,1997 
52 

UN-ORBIT NiH, IPV BATTERY HERITAGE WLE L~PKHER 

L. 

n Technologies Division 3 - 

% 

R 

(D 

V 

F 0 

BATTERY 

X 

PROGRAM 

- NAME 

TOTAL 

CYCLES 

u 

s. 

4 SAR-I 0009 (RNH-65-1) TURKSAT I B 294 

SAR-10017 

(RNH-78-1) ORION F1 

b 

266 

P 

w I SAR 10009 

(RNH-65-1-3) EUTELSAT FM6 22 1 

SAR-1004 I 

(RNH-45-1) ASIASAT 2 177 

SAR-10073 

(RNH-50-57) ECHOSTAR 101 ' 169 

3 - SAR-10057 & 

& (RN H-65-33) , 

cl 

B 

- DOD 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
"ON-ORBIT 9 NiH, IPV BATTERY HERITAGE 

LCRlLL yrmr Technolonies Divi 

!$ 

d CELL1 

=T s 

BATTERY 

!z 

3 SAR-I 0073 

(RNH-50-57) 

PROGRAM TOTAL 

NAME CYCLES 

ECHOSTAR 102 106 

SAR-10057 

, 1 (RN H-65-33) NAHUEL IA 

(RNH-45-3) INTELSAT 801 

I SAR-10041 

(RNH-45-3) INTELSAT 802 

SAR-I 0041 

(RNH-45-3) INTELSAT 803 

- DOD 

70% 

70% 

70% 

70% 

70% : 

, NASA Aerospace Battery Workshop, November 78 - 20, 7997 
EAGLE L9 PICHER E 

b 

' ON-ORBIT NiH2 CPV BATTERY & CELL HERITAGE Technologies Diiision (D 

a G 

9 

CELL/ 

BATTERY 

RNHC-6-1 

RNHC-10-1 

RNHC-I 0-1 

i 

RNHC-10-1 

: IRNHC-10-1 

PROGRAM 

TUBSAT-B (CPV) 

ORBCOMM F1 (CPV) 

ORBCOMM F2 (CPV) 

MICROLAB F1 (CPV) 

MSTI-3 (CPV) 

MSTI-2 (CPV) 

APEX (CPV) 

MGS (CPV) 

TOTAL 

CYCLES 

NASA Aerospace Battery Workshop, November I8 - 20, 1997 
EAQLE E+ PICHER 

Technologies ~ilision 

- DOD 

50% 

31 % 

31 % 

31 % 

31 % 

31 % 

31 % 

31 % 

31 % 

31 % 

31 % 

31% , 

31 %: 

31 % 

31 % 

31 % 

31 % 

31 % 

NASA Aerospace Battery Workshop, November 18 - 20,1997 
2 b ' P ON-ORBIT NiH, SPV BATTERY HERITAGE ~echnologhs Division 

Q 2 CELL/ 

2 BATTERY 3 

F 

9 SAR-10065 

E SAR-10065 

42 SAR-10065 

SAR-1008 1 

SAR- 1 008 I 

SAR- 1 008 1 

% SAR-10081 Y' SAR-I 008"1 

SAR-I 0081 

SAR-I 008 1 

SAR-10081 

SAR- 1 008 1 

SAR-10081 

SAR-10081 

SAR- 1 008 1 f 

& SAR-1008 1 

V 

a 

'2 

SAR-10081 3 

2 SAR-1008 1 

PROGRAM 

NAME 

IRIDIUM 18 (SPV) 

IRIDIUM 19 (SPV) 

IRIDIUM 20 (SPV) 

IRIDIUM 21 (SPV) 

IRIDIUM 22 (SPV) 

IRIDIUM 23 (SPV) 

IRIDIUM 24 (SPV) 

IRIDIUM 25 (SPV) 

IRIDIUM 26 (SPV) 

IRIDIUM 27 (SPV) 

IRIDIUM 28 (SPV) 

IRIDIUM 29 (SPV) 

IRIDIUM 30 (SPV) 

IRIDIUM 31 (SPV) 

IRIDIUM 32 (SPV) 

IRIDIUM 33 (SPV) 

IRIDIUM 34 (SPV) 

IRIDIUM 35 (SPV) 

TOTAL 

CYCLES - DOD 

- 

, NASA Aerospace Battery Workshop, November 18 - 20,1997 
B ON-ORBIT NiH, SPV BATTERY HERITAGE 

CELL1 

BATTERY 

SAR-I 0081 

SAR-10081 

SAR-I 0081 

SAR-10081 

PROGRAM 

NAME 

IRIDIUM 36 (SPV) 

IRIDIUM 37 (SPV) 

IRIDIUM 38 (SPV) 

IRIDIUM 39 (SPV) 

TOTAL 

CYCLES 

82 

82 

82 

82 

EAOLE QPICHER 

Technoloaies division 

- DOD 

31 % 

31 % 

31 % 

31 % 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
WLE @PICHER 

Technologies division 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
EAQLE I+PKHER 

Technologies Division 

+ 

Y, 

3 

5 

2 RNHC- f0- 7 End-of-Discharge- Voltage (EOD V) 

2 

a 

2 

1 

'a .---.---.------------.------..."-,"---..--.---.-.-----"-.------.---.---."."--".-- 

9. 

F, - - - - - - 2- * .?\ 

r? --Uh.p-h 1.- - . . . . . . . - .',b'.=;*.';C- - - . . . . " + - . . $ h.rL.t.*'3L c=d,~= >3& - = - - = -. . *-.I.,.*. 

3 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 9 I 
RNHD-40- 1 LIFE TEST EAGLE @PICHER 

2 Technologies Qivision 

i? 

a G 

R 

m 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
RNHD-60- 7 LIFE TEST 

I 

WLE @PICHER 

Technologies division 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
FULL LENGTH SPV LIFE TESTS 

WLE QPICHER 

Technologies division 

+ 100 minute, 31% DOD LEO cycles 

+ Charge and discharge waveforms are pulsed 

+ Three on-going tests 

Z 14,111 orbits at 10°C 

Z 5,744 orbits at -5°C 

k 2,320 orbits at 5°C 

3 

h 

2 NASA Aerospace Battery Workshop, November 18 - 20, 1997 
31% DOD SPV LIFE TEST 

WLE QPICHER 

Technologies Division 

EODV for EM Life Test 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 

32 .F ............................................. ................ -- ..... - .... -- .. .--.... .... - .. .................... - .... --- 

3 1 ..- -- ...- .- .....-... ....... . 

a, 30 

0 I 29 

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.. - -. .. .. .- .... -- - - . . . 

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0 2000 4000 6000 8000 10000 12000 14000 

Orbit Number 

..... - - . -- ....... . - ............ - ..... . -- -- - -- .... .- - - - -. -- 
", 

'\I 

2 '4. 

2 ACCELERA TED SPV LIFE TEST 

EAGLE I+PICHER 

Technologies division 

R 

'73 

2 

z + 10°C test 

9 

7 P + 100 minute LEO cycles E 

.a0 + 70% DOD 

+ Increased charge rate as necessary to maintain 22V minimum 

+ Completed requirement of 4,612 total cycles 

+ Continued with low rate (31% DOD) 100-minute LEO cycles 

+ Completed an additional 4,812 cycles 

+ DPA revealed cells in excellent condition 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 wmp 
-;1 

5 EAGLE L~P~CHE

e 

b 

HIGH RATE CYCLES - ACCELERATED LIFE TEST Technologies Division 

2 'I 

% 

9 

Graph 1. High Rate EODVvs Orbit Number 

Restart after 1 month open circuit.. 

Restart after 1 day open circuit. 

Charge 

Orbit Number 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
WLE L~PKHER LOW RATE CYCLES - ACCELERATED LIFE TEST Technologies 'Division '1 

Q) 

FZ 

CO = 

r 0 

0 > to- .I 

Graph 2. Low Rate EODV vs Orbit Number 

Restart afier High Rate Test. 

I 

Orbit Number 

NASA Aerospace Battery Workshop, November 18 - 20, 1997 
'r 

'Q 

'0 

U 

s Conclusions 

EAGLE @PICHER 

2 Technologies Division 

+ Total Spacecraft On-Orbit 

+ Total Cells On-Orbit 

+ Total Operational Cell Hours On-Orbit 209,089,724 

b 

a + Longest Running ACC LEO 127,770 + Cycles ?' 

+ Longest Running RT LEO 69,806 + Cycles L 

l + Longest Running ACC GEO 44 + Seasons 

+ Longest Running RT GEO It 42 + Seasons 

$ 

B k w B + Longnest Running On-Orbit LEO 41,410 + Cycles 

% 

2 

@ + Longest Running On-Orbit GEO 

8 

29 + Cycles a. 

5 NASA Aerospace Battery Workshop, November 18 - 20, 1997 , - 
# ,@ ),' COMSAT LABORATORIES 

- 

EOS-A MI Nickel Hydrogen Cell Interim Life Test Report 

- - 

C. W. Bennett 

Lockheed Martin Missles and Space 

D. J. Keys, G.M. Rao 

NASA/GSFC 

H. E. Wannemacher 

Jackson and Tull 

H. Vaidyanathan 

COMSAT Laboratories 

(Performed under NASA Contracts NAS5-32500 and NA S5-32600) 

Abstract LOCKMEED M.;.T$ 

'fiiis papr rcpons the interim results of the Eanh Observing System AM-I project (EOS-AM-I) nickel hytlrogm cell life 

test king conducted undercontrdct to National Aeronautics md Space Adminisvation (NASA)Goddard Space Fli, -ht Centcr 

tGSFC)ilt the Lockheed Mmin Missles and Space (LMMS) facility in East Windsor. NJ: andat COMSAT Labs.. Ckuksh~rg, 

MD. The purpose of the tests is to verify that the EOS-AM-I cell design can meet five yem ofreal-time Low Eanh Orbit 

(LEO) cycling. The tests include both real-time LEO and acceiera~ed stress tests. At LMMS. the first real-time LEO 

simulated 99 minute orbiml cycle sraned on February 7. 1994and the test has been mnning continuously since that time. with 

I R202 LEO cyclescompletrd as of Seplembcr I. 1997. E;sh cycle consisrs of a 64 niinute chaqe (VTat 1.507 volts per cell. 

1.06 UD olio. followed by 0.6 anipere trickle chaqe) and a 35 minutt consmt power discharge at 177 wtts (22.5% DOD). 

A1 COMSAT. the :~ccelented stress test consists of 90 minute orbiml cycles at 60%. DOD with a 30 minute disch:~rgc at 60 

amperes :md r! 60 minute charge at 50 amperes (\Tat 1.54 volts per cell to 1.W C'D ntio. followed by 0.6 ampere trickle 

cirarge). 

'The real-time LEO life test battery consists of seven. 50AH (nameplate mting) Eagle-Picher. Inc. (EPI) .Mantech cells 

m;mufactured into three. :well pack assemblies !there are two place holder cells that are not pan of the life test eleclric:~l 

circuit). The test pack is configured to simulate the conductive thermal design of thespacecraft batterq: including: contluctivc 

aluminuni sleeves. 3-cell pack aluminum baseplate. and honeycomb panel all mounted to a liquid (-5OC) cold plilte. Th: 

entire assembly is located in a thrrmal chamber opentin: at +3"C. The accelerated stress rest unit consists of five cells 

rnounrcd in machined aluminum tcst sleeves and is operating at +I(%. 

Tne real-time LEO life test barter?; has met all performance requirements through the first 18.202 cycles, including: cntl of 

chage and dischilrge cell voluges and voltage gradients: endof ch~e and discharge cell pressures; within cell and betr\ten 

cell temperature gradients: discha~capacity: current and power levels: and all charge parameters. The accslerated strer* 

tcst battery has completed 11998 cycles when the test was terminaad. The svcss test unit met all lest parameters. This pnper 

repons battery performances as a function of cycle life for both the real-time LEO and rhe, accelerated life tes; rcpimes. 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
f - . -5.- 

Battery AssemblylSpacecraft Interface LOCK~ECD MARTIN J' 

Power Equipment Module Battery (PBAT) 

HighGain 

Anlenna Veloc~ty 

/ Y -.- 

Housekeeptng 

Equipment 

Panel Module 

Hex Bay Battery (BBAT) 

Outline LOCKWEED MAR 4 

TIN .f - 

EOS-AM cell parameters 

Real-time LEO test conditions 

Real-time LEO test results 

- Figures 1-9 

Accelerated LEO stress test conditions 

Accelerated LEO stress test results 

-Figures 10-18 

Summary and conclusions 

I997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
EOS-AM Cell Parameters 

Single stack, IPV, Mantech design 

Rabbit ear terminals - - 40, 30 mil slurry electrodes 

Back to back with double layer Zircar, catalyzed wall wick 

31% KOH, nickel precharge 

Average cell weight = 1490 grams 

Average delivered capacity (C12 discharge) 

-1 O°C: 75 AH 

O°C: 71 AH 

+1 O°C: 63 AH 

+20°C: 56 AH 

+30°C: 51 AH 

+lO°C, 72 hour charge retention: 92% lot ave, 89% minimum 

- 4 

Real-time LEO Test Conditions LOCKnEED MARTIN I-'. 

Performed at LMMS, East Windsor, NJ 

Number of cells: 7 

Cell configuration: 

-Conductive thermal design with same configuration as spacecraft 

- Machined aluminum sleeves and baseplate 

- Chotherm and RTV 566 isolation 

-Mounted to spacecraft honeycomb panel with Al face sheets 

- Liquid cooled (-5OC) cold plate 

Discharge regime: 177 watts (total) constant power discharge for 35 minutes 

Depth of discharge: 22.5% nominal 

Charge regime: 

- 12.3 ampereeto 1.507 V/T (per cell), taper to 1.06 CID, 0.6 A trickle 

- 64 minute total charge time 

1997 NASA Aerospace Bartely Workshop Nickel-Hydrogen Session 
LMMS Test History 

Fig 1- 

End of Charge and Discharge Battery Voltage vs LEO Cycle 

Imyx,nn 

Cl~rwc lmnar n-xhrd prnr lo \T limit. Srrrrzl tr4 sIvm% 

Crllsdbrchrrpd - 1 All -illtout rrrharfe. Vlurfc ~tmr Itm. rrxlnl fwttlriu 71' lln~rr 

\T rhmd from 1059 lo 10.fs26 

-I I 

LXscNrpr loxl dunrd lrm 177 rallr I,, XU rslr 

IY.wharpr lo& chrnN lmnv OR wrllr to 177 uattr 

.\l rhrnprd from 10.626 lo 105.W 

Us* prinoulr ~1 IOY~ ! 

Srvcrnl rlmn duraiw~w rl.n!r 

T-t mllr on qrn rtrnttt dt~c to ST!; pml-trnnr 1 Test mllr M qrn CIWII JUC to SlTi pn>hCms I 

T-I rrlls on opcn rirmil euc lo Sl'li anhlrnvr 

Tnr c.ll<on own cirrun~ Jur lo .Qti pmhlrm< J 

Clurp prmomchmycd: Cuma 1235A lo 15.11A. V:T 10.55 111 10.711.C' IJ 1.05 1'1 1.20 

Cbr(r pnrn~mcha~~d.Cun~~ IS.OA lo 12.2SA. V.1' 111.7Utu 10.55.C' 11 I.2Om 1.015 

Chrrp prmelnrchaqcd: Cuma 1Z.ZtA lo IJ.OA. Vr 10.55 lo 10.71l.(71~ 1.05 10 1.20 

Charge pnnmmch.n@: Cvmn I5.M ln IZZ5A. VrT 10.70 la 1035.C~I> 1.20 tu 1.115 

C'harro pmctrnrhm~nt: Cunrn 12.2SA lo IS.llA. Val' 10.55 lo IU.711. C' I> I.lP lo :.ZO I 

t 

('hxqr pnmrlrnrhmpcd: Cumn ISW lo 12.2SA. VrI' tlL7ll In 10.55. ('.I> 1.2010 1.05 

Chrrgc pannlclmchmyrd: Cvmnl I?.?% lo 11.9A. Vfr 10.55 lrr IU.4S.'liirklr ll.6ilA In ZUA. So C:Il. 111% 

rhsrgc lard 177U'lo 86.5U' 

Chgr to urwJatdorbi1: l77W dnsrhrpc. I2.25h chrrpc. VTF-I(LS5V.l.~ Ihl.O.5. Tridlr-U.Nlh 

Rnal orbslal ryck prior loshulhrn for nlovr s, lrcu brilily 

S~rial 1-1srquccc. 1DT rapant). lm.72 hour rhar,zr rnrnniana. IlrK'rnrxci~? I 

1997 NASA Aerospace Battery Workshop 

Cyrlr 

?00 

Ill0 

3x0 

3750 

3050 

3Sm 

S2oh.W 

ONXl 

72m 

RlW 

8700 

12950 

15501 

15601 

16101 

16201 

lbaol 

16201 

lPWl 

18101 

18202 

IS?* 

Evcnl 

I 

2 

3 

4 

5 

6 

7 

R 

Nickel-Hydrogen Session 

Dtlc 

Y:lrUJ 

uWU.4 

1WZllUd 

I(Y?8'94 

11:5.'94 

l1.7'U1 

2719..CS126U5 

Y 15.VS 

9 

10 

11 

I2 

13 

14 

IS 

16 

17 

18 

19 

20 

21 

22 

' u?&u~-7,?~ 

9:IIPL9:2R'95 

111109.(-11 1195 

9!liOh 

2'24'97 

31397 

~7'07 

Af15!97 

4i2R.97 

5'5:YT 

8'1&97 

6'2!197 

9:llQ7 

9'3'97-gl0 97 
Fig 2 -. -A. LOCKNEED MAUrlN ,'- 

Minimum and Maximum Cell End of Charge Voltage vs LEO Cycle 

- 

Fig 3 - ..--A 'tocKHEEo mA,trrw .A 

Minimum and Maximum Cell End of Discharge Voltage vs LEO Cycle 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Fig 4 

n -- .- LOCKHEED MARTIN !" 

End of Discharge Cell Voltage Gradient vs LEO Cycle 

Fig 5 -aLo .... -- ---f',7 ED MA..,. ., 

CID Ratio (Total Charge CapacitylTotal Discharge Capacity) vs LEO Cycle 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Fig 6 

f 

LOCWHEED MA R-i-le,- 

Cell 1 Pressure vs LEO Cycle 

Fig 7 

Individual Cell Charge Voltage for Cycle 17830 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Fig 8 

.4 -. ..-- LOCKHEED MARTIN I. 

- 

Individual Cell Discharge Voltage for Cycle 17830 

Cell 1 Temperature Profile for Cycle 11800 

1997 NASA Aerospace Battev Workshop Nickel-Hydrogen Session 
Accelerated Stress Test Conditions LABORATORIES 

; Performed at COMSAT Laboratories 

Number of cells: 5 

0 Cell configuration: 

- Machined aluminum sleeves 

- Mounted on a cold plate 

- Chotherrn and RTV 566 isolation 

Test temperature: 10°C 

Discharge regime: 

- 60 amperes for 30 minutes 

Depth of discharge: 60% 

Charge regime: 

- 40 ampere to 1.54 VTT (per cell), taper to 1.09 CID, 0.6 A trickle 

- 60 minute total charge time 

Fig 10 COMSAT LABORATORIES 

Voltage and Current Profiles for Cycle 11998 

3 55 

150.. 

145 -. -....__.._, 

'-1 ... I 

1 j .. ic 

/ .:> 

I E 

I 1 i. 

t2s{ I i i .- --CC'lAlp . 

t XI Dnco.<ge . . . . . .trs,& 

................. ...... .......... i 7: 

I IS 1'- 1 10 C I0 I?" 32 0 a 5a w C w z 

TIME [nltll] 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Fig I1 LABOAATORICT. A 

CID Ratio 

CYCLES 

Fig 12 LABOR/ITORIF< 

End of Discharge Sleeve Temperature 

5.0 

0 1000 20W 3000 4000 5000 BOO0 7000 8000 9000 10000 11000 :ZOO0 

CYCLES 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Fig 13 LABORATORIES 

End of Discharge Average Cell Voltage 

1.DOo 

0 2000 4WO 6000 8000 10000 12000 

CYCLES 

Fig 14 LABORATC)R1C? 

End of Discharge Individual Cell Voltages 

-..... C." : -..- CCI 3 

CYCLES 

1997 NASA Aerospace Balteiy Workshop Nickel-Hydrogen Session 
End of Charge Pressure 

350 0 

o IOW 2000 3000 4000 5000 6ow 7000 eooo 9000 ioooo :iooo 1:ooo 

CYCLES 

Fig 16 LABORATORIE: 

End of Discharge Pressure 

I 

0 1000 2000 3000 4000 5000 6000 7000 BOO0 BOO0 10000 1?000 :ZOO0 

CYCLES 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Sfi3.43 

0 OOOI OWL OOOC OOOV 0005 0009 OGOL 0008 0006 00OOL OOOLL 000Ll 

0'0 

$ an u 

s313i.3 

0 0001 0002 000C OOOI 0004 0008 OOOL 0008 0006 OOOOL OOOLL OCOLL 

""" ''_'"*' "' :.._'.''' . :. . I. 
Summary and Conclusions - 

Real-time LEO life test at Lockheed Martin Missles and Space 

- Cells have completed 18203 cycles as of 9/1/97 

- End of discharge voltage decreased initially and has stabilized 

-Variation in end of discharge cell voltage has converged with cycling 

-Cell EODP and EOCP stabilized after 8000 cycles 

-- 1.51 VTT limit and 1.06 C/D insufficie.nt to makeup for early test 

problems (cells on open circuit during all test outages) 

-There was no full reconditioning performed during the first 18203 cycles 

Accelerated LEO stress test at COMSAT Labs. 

- Cells have completed 11 998 cycles 

- End of discharge voltage decreased initially and has stabilized 

-Charge termhation at 1.55 V and 1.08 C/D is appropriate for 60% DOD 

cycling 

- EOC and EOD Pressure have increased with cycling 

- Two reconditioning cycles performed 

- - Temporary voltage improvement for approximately 100 cycles 

1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
Characterization and Initial Life-Test Data 

- - for Computer Designed Nickel Hydrogen Cells 

A. Zimmerrnan, J. Matsumoto, A Prater, 0. Smith, and-N. Weber 

Electronics Technology Center 

The Aerospace Corporation 

El Segundo, California 90245 

Abstract 

Five advanced nickel hydrogen battery cells have been acquired based on specifications 

developed from a first principles battery performance model, with the objective of reducing internal cell 

impedance by SO%, while improving cycle life and performance. The initial performance of these cells 

has been characterized and is reported here, and has been found to match the performance predicted by 

the model quite well. The cells have been placed in a cycle life test at 60% depth-of-discharge and 10 

deg C using a standard Vff charge control method. A recharge ratio between 102 and 103% was 

obtained employing a charge voltage level of 1.488 volts/cell, giving stable performance for the first 830 

cycles. At 830 cycles an equipment failure subjected the cells to 9 hr of overcharge at 1.2C, after which 

capacity was degraded and recharge voltages increased. Cycling is being continued at a higher voltage 

level. 

The standard individual pressure vessel nickel hydrogen cells used in modem satellites have 

been optimized over the years to have a relatively high energy density, in many respects compromising 

the voltage, power, or cycle life capability of these cells relative to the ultimate capability of nickel 

hydrogen battery technology. For applications not involving low altitude orbits, cell weight is often not 

the overriding battery issue. Efficiency, high pulse power capability, improved voltage regulation, or 

improved life capability (operation at greater depth-of-discharge) are often highly desirable battery 

performance traits. To evaluate how nickel hydrogen battery technology may better power these kinds of 

satellites, an effort has been made to optimize the state-of-the-art nickel hydrogen technology to reduce 

internal cell impedance by at least 50%, while enhancing the cycle life capability and overall robustness 

of the cell to handle high rates and high depths-of-discharge. This optimization was performed utilizing a 

first-principles nickel hydrogen battery model developed at The Aerospace Corporation. This model 

performed a full optimization of the standard MANTECH design with the constraint that no redesign of 

existing electrode types was allowed, and considered hydrogen, oxygen. electrolyte, and thermal 

management issues in addition to electrical performance. 

The optimized cell design basically consisted of a 'dual anode" stacking arrangement, in which 

each nickel electrode has both of its faces in contact with separator and a hydrogen electrode. This 

stacking arrangement requires twice the number of hydrogen electrodes as does the more standard 

back-to-back stack design, but has the advantages of equally utilizing both faces of the nickel electrodes, 

doubling the available surface area in contact with the separator, cutting mass transport and conduction 

distances within the nickel electrodes in half, and doubling the oxygen management capability of the cell. 

The computer model found that a nearly optimum electrolyte level was obtained using a single layer of 

Zircar separator, which provides a total electrolyte fill approximately the same as that obtained in the 

standard back-to-back stack design having two layers of Zircar separator. The optimized cell design also 

contained a zirconium oxide wall wick, twice the number of leads from the plates, 31% KOH electrolyte, 

a nickel precharge, and an axial cell terminal configuration. Because of the doubling in the number of 

hydrogen electrodes and the improved oxygen gas management capabilities of the design, the 

advantages of a recombination wall wick were found to be less significant than in the standard back-to1997 NASA Aerospace Battery Workshop Nickel-Hydrogen Session 
back design, and therefore was not included. The optimized cell design utilizes a 4.5 inch diameter cell 

assembly, and 30 mil thick nickel electrodes having 80% porosity plaque. 

Following these general design specifications, five cells were built by Eagle-Picher Industries, 

and are referred to as RNH 60-3 cells. Several compromises in the optimized design were made to 

obtain these cells using components that were readily available. First, a 3.5 inch rabbit-ear cell 

assembly was employed, giving a longer cell and longer thermal dissipation paths than were optimum. 

The 3.5 inch diameter cell hardware also did not provide enough core space to contain a lead from each 

plate, thus resulting in leads from each plate only for the plates in the lower half of the stack. This 

change was predicted to reduce the improvement in cell impedance by about 25%, but still giving an 

approximately 50% impedance improvement over the more standard cell designs. Finally, 35 mil thick 

nickel electrodes were used (slurry type), a change that was not expected to have a significant impact on 

performance. The maximum pressure expected in these cells is about 600 psi of hydrogen at full 

charge. 

The five cells were placed into an extensive characterization test regime to evaluate their 

capacity, voltage behavior, charge efficiency, and self-discharge rates as functions of state-of-charge 

and temperature. Characterization was done at 20, 10, 0, and -10 deg C. Figures 1 and 2 indicate 

charge and discharge voltages respectively over this temperature range for a standard charge (16 hr at 

C110) and discharge (Cl2 to 1.0 volts). As predicted and indicated in -Fig. 3, these cells have a higher 

discharge voltage and a lower charge voltage than do standard cells. Voltage behavior at a given 

temperature can best be summarized in terms of a set of IN curves which track voltage as a function of 

current at discrete states-of-charge. Figures 4 and 5 provide a summary of UV behavior at 0 and 20 deg 

C for 25, 50, and 75% states-of-charge. The slopes of these IN lines at high currents provide the best 

indication of internal cell impedance, including electrochemical reaction impedance, lead resistance, and 

the resistance of the polarized separators and plates. These cells have impedances of about 1 mohm at 

the higher states of charge, with a slightly higher impedance being seen at 0 deg C than at 20 deg C. 

Cells of this size (60 Ah) typically have in internal impedance of 1.5 to 1.8 mohm when all polarizations 

are included. 

The capacity of the cells was found to range from 58 Ah at 20 deg C, to nearly 80 Ah at -10 deg 

C. The capacity trend is indicated in Fig. 6, and shows that the capacity is best at about -10 deg C, 

having appeared to level off just below 80 Ah. Thus, these cells are conservatively rated at 60 Ah for 

usage in the normal temperature range of 0 to 10 deg C. For these cells, there appears to be no reason 

why the lower operating temperature range could not be extended down to -10 deg C. Figure 7 indicates 

charge efficiency as a function of state-of-charge for each test temperature. The trends in Fig. 7 mirror 

the cell capacity function of Fig. 6, with the quite high charge efficiency obtained at low temperatures 

giving a very high stored capacity. 

The self-discharge rate in this cell design based on the variations in observed total capacity from 

different duration discharges. Figure 8 indicates self-discharge as a function of state-of-charge at 20 deg 

C. As expected, the self-discharge rate increases with increasing state of charge. This increase is 

because the hydrogen pressure increases with state-of-charge and because the oxygen evolution rate 

increases exponentially at high states-of-charge. Fig. 8 indicates that even at 20 deg C, 100% state-ofcharge can be maintained with a Cl60 trickle charge rate. Trickle charge rates as low as Cl240 can 

maintain a 70% state-of-charge. A standard 72 hr charged stand test was conducted on these cells at 20 

deg C, with the result that they retained 83.9% of their capacity. These results indicate that the increases 

in plate area exposed to the separators and hydrogen electrodes have not resulted in a significant 

increase in charge losses for this cell design. 

After completion of characterization testing, the cells were placed into a cycle life test at 60% 

depth-of-discharge, utilizing a 90 minute cycle (30 minute discharge, 60 minute charge). The average 

cell temperature during the life test cycle was held at 10 deg C. The discharge was at 72 amps for the 

full 30 minute period. Recharge utilized a peak charge rate of 72 amps, with a voltage limit that at 

present is 1.488 volts per cell. When the voltage hits this level, the current is tapered back at a rate that 

maintains a constant voltage level within a 5 mv band. The goal at present is to return 102 to 103% of 

the discharged capacity during recharge for each cycle. The voltage level of 1.488 volts per cell gave 

102-1 02.4% recharge ratio and stabilized performance for about 830 cycles, after which a test equipment 

failure resulted in the cells being overcharged at 72 amps (1.2C) for 9 hr. After this event, as indicated in 

Fig. 9, the charge return ratio dropped to just above 100%, and the cells quickly ran down in capacity due 

to inadequate charge return. Re-characterization of the cells showed an increase of at least 30 mv in the 

1997 NASA Aerospace Battery Workshop -472- Nickel-Hydrogen Session 
recharge voltage curves for the cells, as well as about a 25% reduction in cell capacity. However, the 

cells all still delivered over 52 Ah of capacity at the 72 amp discharge rate. The cells are presently 

resuming l@e testing with an increased charge vdtage level. 

Figures 9-1 1 indicate the end-of-discharge voltage, end-of-charge voltage, and recharge ratio 

during the first 1100 cycles of the life test. At about cycle 200 the cells went through about 16 cycles 

with no discharge, but repeated recharge. Between cycles 200 and 830 the end-of-discharge voltages 

dropped about 6 mv, and the end-of-charge voltages have spread apart 3-4 mv. These changes are 

quite small, thus at present it is not clear if they signify drift in cell performance or simply a stabilization 

at slightly different levels. There was no significant change in end-of-charge or end-of-discharge 

pressure since cycle 200 up to the test anomaly at cycle 830, suggesting that there was not any 

significant loss in stored capacity. Performance will continue to be trended after the life test is restarted, 

again trying to attain the minimum stable recharge ratio that keeps end of discharge voltage above 1.1 

volts. 

The intent in this life test is maintain the recharge voltage at the minimum level that is just 

sufficient to maintain capacity of the cells. Failure is defined as voltage falling below 1.0 volts, combined 

wifh an inability to correct the low voltage by increasing recharge voltage level. It is anticipated that 

many years of cycling will be completed prior to cell failures, even after the accelerated aging that the 

cells experienced as a result of the recent test anomaly. Cells will be disassembled and analyzed when 

any failures occur, to determine the cause of the failure and whatever cell degradation modes are 

occurring along with the life cycling or as a result of the test anomaly. # 

1997 NASA Aerospace Battev Workshop Nickel-Hydrogen Session 
Figure i , 

C110 Charge Voltage vs. Temperature 

'I 

-10 deg C 

. . 

.................................................................................................................................................................. 

1, I 

........................................................................................................................................................ ,'................ 

I I I I I I I I I 

6 8 10 

Charge time (hr) 
Figure 2 

Cll 0 Disharge Voltage vs. Temperature 

.......................................... 

10 deg C 

20 deg C 

I 1.5 2 

Disharge time (hr) 
ChargeIDischarge Voltages for 

Different Cell Designs 

50 100 

Amp-hours (% of rated capacity) 
Figure 4 

I I 

IN Curves for Dual Anode Cells, 0 C 

at 25%, 50%, and 75% State-of-Charge 

1.35 

Cell Voltage 
Figure 5 1 

IN Curves for Dual Anode Cells, 20 C 

at 25%, 50%, and 75% State-of-Charge 

1.35 1.45 

Cell Voltage 

-- 

-- 

-- 

:# -- 

...................................................... 

...................................................... 

-3.43 mohm 

...... .... ..........."........ ...... ......... .......................... 

/'lhm I, 

".." ""' '."" "" ..................................................................................................................................................................... I 

I 

I I 

I I I I 
Figure 6 

I 1 

Standard Capacity vs. Temperature for 

Dual Anode Cells 
Figure 7 

Charge Efficiency for Dual-Anode Cells 

CI10 Charge Rate 
Figure 8 I 

Self-Discharge Rate at 20 deg C 

for Dual Anode Cells 
Figure 9 

Dual Anode EODV in Life Test 

400 600 800 I000 1200 

Cycle Number 

Figure 11 

Dual Anode Recharge Ratio in Life Test 

0 200 400 600 800 I000 I2OQ 

Cycle Number 
i 

I 

i 

Y' 00 Development of a Large Diameter 

Nickel-Hydrogen (NiHd Individual 

Vessel (IPV Battery Cell 

1, 11 I 

3 

B Dwight Caldwell, Jim Morgan, Jack Brill 

k Tebhnologies Division, Eagle-Picher Industries, Inc. I 2 

Q 8 

3 Joplin, MO 64802 

Fa 

8 

6' I 

623-8000 3 
OUTLINE 

DESIGN RATIONALE 

. . DEVELOPMENT PLAN 

CELL DESIGN 

Mechanical 

Electrical 

Thermal 

ACCEPTANCE TESTING 

LIFE TESTING 

FWTURE DEVELOPMENT PLANS 

SUMMARY 
NiH, ADVANTAGES i I 

Reliable, Durable 

High Energy Density 

High Power Density 

Sealed, Maintenance-Free Battery 

Excellent Overcharge, Overdischarge and 

Deep Cycle Capa bilifies I# I I 

SOC Easily Determined by Internal 

Hydrogen Pressure I ! 

No Toxic Materials 

Competitive Life Cycle Cost 
NTS-2 (US NAVY) [35 AH]' 

lNTElSAT V 

(FACCANTELSAT) [30 

G-STAR (RCAIGTE) [30 

SPACENET 

(RCAISOUTHERN PACIFIC) [40 

ASC (AMERICAN SATELLITE 

CORPORATION) [35 AH]' 

SATCOM K (RCA SPACE 

COMMUNICATION) [60 

ITALSAT (FACCISELENIA) [30 AH]' 

MILSTAR (LMSCIAIR FORCE) 176 

ASTRA IA, 1 B (GUASTRO)'[~O 

ANIK-E 

(RCAtTELESATCANADA) [30 

PAN AM SAT (GEIASTRO) 135 AH]' 

INTELSAT K (GEIASTRO) [SO AH]' 

ASC II (GWASTRO) 140 AH]' 

C3lC4 (GUASTRO) [50 

AURORA II (GWASTRO) 140 AH]' 

TELSTAR 401 (GEIASTRO) [50 AH]' 

INMARSAT 3 (GWASTRO) [60 

8482 (TRW) [~SAH-76 AH] 

SALT 

(INTRASPACUNAVY) [I2 AH] 314 

EOS (GUASTROINASA) 160 AH] 

TESS(McDonne1l Douglas1 

NASA) [50 AHI3 

ORBCOMM 

(ORBITAL SCIENCES) [I 0AH]' (2)a 

GPS IIR (GEIASTROI 

AIR FORCE) 140 

KOREASAT (GUASTRO) [42 

SUPER (MARTIN-MARIETTA1 

AIR FORCE) [50 AHI3 

BS3N (GEIASTRO) [50 AM1 

AXAF (TRWINASA) [40 AH] 

AXAF (NASA-MSFC) [50 AH] 

SPACE STATION FREEDOM 

(SSILORAUNASA) [81 AH] 

GEOSAT Follow-on 

(BALL AEROSPACUNAW) 120 AHI4 

TUBSAT-B (TECHNICAL UNIV. 

OF BERLIN)[G AH]'t4 

GEl (MARTIN-MARIETTA ASTRO) 

[I 00 AH]' 

GE2 (MARTIN-MARIETTA ASTRO) 

[I 00 AH]' 

TEMPO (SSILORAL) [I 20 AH]' 

LANDSAT '/(MARTIN-MARIETTA1 

AIRFORCE) [50 AH] 

GLOBALSTAR (SSILORAL) [64 AH] 

PanAmSat 6 (SSILORAL) [I20 AH] 

AGlLA 2 (SSILORAL) [I20 AHJ1 

2828 (TRW) 198 AH] 

TUBSAT-C (TECHNICAL UNN. 

OF BERLIN) [I0 AHl4 

MICROLAB 

(ORBITAL SCIENCES) [I0 

TELSTAR 501 (SSILORAL) [I 20 AH]' 

GE3 (Lockheed-Martin) [I 00 AH11 

INSAT 2E (ISRO) [55 AH] 

APSTAR 2R (SSILORAL) [I 20 AH]' 

P-81 (Lockheed-Martin) 177 AH]I 

TELSTAR 402R (GUASTRO) [45 AH11 

SSTl Lewis (TRW) 123 AH114 

lSSA (SSILORAL, NASA) [81 AH1 

GOES (SSRORAL, NASA) [I 6 AH] 

DBSC (Lockheed-Martin) [I00 AH] 

PanAmSat 7 (SSILORAL) [I20 AH] 

PanAmSat 8 (SSILORAL) [I20 AH] 

CHINASTAR (Lockheed-Martin) [I00 AH] 

PIONEER 1&2 (SSRORAL) [I20 AH] 

Whitetail I& 2(Lockheed Martin 

Missiles & Space)[lOO AH] 

ACeS (LOCKHEED MARTIN 

MISSILES 81 SPACE) [I31 AH] 

SMTS (TRW) [60 Am 

IR&D 

(NAVAL RESEARCH LAB) [SO AY]b 

EOS Common (TRW) [I60 AH] 

QUICKBIRD (BALL AEROSPACE) [40 AH14 

MTI (BALL AEROSPACE) [4O AH14 

41 71 (TRW) [86 AH]I 

MILITARY (CLASSIFIED)' 

One (or More) Spacecraft in Orbit 

(128 Total) 

Waiting for Launch Availability. 

Flight Qualified Hardware Delivered, 

Program on Hold or Temlnated 

Common Pressure Vessel (CPV) 

Dependent Pressure Vessel (DPV) 
1. FLIGHT EXPERIMENT 18. BRILLIANT EYES 

(US AIR FORCE) [50 AH]' (ROCKWELUAIRFORCE) [30 AH] 38. 3890 [90 AH] 

2. OLYMPUS (BAelESA) [58 AH]' 19. TIMED (JACKSON & TULU 39. 3893 190 AH] 

3. EUTELSAT II NASA-GFSC) [I9 AH]' 40. EUTELSAT W24A 

(AEROSPATIALE) [58 (AEROSPATIALE) [68 AH] 

4. WSAT II 

20m sPV BATTERY (TRWIIR(LD) [36 41. ARABSAT 2B (AEROSPATIALE) P2 AH] 21. SMALLSATIIRIDIUM (Lr'SC) t50 "I6 42. NAHUEL I B (AEROSPATIALE~ 158 (AEG - TELEFUNKEN) [30 AH]" 

6. HUBBLE SPACE TELESCOPE 22' (AEROsPATIALE) 43, IRAD46 (LOCKHEED MARTIN) 115 AH16 

(LMSCINASA) 190 AH11 23. NAHUELIA 

(AEROSPATIALE) [58 AHll 44. IRIDIUM (LOCKHEED MARTIN) 

6. TELECOM I1 (MATRA) f18 AH]1(4) 24. CLEMENTINE [I 5 AH]"*6 160 AH]"(ls)fi 

7. HISPASAT (MATRA) [78 AHI"(2) 26. IRIDIUM (LMSC) [SO AH] 1(20)*6 

46. MSP (SPECTRUM ASTRO) [I6 AH14 

8. ORION (BAe) [78 AH11 26. TURKSAT IC 46. GPS IIF (ROCKWELLIBOEING) f16 AH] 

0. TURKSAT IB (AEROSPATIALE) [58 AH] 1 47. SICRAL (ALENIA SPAZIO) 155 AH] 

(AEROSPATIALE) [65 AH]" 27. SSTlClark (CTA) [I5 AH] 48. STARDUST (LMA) [I 6 AH]' 

10. APEX (ORBITAL SCIENCES1 

AIR FORCE) [6 AH1114 

28. MGS (MARTIN-MARIElTA) [20 AH] 14 

11. SEA STAR 29. HST Follow-On (LMSC) [90 AH] 1 One (or More) Spacecraft in Orbit 

(ORBITAL SCIENCESINASA) ~0 ~~11~4 30. DMPS (LOCKHEED MARTIN) [IOO AH] ' (128 Total). I 

12. INTELSAT 8 (GEIASTRO) 145 & 50 

2 Waiting for Launch Availability. 

(AEROSPATIALE) 152 AH] 3 Flight Qualitied Hardware Delivered, 13. P9l-IIARGOS 32. SAFIR (OHB) [4 AH] Program on Hold or Terminated 

(ROCKWELUAIR FORCE) 145 AH] 4 Common Pressure Vessel (CPV) 

14. ASIASAT 2 (GEJASTRO) 145 AH]' 33. 3810 [90 AH] 5 Single Pressure Vessel (SPV) 

16. ECHOSTAR (GEIASTRO) [50 AH]'(3) 34. 3822 [90 AH] 

16, MSTI-2 36. 3857 [90 AH] 

(SPECTRUM ASTROISDIO) [I0 AH1114 36. 3866 190 AH] 

17. MSTI-3 37. 3889 I90 AH] 

(SPECTRUM ASTROISDIO) 112 AHJ1*4 
4 RATIONALE FOR 5.5 INCH CELL DEVELOPMENT 

b 

\O 

7 * Satellite Manufacturers are Demanding Increased 'Power 

* Large Diameter, High Capacity Cells Demonstrate the 

Following Benefits 

* Reduced Cell and Battery Weight 

lncreased Specific Energy 1, 

2 6' t 

E- * Decreased Length 
I 

G 

I 

COMPARISON OF 4.1 in. AND 1.1 in. DIAMETER 22 CELL BATTERY CONFlGURATlONS 

I 

200 250 300 360 ' 400 460 600 

CELL CAPAClN (AH) 
b0 COMPARISON OF 4.5 In. AND 5.5 in. DIAMETER 22 CELL BATTERY CONFIGURATIONS 

i 

V 

3 200 260 300 360 400 460 

CELL CAPACITY (AH) 
s 

I 

5.5 INCH CELL DEVELOPMENT PLAN 

Guideline: The 5.5 lnch Technology is ScaIed4.Jp 

Directly from the 2.5, 3.5, and 4.5 lnch Designs 

Electrochemistry is Unchanged 

Stack Component Materials and Processes 

are /dentical '* 

Positive Electrode 

Negative Electrode 

Separator 

Gars Screen 
s 

3 5.5 INCH CELL DESIGN 

b 

P Pressure Vessel - laconel 718 

Max Design Pressure - 800 psi 

3.5 : . f Safety Factor 

Ail Pressure Vessel We/ds 

Laser or Ejectran Beam 

MIL-STD-7529A Cyci@ and Burst resf 

3 

ct Comp8ete s. 

& k a Core/End Pjates Designed Proportional ;t to Expected Loads and Growth 3 

GY c ZIrcsnCum Oxide Wall Casting 5. 

3 
.n" 5.5 INCH CELL DESIGN 

ELECTRICAL: 

"Mantech" Design, "Back to Back9' Electrode 

Arrangement 

Slurry Slnter Positive Electrode 

LighWeight Catiljfls"~ Hydrogen Negative 

Electrode 

31% KQN ElectroIfle 

Zireaa Separator 

e' PolypropyIene Gas Screen 
.a0 5.5 INCH CELL DESIGN 

u3 THERMAL: 

? 

ThermaB CraassPSecti~n Majntain ed 

in Siring Cell Components 

Therma8 Gradjen9 Optimized 

Large Thermal Transfer Area 

Along B".VV Wall I 

Minimal Sfack Edge to Vessel Wail Gap 3 ThermaB Mode\ wi/I be Deveioped 

k '-4 Q 

Early Next Year 

a 

2 

2 

M 

6 

3 
5.5 INCH CELL DESIGN SUMMARY 

Cell Type 

Nominal Voltage 

Rated Capacity 

Actual Capacity 

Weight 

Diameter 

Length 

Specific Energy 

vessel Wall Thickness 

MEOP 

Safety Factor 

RNH 200-7 

7.25 Volts 

200 Ampere-hours 

205 Ampere-hours 

(C/2 to ?.OV @ 10°C) 

12 lbs. (5.45 Kg) 

5.7 1 " (7 4.50 cm) 

74.47: (36.60 cm) w/terminals - axial design 

4 7.02 Wh/Kg 

0.040" (I. 02 mm) 

800 psig (55. 77 bar) 

3.5: 7 
4 5.5 INCH ACCEPTANCE TESTING I I 

j 

-L 20°C Capacity Test: 188 Ah (Avg.) 

\O 

00 

I 

70°C Capacity Test: 205 Ah (Avgn) 

10°C Charge Retention: 185 Ah (Avgm) (90% of 10°C Capacity) 

(72 Hour Open Circuit) 

0°C Capacity Test: 220 Ah (A vgm) 

-lO°C Capacity Test: 210 Ah (Avgn) 

Note: Charge rate - C/lO or 20 Amps 

Discharge Rate - C/2 or 100 Amps 

Capacity Measured at 1.0 Volt 
5.5 INCH CELL LlFE TEST 

RNH 200 -1 LlFE TEST SUMMARY (LEO) 

0.9 

0 10 20 30 40 50 60 90 80 90 100 

CYCLE NUMBER 

+ EODVl 

* EODV2 

+ EODV3 

-M- EODV4 

-!+ EODV5 

+ EODVB 

+ EODV7 

- EODV 8 

- EODV9 

+ EOCV1 

-B- EOCVP 

+ EOCV3 

-I- EOCV4 

+ EOCV5 

+ EOCVB 

+ EOCV7 

- EOCV8 

- EOCVQ 
a" FUTURE 5.5 INCH DEVELOPMENT PLANS 

Parts in Production 

Activation and Testing Scheduled Early 1998 

, 100 Ah CPV Cell 

* Parts in Production 

Activation and Testing Scheduled Early 1998 
SUMMARY 

clr, 

E 5.5 INCH RNH 200-1 CELLS BUILT AND ON TEST 

CELL PERFORMANCE AS PREDICTED 

FURTHER DEVELOPMENT IS PROCEEDING RAPIDLY 1. I 

I 
Page intentionally left blank 
Charge Temperature 

Characterization of Nickel-Hydrogen 

Batteries 

D. Coates, R. Smith, D. Judd, J. Brill 

J. McDerrnott* and G. Gates* 

Eagle-Picher Technologies 

Joplin, Missouri' 

"Lockheed-Martin 

Denver, Colorado 
Nickel-Hydrogen Battery 

Charge Characterization Test 

The purpose of the test is to determine the 

V/T level charging characteristics and 

charging efficiency of the battery as a 

function of temperature and current under 

realistic mission-oriented conditions. 
Test Article 

Eight-cell pack 

RNH- 100-9 cells 

Dual-stack IPV 

Flight configuration hardware 

Sleeve mounted 
Test Configuration 

Eight cells sleeve mounted into battery pack 

Pack mounted on coldplate in chamber 

No chamber refrigeration 

2 fans for air circulation 

4 baseplate thermocouples \ 

1 chamber air thermocouples 

1 thermocouple per cell 
Test Conditions 

V/T level 7 thru 12 (0.5 step increments) 

Testing from 0" to 20°C (5°C increments) 

Charge rates of 35A and 50A 

155 steps total 

Two charge characterization sequences 

', l 

98 minute cycles (63135) I 

Discharge load profile 
Test Sequence 

Charge to the V/T level 

0 Taper charge to 1 05 Ai return 

- if 105% is reached - switch to trickle charge 

Perform constant current capacity (CI2) 

Re-charge to V/T level as above 

Cycle to stabilization I 

Perform constant current capacity ((212) 
Charging Methodology 

63 minutes total charge time 

Charge to VIT level (35A or 50A) 

Taper charge until: 

- 105% recharge level reached 

Trickle charge 1.57A 
Stabilization Criteria 

The EOD voltage over 5 consecutive cycles 

changes by less than 0.003V 

The EOD pack voltage drops below 9.164V 
Discharge Load Profile 

149 Watts for 10.5 minutes 

227 Watts for 7.4 minutes 

635 Watts for 16.1 minutes 

753 Watts for 1 minute 
VIT 9 STABILIZATION CYCLE 

30 40 50 60 

ELAPSED TIME (minutes) 
TYPICAL DISCHARGE LOAD PROFILE 

VIT 8 DISCHARGE 

753 Watts 1 minute 

635 Watts 16.1 minutes 

149 Watts 10.5 minutes 

227 Watts 7.4 minutes , 

I 

IN I I I,( 4 

0.3 

TIME (hours) 
STABILIZATION CYCLES 

V/T LEVEL 
50A CHARGE STABILIZATION CYCLES 

TIME (hours) 
CHARGE STABILIZATION CYCLES 

TIME (hours) 
CHARGE STABILIZATION CYCLES 

9.5 

9.0- III,,,,,, 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 

TIME (hours) 
CHARGE STABILIZATION CYCLES 

1 

TIME (hours) 
50A I 35A - 10°C CHARGE CHARACTERIZATION 

-- O - -5OA 

VOLTAGE . 

VIT LEVEL 
50.44 - W°C CHARGE CHARACTERIZATION 

7 7.5 8 8.5 9 9.5 10 10.5 I I 11.5 12 

VIT LEVEL 
VIT 9 IMPEDANCE ANALYSIS 

STATE-OF-CHARGE (%) 
Conclusions 

10°C testing completed 

Optimal charging occurs at: 

Charge efficiency is similar at both rates 

Temperature effects have not ( yet been 

determined 
MEETING THE DEMANDS OF THE 

COMMERCIAL AEROSPACE MARKET 

Jeff Dermott and Dwight Caldwell , I 

1 

Eagle-Picher Industries, Inc. 

Joplin, MO. 
Outline 

BackgroundIHistory of 

Matrix Life Test Data 

Acceptance Test Data 

Acknowledgements 

Facility 
+ Conversion to Nickel-Hydrogen Production Began in 1994. 

+ Manufacturing Philosophy 

!? 

Z 

9 

F 8 

Implement a Paperless System for Data Management, 

Quality Control, Process Control. 

+ Originally Designed for High Volume Nickel-Iron Production. 

Reduce Inspection Points by Incorporating Operator 

Inspections. 

I I 

, 

3 

P 

& Y 

2 

3 

Build Quality Into the Product by Using Process control. 
Qualified Processes 

+ Since 1994 the Following Processes at RLP Have 

Been Qualified for Flight Use: 

* Positive Electrode Manufacturing and Testing 

Negative Electrode Manufacturing and Testing 

* Pressure Vessel LASER Welding 

* Pressure Vessel Cycle and Burst Testing 

Pressure Vessel Flame Spraying 

IPV Assembly I SPV Assembly 

.SPV and IPV Acceptance Testing 
Qualified Designs I\ 

+ 10 Inch Diameter, 50 and 60 AH Single Pressure 

Vessel (SPV) Batteries. 

+ 4.5 Inch Diameter, 120 AH Independent Pressure 

Vessel (IPV) Cells 

+ 3.5 Inch Diameter, 64 AH IPV Cells 

+ All Three Designs Are Produced on Programs I Requiring Monthly and/or weekly Deliveries. 
RNH 64-1 .Matrix Life Test '1 

+ First RLF, IPV Life Test 

+ Test Uses High Rate 70% DOD Cycles to 

Accumulate Wear on the Cells. 

+ Every 1000 Cycles the Cells Are Run for 100 

Cycles at 40% DOD As a Health Check. 

\ I t 

, 
Matrix Life -Test Cycle Regime I\ 

+ Test Temperature = 0 C 

+ 70% DOD Cycles: 

Charge at a C Rate. With a Taper at the End of Charge. 

Discharge at a 1.2C Rate for 35 Minutes. 

* RR = 110%, Total Cycle Time = 90 Minutes 

+ 40% D.0D Cycles: 

Charge at a Cl2 Rate With a Taper at the End of Charge. 

Discharge at a C11.45 Rate for 35, Minutes. I 

I 

RR = 110%, Total Cycle Time = 90 Minutes 
b 

'0 

'0 

'4 

5 

fa b 

8 

rr 

m 

P 

a 9 

? 8 E 

% 

W 

? 

$ 

B & V Sa. a tl 

Descriptiqn of Cells in Test 

I 

I 
Cycle Status 

1 + Currently the Cells From Boiler Plate Lot Have 

Accumulated 13,400 70%DOD Cycles and 1,300 

40%DOD Cycles. 

+ The Engineering Model Cells Have Accumulated 

-11,400 70%DOD Cells and 1,100 40%DOD Cells. 
Discharge Rate = I .2C 

I 

r 

RNH 64-1 MATRIX LIFE TEST 

70% DOD CYCLES I 

1.250 

1.200 

1.150 

> g 1.100 

W 

8 1.050 

i3 9 1.000 

4 

0.950 

0.900 

0.850 

0 2000 4000 6000 8000 I0000 12000 14000 16000 

CYCLE NO. 
CONTROL CELLS 

40% DOD CYCLE DATA I( 

1.25 1 

Discharge Rate = Cl1.45 RR = 1.05 

................................ 

- After -1 3000,70%DOD Cycles 

W 

(3 

1.21 .................................................................... .......... 

W 

> 

a ............................................. 

............................................. 

\ 

1.18 ! I I I t I I -- : 

0 10 20 30 40 50 60 70 80 90 100 

CYCLE NO. 

I 
I 

* 

BP-V3 CELLS 

40%.DOD CYCLE DATA i 

1.25 -. 

Discharge Rate = CII .45 RR = 1.05 

w 

(3 

1.21 

W 

> 

4 1.2 

1.19 

1.18 

.......

............................................................................ * .....------...-.--. 

................................................................................................ 

\ 

t I I I I I I t I 

0 10 20 30 40 50 60 70 80 90 100 

CYCLE NO. 

. - ---- --~ 
Discharge Rate = Ci1.45 RR = 1.05 

--*-.-..--..--- 

-After -1 3000,70%DOD Cycles 

--.--.....-.-.. 

--..L..~~.~.--.~.~-.....--..-----.-..-.........--..----....-...............--.......~........~.

\ 

I I I I I I I I I 
I 

EM-V2 CELLS 

40% DOD CYCLE DATA '\ 

1.25 

1.24 

1.23 

> 

Discharge ate = C11.45 RR = 1.05 

................................................................................................ 

................................................................................................ 

cl 

........................ 

. . 

1.18 1 

0 10 20 30 40 50 60 70 80 90 100 

CYCLE NO. 

\ 

1 1 t 1 I I 1 1 1 

I 

1 
Summary ,I 

+ All 0.035" Slurry Designs Are Still Performing Well. 

+ The RLP Impregnated Electrodes (EM-V2) Show the 

Least Amount of Voltage Decay. 

+ Three Additional Cell Designs Have Been Added to 

the Test Recently. 

+ A Control Cell DPA Is Planned to Determine Cause 

\ I 

of EODV Decline. I 
RNH 120-1 Acceptance Test Data ,I 

+ First 4.5" IPV Flight Programs (800 Cells Delivered). 

+ Represents a Mixture of Data Prom C Street 

Impregnated Positives and RLF Impregnated Positives. 

+ Cell Lots Are Being Shipped on 4 to 6 Week Centers 

With 6 Month Build Times. 

I I , 

1 
RNH 120-1 CAPACITY DATA 

ACCEPTANCE TEST 

RESULTS 

0 Id 

n 

- " 

" 4-, u 

60 I I I I t t I I I 

0 1 2 3 4 5 6 7 8 \ 9 

LOT NO. I 

I 
I 

RNH 120-1 WEIGHT DATA 

\ 

2700 

iij. 

3 2650 

C) 

V 

.................................................................................................. 

I- 

- - - * - . - - -' 

' A d 2550 

0 

W 

C) 

2500 

5 - 

a 

2450 

................................................................................................. 

................................................................................................... 

I I I I \ I I I I 

I I I I 1 

1 2 3 4 5 6 7 8 9 10 11 

CELL LOT 

~-~ 
RNH 120-1 CAPACITY DATA I 

Discharge at 75 amps to I .O V 

CELL CAPACITY (AH) 
Summary 

+ RNH 120-1 Data Has Established the Ability to 

Repeat Cell Performance Over Multiple Lots. 

+ Tracking This Data Has Aided in the 

Consistency Achieved. 
Acknowledgements 

+ John Mauer: Eagle-Picher Industries, Inc. 

+ David Judd: Eagle-Picher Industries, Inc. 

+ Dan DeBiccari: Space SystemsILoral 

+ John Hall: Space Systems/Loral 
Page intentionally left blank 
Nickel Electrode Design Focused Session 

1997 NASA Aerospace Battery Workshop -545- 

-- - 
Page intentionally left blank 
Cause for Second Plateau Discharge in Nickel 

Electrodes 

A. Zimmerman and N. Weber 

Electronics Technology Center 

The Aerospace Corporation 

El Segundo, California 90245 

for presentation at 

The 1997 NASA Battery Workshop 

18-20 November 1.997 

Huntsville, Alabama 

1997 NASA Aerospace Battery Workshop -547- Nickel Electrode Design Focused Session 
- 

Cause for Second Plateau Discharge in Nickel Electrodes 

A. Zimmerman and N. Weber 

Electronics Technology Center 

The Aerospace Corporation 

El Segundo, California 90245 

Abstract 

Nickel electrodes typically exhibit a lower discharge plateau at about 0.9 volts vs. 

hydrogen. This lower plateau can become quite significant in degraded nickel hydrogen cells. 

The possible mechanisms for this lower plateau are presented, and the data either supporting 

or refuting each mechanism described. It will be demonstrated that~nly one mechanism for 

the second plateau is fully consistent with the existing data base. This mechanism predicts the 

second plateau to result from coupling between parasitic oxygen reduction and the IN 

behavior of a barrier layer at the junction between the semiconducting active material and the 

current collector. Methods will also be described that are capable of either accentuating or 

reducing the second plateau discharge in nickel hydrogen battery cells. 

Capacity loss from nickel electrode degradation in nickel hydrogen battery cells is often 

manifested by the loss of discharge capacity from the normal discharge voltage plateau at 

about 1.25 volts, accompanied by an increase in discharge capacity at a lower (or "second") 

plateau at about 0.9 volts, as illustrated in Figure I. While this lower discharge level is always 

present to some degree in the discharge of nickel electrodes, it is only in the nickel hydrogen 

cell environment that is seen to become accessible for lengthy discharge at quite high 

discharge rates. A number of mechanisms for this lower plateau have been proposed over the 

years. Here the consistency of each of these mechanisms with the available data base will be 

evaluated, for the purpose of determining the underlying mechanism for the second plateau. 

The following list of mechanisms have been either proposed, or have the possibility of causing 

a lower discharge plateau in nickel electrodes: 

1. An alternative electrochemical reaction. Possibilities include reduction of oxygen gas, 

reduction of gamma-phase active material, or the reduction of some unidentified material. 

These possibilities all require significant quantities of an alternative active phase. 

2. The drop from the upper discharge plateau is caused either by depletior, of charge in the 

active material or by the development of a high resistance in the active material due to 

charge depletion, and probably in most instances is caused by a combination of these two 

effects. If conductivity can increase with decreasing electrode voltage, and there exists a 

parasitic electrochemical process that can pin the voltage at the interface between the 

active material and the current collector, a lower discharge plateau can occur at the pinned 

voltage level. This mechanism can give multiple lower plateaus if there are different types 

of interfaces present, since each interface can pin the voltage at a different level. 

I997 NASA Aerospace Battery Workshop -548- Nicke 1 Electrode ~esi& Focused Session 
3. Physical variations in contact area of active material to current collector can shift the 

overpotential for the normal discharge reactions to greater values, thus moving the 

discharge plateau downwards. If there are two populations of active material, distinguished 

by either very good or very poor contact area with the current collector, it is possible for 

these populations to discharge at different voltage levels. - - 

To differentiate between these or other mechanisms, it is necessary to have a collection of 

observed data. The following list is a collection of observations made for nickel electrodes that 

. are pertinent to the mechanisms for the second plateau. 

1. In some cases, the amount of capacity discharged in the second plateau can exceed the 

amount discharged in the upper plateau. 

2. The second plateau disappears when oxygen gas is not present at the nickel electrode, i.e. 

evacuating all gas from a nickel hydrogen cell then backfilling with hydrogen. However, the 

capacity encompassed by the second plateau remains in an undischargeable state. This 

behavior is indicated in Figure 2. The regeneration of oxygen allows the lower plateau 

material to be again discharged. 

3. The gamma-NiOOH phase, in a homogeneously discharging electrode, discharges at a 

voltage about 40 mv below that of the beta-NiOOH phase. This reaction occurs with 

kinetics that are not significantly poorer than discharge of the normal beta-phase, as 

indicated in Figure 3, For this reason it is thermodynamically impossible for significant 

gamma-phase active material to exist in close proximity to the current collector when the 

voltage has dropped down to the second plateau. 

4. The depleted nickel oxyhydroxide active material is a semiconductor, having a doping 

density equal to the density of charged sites remaining. Thus, if a voltage exceeding a 

bandgap of about 0.3 volts develops across a layer of this active material, its electronic 

conductivity will increase exponentially as the voltage across the layer increases. 

5. The heat produced by a nickel electrode while discharging on the second plateau gives a 

enthalpy change for the second plateau reaction that is identical to that for the reaction on 

the normal discharge plateau, i.e. the added voltage drop between the upper and lower 

discharge plateau all shows up as heat according to the product of the voltage drop and 

the current. Data measured utilizing dynamic calorimetry, which has a time constant of 

less than 30 seconds, indicate this in Figure 4. 

6. In the hydrogen gas environment of the nickel hydrogen cell, extensive second plateaus 

are often seen, with the amount of second plateau capacity generally being proportior~al to 

the amount of gamma-phase active material generated. However, typically there is much 

more overcharge when larger amounts of gamma-phase are formed. When the same kind 

of sintered electrode is evaluated in a flooded test cell, there is no good correlation 

between the amount of second plateau seen and the amount of gamma-phase or the 

amount of overcharge. 

7. The reversible potential seems to be the same on both the upper and lower discharge 

plateaus. 

8. Increasing the amount of cobalt additive in the active material shifts capacity from the lower 

plateau to the upper plateau, and also increases the rate of recovery from the lower 

plateau to the upper plateau if the discharge current is decreased. 

9. The voltage level associated with the second plateau depends on the pressure of 

hydrogen gas in the nickel electrode, increasing with increasing hydrogen pressure. 

10. The kinetics for the discharge reaction undergo a significant change, as indicated by the 

IoglN curves in Figure 5 for a nickel hydrogen cell undergoing discharge on the upper and 

lower plateaus. 

1997 NASA Aerospace Battery Workshop -549- Nickel Electrode ~esi~n Focused Session 
The first mechanism to be considered is simply that reduction of oxygen, either 

adsorbed or gaseous, causes the second plateau. There are several reasons why this is not 

likely to be a valid mechanism. First, the amount of oxygen needed to give 50% of the 

electrode capacity in a second plateau is prohibitive, particularly in.. a nickel hydrogen cell 

where the oxygen will bum in the hydrogen atmosphere. Second, the reversible potential and 

enthalpy change due to oxygen reduction should be slightly different from the normal 

discharge process, a difference that has not been detected. Third, the second plateau position 

should vary with oxygen activity in the nickel electrode. The second plateau voltage level 

actually seen does not vary significantly, but seems pinned somehow to a relatively narrow 

range of potentials. We have constructed a nickel-oxygen cell employing a sintered nickel 

electrode, which when filled with 114 psia of oxygen gas could continuously support oxygen 

reduction at 0.1 ma/cm2 at an overpotential of 800 mv. Clearly this kind of process cannot 

support the 10 ma/cm2 discharge rates seen on second plateaus in nickel hydrogen cells. 

The second mechanism to be considered is that electrochemical reduction of gammaphase to some other reduced phase causes the second plateau. Since gamma-phase is 

reduced at voltages about 40 mv below beta-phase reduction (Fig. 3), it is thermodynamically 

impossible to retain significant gamma-phase in the face of a 300-400 mv overpotential, unless 

it is stabilized by a high resistance layer between the gamma-phase--material and the current 

collector. If this is the case, then it is the high resistance layer and its conduction properties 

that are responsible for the lower discharge plateau. In addition, it is difficult to envision any 

electrochemically active material that always discharges at an overpotential of 300-400 mv 

independent of rate, structure, or other parameters that normally alter electrochemical 

overpotentials. 

The third mechanism to be considered is the possibility that two distinct populations of 

active material exist, one having a very high surface area contact with the current collector and 

the other having a very low surface area contact. This is consistent with the observation that 

the amount of second plateau seen in nickel hydrogen cells typically increases as the 

electrodes degrade. Here again, one would expect to see second plateau levels covering the 

entire voltage range from the upper level and downwards, as the physical structure of the 

active materiat shifts and degrades. Since this behavior is not seen, this cannot be regarded 

as a viable mechanism. 

The final mechanism to be considered attributes the lower voltage plateau to the 

intrinsic semiconducting characteristics of the active materials in nickel electrodes. As the 

discharge process begins to deplete the active material at its junction with the current collector, 

a layer of depleted material begins to grow out from the current collector. Because, this layer 

has poor conductivity, a significant voltage drop develops across the layer, i.e. between the 

current collector and the remaining charged active material. When this voltage drop exceeds 

the bandgap of the nickel oxyhydroxide, about 0.3-0.4 volts, the depleted layer wiH increase in 

conductivity much like a diode responds when forward biased. While this enables electronic 

conduction through the depleted layer and discharge of othewise isolated charged material, it 

cannot give a second plateau unless some electrochemical process can act to pin the potential 

at the junction between the active material and the current collector. The oxygen reduction 

process provides just such a process, requiring only microamps or less current to support the 

potential at the current collector surface, while the bulk of the current is carried electronically 

through the depleted layer to discharge remaining charged active material. If there is only a 

very limited supply of oxygen gas at the surface of the current collector, the second plateau will 

terminate. The diagram in Fig. 6 illustrates this mechanism. 

This last mechanism seems to be fully consistent with all the observed. behavior of 

nickel electrodes. By definition, it involves identically the same reactions on the lower plateau 

I997 NASA Aerospace Battery Workshop -550- Nickel Electrode Design Focused Session 
as on the upper plateau. Since it requires electronic conductivity through a depleted layer, it 

should be sensitive to any materials that can alter the conductivity of the depleted layer. 

Cobalt additives provide sites that can significantly improve proton diffusion rates, thus 

increasing the ability of the active material to maintain conductivity. Similarly, hydrogen gas 

can undergo catalytic reaction at the metal surfaces at low rates, resulting in the direct injection 

of protons into the depleted active material and its enhanced ability to conduct charge through 

the depleted layers. It is this exchange between hydrogen gas and protons in the active 

material that enhances the ability to obtain lower plateau discharge at high rates in the 

gaseous environment of the nickel hydrogen cell. 

The mechanism described above and illustrated in Fig. 6, which is consistent with the 

body of performance data, indicates precisely how one should go about minimizing the second 

plateau. First, assure that all active material is in intimate and close contact with the current 

collector. The sintered electrode, until it undergoes significant expansion, blistering, or 

corrosion after extensive cycling, is an excellent example of this kind of structure. Second, 

improve the proton diffusion rate in the active material. This will delay the formation of 

depleted layers, as well as improving electronic conductivity. The use of cobalt 'additives 

provides this function, when the cobalt occupies a site in the active material lattice that 

otherwise would contain a nickel ion. Additives that do not occupy nickel sites within the lattice 

can also improve conductivity by enhancing conduction across the gfain boundaries between 

the active niaterial crystallites. 

1997 NASA Aerospace Battely Workshop Nickel Electrode Design Focused Session 
Figure I 

Cl2 Disharge Voltage for NiH2 Cell 

Showing second plateau 

1 1 .5 

Disharge time (hr) 
Figure 2 

C/2 Disharge Voltage for NiH2 Cell 

Showing second plateau I 

........................................................................................... 

0 0.5 I 1.5 2 2.5 

Disharge time (hr) 

1.2 - 

a, cnl,q - 

Cb 

m 

.c, 

>O 

m 

m 

0.9 

0 s 8 

0.7 

0.6 

-45 

- ................................................................................................................. ......................................... 

-. ......................................... 

I 

I 

-- .................................................................................................................................... ............................... 

1 

I 

-- I 

1 

-- ..................................................................................................................................... ....................................... 

O2 removed! -- ....................................................... ........................................................................ I 

I , 

I 

I 

:~-+-{~t-$--tClf--( --'I
Relative Redox Potentials of P-NiOOH and y-NiOOH 

Voltage vs. Hg/HgO 
Figure 4 

~nthalpy Change for Discharge Reaction During . . - . . - - 

Discharge on Upper and Lower Plateaus. 

1 

...................................... .-- 

....... ............___ .................-. 

. . 

................... ...................... 1.60 & 

- .clr z CII 

. . ..................................................................................................................................... P1 -- ........................... * 3 n 

0 .- 

- .I Q. 0 g ..... ......--... CII 

0 = s 

W 

0.8 

0.7 - 

- r--. ... "----............................................ 

0 0.5 I 1.5 2 2.5 

Disharge time (hr) 
Figure 5 

I 

IN Plot for Nickel Electrode 

on Upper and Lower Voltage Plateaus 

0.8 0.9 1 1 .I 1.2 

Voltage vs. Hydrogen (50 psia) 
4 Figure 6 

e of Barrier Layer 

in Producing Lower Vo 

in Nickel Electrode 

Electrolyte 

kt site A: 

O2 + 2H20 + 4e- = 40H- 
Page intentionally left blank 
b 

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52 

EAGLE-PICHER INDUSTRIES INC. 

k TECHNOLOGIES DIVISION 8 

R 

r3 POWER SYSTEMS DEPARTMENT 

2. 

X 

9 

COLORADO SPRINGS, CO 80911 

3 

% 

b" 

FABRICATION OF NICKEL 

ELECTRODES USING ETHANOL 

SOLUTIONS 

DAVID F. PICKETT, JR. 
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S B - STUDIES OF AQUEOUS El PROCESSES AT BELL LABS. 

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- CONTRACTUAL STUDIES 

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DEVELOPMENT HISTORY 

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7 -WRIGHT LABORATORIES BENCH DEVELOPMENT 

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MAURER, MACARTHUR, McHENRY, BEAUCHAMP, 

O'SULLIVAN, WILLIS 

ACTED AS ADVISOR TO USAF AND INDUSTRY 

- IN-HOUSE EFFORTS AT WLIAFAPL (AQUEOUS, ETHANOL) 

LANDER, PICKETT 

- INDUSTRY(C0NTRACTS THROUGH USAFI WL) 1, 

SPECTROLAB, EAGLE-PICHER 

- NASA LeRC 

CONTRACT WITH HUGHES 

MANUFACTURING TECHNOLOGY (USAF) 

- EAGLE-PICHER 
r, 

'0 'a 

'4 

2 b 

$2 

b FLIGHT QUALIFICATION 9 1 

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NICKEL HYDROGEN PROGRAMS AT 

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3 f: WHY ETHANOL? \ 

n - 

CAN BE CARRIED OUT AT LOWER TEMPERATURES 

THAN AQUEOUS SOLUTIONS 

- FORMS AZEOTROPE WITH WATER 

- REDUCES SURFACE TENSION 

RETARDS NICKEL ATTACK ON SUBSTRATE WITHOUT 

FURTHER PASSlVATlON OF SINTER 

- USED TO IMPREGNATE NICKEL PLATED CARBON 

FIBERS 

REDUCING AGENT DOES NOT BUILD UP IN SOLUTION I 

ENHANCES PRECIPITATION IN PORES OF SINTER 

PROVEN TO GIVE LONG LIFE ELECTRODE 

- LIM AND SMITHRICK REPORT >20,000 LEO CYCLES 

AT 80% DOD 

- IN-HOUSE STUDIES AT HSC 
4 

8 '4 

5 

2 A 

WHAT MAKES THE BEST ELECTRODE - 

2 AQUEOUS OR ALCOHOL PROCESS? I 

4 

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i3 

3- "IT'S NOT WHAT YOU LINE UP WITH, IT'S WHO YOU LINE UP WITHwP DARRAL ROYAL 

B 

.a0 -THE USAF,NASA, HUGHES AND EAGLE-PICHER (COLO SPGS) ARE THE 

"FOUR HORSEMEN" WHO MADE THE ALCOHOL PROCESS 

SUCCESSFUL 

- THE BELL LABS HIGH TEMPERATURE PROCESS ALSO MAKES I 

VI 

a 

f 

EXCELLENT ELECTRODES 

BELL LABS., TYCO, GENERAL ELECTRIC, AND EAGLE-PICHER 

(JOPLIN) PARTICIPATED IN ITS DEVELOPMENT 

3 

APPLICATION AND ECONOMICS, NOT ONLY CYCLE LlFE PERFORMANCE, 

P ARE ALSO KEY IN DECIDING WHICH IS THE BEST ELECTRODE FOR AN 

9 APPLICATION. 

2 

1 

a - GEO ONLY REQUIRES -2500 CYCLES FOR 'MISSION I 

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P 

2. - PLANETARY MISSIONS USUALLY REQUIRE LONG STAND LIFE AND 

2 

2' 

c, 

MINIMAL CYCLE LlFE 

$ 

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- LONGEST LEO CYCLE LlFE HAS BEEN DEMONSTRATED USING 26% 

k? 

8 

KOH AND ALCOHOL ELECTRODES. 

6' 

3 QUALITY IS NO ACCIDENT - THE BEST ELECTRODE WILL BE MADE BY 

THE ORGANIZATION WHO HAS THE COMMITMENT AND DEDICATION TO 

MAKE IT - BE IT AQUEOUS OR ALCOHOL BASED. 
CURRENT USES OF ELECTRODES 

2 

.a" 

!2 IMPREGNATED WITH ETHANOL SOLUTIONS 

9 HSC NICKEL HYDROGEN CELLS 3 0 

- GEO COMMUNICATIONS SATELLITES 

EAGLE-PICHER SUPER NiCd CELLS 

- GEO 

- LEO 

- SPACE PROBES 

- SPACE STATION EMERGENCY LIGHTING 

MAINTENANCE FREE AIRCRAFT BATTERIES 

- BOMBERS \ 

- FIGHTERS 

- MILITARY TRANSPORT 

- HELICOPTERS 

- OTHER 
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F S E 

4 

LIGHTWEIGHT NICKEL ELECTRODES 

- HSC 

- NASA LEWIS RESEARCH CENTER 

SPECIALTY APPLICATIONS 

- EAGLE-PICHER POWER SYSTEMS BEPT. 

MODELING OF IMPREGNATION 

PROCESS 

- ORDIUNIV. OF S. CAROLINA 
Effects of Cycling Conditions of Active Material 

from Discharged Ni Positive Plates Studied by 

c(n 

Q\ 

Y Inelastic Neutron Scattering Spectroscopy 

3 Juergen Eckert t 

trl F 

2 

Los Alamos National Laboratory I 

a I 

?3 

I 

9 

and 

5 

'T X 

Ravi Varma, Lisa Diebolt and Margaret Reid E 

U 

R 

Fa 8. 

3 
52 Objectives 

3 

Ct 

g 

3 

$ 

$' Identify atomic-level signatures of electrochemical 

activity of the active material on the Ni positive plate 

of Ni-H, batteries. 

Relate findings to cycling conditions and histories 

Develop INS spectroscopy as a non-destructive 

testing technique for the evaluation of Ni-positive 

plates of Ni-H, batteries. 
ChargelDischarge of 

(a,P)-Ni(OH)2 1 (y,P)-NiOOH Couples 

Charging I Discharging 

4- oxygen - r-- n1cr.t 

Oxygrn +- Hydrogrn 

I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
4 

a 2 Fundamentals of Vibrational Spectroscopy by 

"nelastic b Neutron Scattering 8 

neutrons are scattered by the atomic nuclei and not 

the electrons ( as are photons ) 

- scattering cross-sections a nuclear property 

- H scatters neutrons >I0 times more strongly than other 

atoms 

lJ, 

4 

7 absorption cross-sections for neutrons are very low: 

- probe the bulk of the sample 

- in-situ methods are easy(no windows required) 

3 

fi all vibrational modes are observable 

2 

B o - intensities are weighed by nuclear kross-sections:INS : 2 

b 

(D spectra are dominated by modes involving large * 

3 $displacements of H atoms. 

- intensities are readily quantifiable and are proportional to the 

number of scatterers. 

Ids Alm~ros Neutron Seierrce Cerrfer 
0 Incoherent scattering cross section 

Coherent scattering cross section 

@ Absorption cross section 

1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
s '4 

INS Vibrational Spectroscopy !2 

b 

8 $ 

W 

d technique is well suited for application to battery 

P 0 3 material Gi 

's. 

4 - bulk probe 

- sensitivity to protons (H) 

& 

' experiments are carried out at the Lujan Center of 

4 

Y LANL 

- 5 - log samples from battery plates 

3 

- FDS instrument;.AE = 50 - 4000 cm-I 

- 12-24 hrs. data collection time 

- T=15K ( 

Los A lartros Neulrorr Science Cerrter 
r EXTERIOR 

SHIELDING 

I997 NASA Aerospace Battety Workshop Nickel Electrode Design Focused Session 
Assignment of P-Ni(OH), vibrational bands 

Wave Number (cm-') 

I997 NASA Aerospace Battev Workshop 

INS (cm-1) IR, Raman Assignment 

(cm-1) 

Nickel Electrode Design Focused Session .. 
- 

- Hydrogen disorder in brucite structures 

1'997 NASA Aerospace Battely Workshop Nickel Electrode Design Focused Session 
Raman Scattering Spectra of Ni electrode 

materials - - 

B. C. Comilsen and collaborators 

Fig. 1. Raman spectra of nickel electrode active mass and model compounds, (a) Charged 

active mass; (b) discharged ar active mass; (c) recrystallized /3-Ni(OH)2; (d) first precipitate &phase. 

Raman spectra of: (a) discharged active mass (ID no. 1W9); (b) 'phase-X 

.- 

I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
KQH conBPI 21 % 

BP4 31 46' 

BPlO 23.5% 

400 600, 800 1000 

Wave Number ( cm-1) 

1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
Vibrational Modes of Hydration Water 

Internal modes 

Vl' Vsym) '3' Vasym) 

Librations - 

R)(waa) Rt (twist) Rr(r-) 

frequency ranges ( cm-1) 

stretching modes ( v ) 

bending modes ( 6 ) 

librations ( R ) 

translatory modes ( T ) 

1997 NASA Aerospace Battev Workshop -578- Nickel Electrode Design Focused Session 
INS Spectra of Reference Compounds - - 

F. Fillaux et al., Physica B 213&214,637 (1995) 

1997 NASA Aerospace Battery Workshop 

4 

I 

1 000 2000 30b0 4600 

Energy transfer (an-') 

Nickel Electrode Design Focused Session 
Wave Number (cm-') 

1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
s '4 INS results 2 

discharged materials are mainly P-Ni(OH), 

changes in the Ni-0 stretching and bending regions: 

- a decreases from 3.1 3 A (P-Ni(OH), ) to 2.89A (P-NiOOH ) 

- distortion of NiO, octahedron 

- frequency shifts and band splittings result 

water librations above - 500 cm-I 

- vacancies may allow formation of Ni(H,O) 

protons in 0-Hm-0 hydrogen bonds: P-NiOOH 
Structural Models for Hydrogen in - - NiOOH 

and bound H,O 

I997 NASA Aerospace Battery Workshop -582- Nickel Electrode Design Focused Session 
C Greaves el al. / Structure of the mixed conductor Ni203H 

FIGURE 2 

Schematic representation of the structure of 

Ni203H 

FIGURE 3 

Fourier section, y = 0, based on observed intensities and calculated phases 

1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
'0 'n 

Conclusions F2 

B 

a 

(D 

'a 9 - (1)lrreversible formation of NiOOH; scales with number 

i? 

F 0 

of cycles 

- (2) additional protons are bound in the lattice to form Ni- 

(H,O) complexes; increases with KOH concentration in 

the cell. 

- (3)These processes occur only in the outermost layers 

of the plate material but lead to the failure of the battery 

cells. 

- 

Los Alanros Neutrorr Science Center 
Support from the NASA Lewis Research 

Center (NASA grant NAS-3-119) 

DOE/OBES for use of facilities at the 

Manuel Lujan Jr. Neutron Scattering 

Center of Los Alamos National 

Laboratory 

1997 NASA Aerospace Battery Workshop -585- Nickel Electrode Design Focused Session 

-- - 
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- V STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT ON 

NICKEL HYDROGEN 'CELLS 

Y. BORTHOMIEU 

SAFT Defense and Space Division 

POITIERS FRANCE 

File: 0:\s'17\\r~lUXSlGOG-97.1W 

N. SAC-EPEE 

Laboratoire de RCactivit6 et de Chimie des Solides du CNRS 

AMIENS FRANCE 

T. JAMIN I 

CNES 

TOULOUSE FRANCE 

1997 NASA Aerospace Battery Workshop 

Huntsville Hilton, AL 

November 18-20,1997 

1997 NASA Acrospncc B:lllcry Workshop, Hitntsvillc I-lilton, AL 

Novc~lll~cr 18-20, 1997 
h STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Division ON NiHz CELLS 

OUTLINE 

1- NiCd CELL SECOND PLATEAU 

b 

8 $' 2- EFFECTS OF HYDROGEN PRESSURE AND NiHz CELL SECOND PLATEAU 

File: ~:\sLI\~~~I~D\sI~o~-~~.I~oc 1997 NASA Aerospace Battery Workshop, Huntsville Hilton, AL 

Novc~nbcr 18-20, 1997 
,I 

V mEma STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

THE SECOND PLATEAU 

WHAT IS IT ?\ 

1997 NASA Aerospace Battery Workshop, Huntsville Hilton, AL 

November 18-20, 1 997 
~15~~t4t102,, I"" 116 IIII,,, J 

I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
- v STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battety Group 

Defense and Space Division ON NiHz CELLS 

INTRODUCTION 

SECOND PLATEAU IS GENERALLY OBSERVED IN ALL NICKEL COUPLES : 

NiCd, NiMH and NiH2 

BUT IS EXACERBATED IN NiH2 CELLS 

OUR APPROACH : 

a MANY OF ELECTROCHEMISTRY STUDIES HAVE BEEN PERFORMED 

(ZIMMERMAN AND BARNARD) 

File: o:\snwoalnsr ~o6.97.I)C~C 1997 NASA Aerospace Battery Workshop, H~lntsville Hilton, AL 

Nove~nber 18-20, 1397 
v , 

m2Rm STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

i$' Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

T 2 INTRODUCTION E 

4 DIAGRAM OF THE STUDY 

ACTIVE MATERIAL BEHAVIOR NICKEL HYDROGEN CELL 

BEST KNOWLEDGE OF THE SECOND PLATEAU PHENOMENON 

File: ~:K~~U~~I~I)\SIGOG-~~.I)OC 1007 NASA Aerospmc Baltcry Worksltop, Huntsville Hillon, AL 

November 18-20, 1997 
V Emm STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and lnduslrial Battery Group 

Defense and Space Division ON NiH2 CELLS -- 

1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

1.1 EFFECT OF THE Y PHASE 

1.2 ROLE OF THE INTERFACE 

1.3 STUDY OF y/P(II) TRANSITION 

File: o:\s'l\wolto\sIGoG.97.t~Oc 1997 NASA Aerospace Battery Workshop, Huntsville Hiltort, AL 

Novc~nber 18-20, 1997 
h STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battefy Group 

Defense and Space Division ON NiH2 CELLS 

1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

1.1 EFFECT OF THE y PHASE 

PREPARATION OF VARIOUS OXIDIZED PHASES FROM P(II) 

LEVEL OF OXIDATION CONTROLLED BY : 

- THE QUANTITY OF KOH 

- THE TEMPERATURE (55 TO 80 OC) 

I 

OXIDATION LEVEL BETWEEN 2.8 AND 3.6 

1007 NASA AC~OS~IICC B:lltcry Workshop, I-111ntsvillc Milton, AL 

Novc~nbcr 18-20, 1997 
4 v 

w m STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

!? Advanced and Industrial Battery Group 

2, Defense and Space Division ON NiH2 CELLS 

9 

T B 

% 1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

43 

1.1 EFFECT OF THE y PHASE 

X-RAY DIFFRACTION PATTERN STUDY 

ELECTROCHEMICAL STUDY (BOILER-PLATES) 

1997 NASA Aerospace Battery Workshop, Huntsville Hillon, AL 

Noven~bcr 18-20, 1997 
1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
0 0.2 0.4 0.6 -0.8 1 

Relative amount of 

y(~~~) phase in the sample 

0 0.2 0.4 0.6 0.8 1 

Relative capacity on the second voltage plateau 

1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
m 

v 

STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Division ON NiHz CELLS 

5 8 E 1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

G 

(k, 

\O 

PO 

I-IOW TO EXPLAIN THE y PI-IASE DURING THE CHARGE ? 

2 0 

8 WI-IICFI PARAMETERS DO DETERMINE THE y PHASE FORMATION 

P - FROM AN ELECTROCHEMICAL POINT OF VIEW ? 

2 

File: o:\Sl\WOlt~S1606-97.1)0C 1997 NASA Aerospace Battery Workshop, Huntsville Iiilton, AL 

Novcm\>cr 18-20, 1 997 
4 

U 

5 

2 

b '1 

a 

% v 

8 (D Emm STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

2 Advanced and lnduslrial Battery Group 

Defense and Space Division 2 

ON NiHz CELLS 

4. 

F 8 

g 1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

4 

1.2 EFFECT OF THE COLLECTOR 

1997 NASA Aerospace Bi~ttery Workshop, Huntsville Hilloo, AL 

Novcoiltc~. 18-20, 1997 

IEC ELECTRODE ON NICKEL 

SINTER 

PHYSICAL IMPREGNATION (PI) OF A 

NICKEL PASTE ON A NICKEL FOAM 

, WITH VARIOUS PORE SIZES I 

I 
PI electrode 

with small pores 

24 26 

time (h) 

24 26 28 30 

time (h) 

Pi e!ectode 1.5 

with large pores 

E z1 

9 

- C Q 

>" 0.5 

0 

20 22 24 26 28 30 Time (R) 

I997 NASA Aerospace Battery Workshop -600- Nickel Electrode Design Focused Session 
X-RAY DIFFRACTION PATTERNS AT THE 'END OF CHARGE 

ECI . 
\ 

\O 

\O 

'4 

s 

52 

i? 

a '\ 

G 

f: 

a dim STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

B Advanced and Industrial Battery Group 

... 

Q Defense and Space Division m 

ON NiH2 CELLS 

1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

L 

4 ROLE OF THE COLLECTOR 

File: o:\snwollI)\s1606-97.1)0~ 

FOAM 

1997 NASA Aerospace Battery Workshop, Huntsville Hilion, At, 

Novc~nbcr 18-20, 1997 
m!ka STUDy OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

' POOR CONTACT BETWEEN COLLECTOR 

AND ACTIVE MATERIAL 

I 

t 

OVERCHARGE LOCALIZED AT THE INTERFACE 

I FORMATION OF y PHASE I 

1997 NASA Aerospace Battery Workshop, I-Iuntsvillo Hilton, AL 

Nove~i~ber 18-20, 1997 
g 

3 & STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

b Advanced and Industrial Battely Group f? Defense and Space Division ON NiH2 CELLS 

9 

7 

i? 1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

% 

1 .3 .)I/P (11) TRANSITION 

File: o:\s'nwowmsl~o6-c,7.l,oC 

BODE DIAGRAM : 

1997 NASA Acrospacc Battery Worksllop, fintsville Ililton, AL 

Nove~nbcr 18-20, 1997 
1997 NASA Aerospace 
4 

V3 

V3 

'4 

s 

2 

b 

2 'I 

% V 

1 

B 

STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

2 Advanced and Industrial Battery Group 

Defense and Space D~v~sion ON NiH2 CELLS 3 

F B i? 

1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

% 

cS\ 

0 

?' 

DIRECT TRANSITION y 

Pile: ~:\~II~~~oI~~\sI~~G-~~.Do~ 

NIGH RESOLUTION TEM ANALYSES SI-10 W : 

P(II)ucr EXHIBITS SAME MOSAIC STRUCTURE AS Y 

PLATELETS ARE MAINTAINED FROM y TO p(I1) 

1 TI-IE REACTION ISPSEUDOMORPHOUS 

1997 NASA Aerospace Battery Workshop, Hon(svil1e Hillon, AL 

Novcn~bc~. 18-20, 1 997 Page- 1 -i 
1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
I 

I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
-,-..- . -- .- . .. - . . . 

Next 

C 

H 

A textural memory 

R 

G 

E Y (W B(m) 

1997 NASA Aerospace Battely Workshop Nickel Electrode Design Focused Session 
F 8 

$ 1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

4 

File: o:\snwoltrl\sl60~-97,1)oc 

MEB, SAED, XRD TECHNICS CONFIRM THIS TRANSITION 

A RECHARGE ON P(II), , YIELDS IMMEDIATELY y PHASE 

1997 NASA Aerospace Battcry Workshop, Huntsville I-Iilton, AL 

Novcmbcr 18-20, 1997 
CHARGE begins at i .SGV 

(no 0.8 V plateau) 

- 1.4 I...,. ..,...- 

- N 

- 

I 

a 

..:(111) at the beginnino .# of 

charge: psx.,! --> *; 

I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
LA 

m 

2 V 3 

cp Emm STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

2 Advanced and Industrial Battery Group 

Defense and Space Div~sion ON NiH2 CELLS 

% 

3 . 

P a 1- STUDY OF SECOND PLATEAU APPEARANCE IN NiCd 

B 

SUMMARY ON ACTIVE MATERIAL BEHAVIOR : 

EZ 

2 THE SECOND PLATEAU CAN BE INFLUENCED BY THE COLLECTOR INTERFACE n 

3 I 

1007 NASA Acrospilcc Ballcry Workshop, Ntt~ltsvillc Hilloo, AL 

November 18-20, 1997 
'r 

'Cl 

3 

2 

52 

% I\ 

-i $ 

S 

STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group i? Defense and Space D~vision 

ON NiHz CELLS 

3' 

3: .? 2- EFFECTS OF HYDROGEN PRESSURE 

E 

G 

2.1 STORAGE IN KOH 

1997 NASA Aerospace Battery Worksliop, Huntsville Hilton, AL 

Nove~nbcr 18-20, 1997 
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2 b 

2 

b 

(D Y 

V 

I1 8" 

n 

(D m!Em STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

& Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

3 

3: a 2- EFFECTS OF HYDROGEN PRESSURE $ 

4 

2.1 STORAGE IN KOH 

P(I1) WITI-I SMALL PLATELETS (30 A * 200 A) 

TREATMENT AT 20 OC (AND 50°C) UNDER KOH 8N 

EVALUATION OF THE PARTICLE GROWTH 

2 400 V 

I Time (Days) I 

1007 NASA Acrosp:~cc B:~llcry Workshop, I~lunlsvillc I-lilloa, AL 

Novclnbcr 18-20, 1997 
'\ 

& STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space D~vision ON NiHz CELLS 

2- EFFECTS OF HYDROGEN PRESSURE 

File: o:\snwo1c1)\~1606-97.l,oC 

DISSOLVING OF THINNEST PLATELETS AND ' 

RECRYSTALLIZATION ON THE THICKEST ONE 

1 

PARTICLE SIZES INCREASE 

BUT STABILIZE 

AFTER 24 HOURS 

1997 NASA Aerospace Bi~tlc~y Worksltop, FIunlsville 14il1011, AL 

Novc~t\bcr 18-20, 1397 
I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
v 

2 - STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

X Defense and Space Division ON NiHz CELLS 

\. 

F & 2- EFFECTS OF IIYDROGEN PRESSURE 

* 

P(II1) Is OBTA~NED WITH CHEMICAL OXIDATION OF p(I1) : (u 1 3 00 A) 

TREATMENT 80 BARS (1 100 PSI) OF Hz IN KO13 AT 60°C 

EVOLUTION OF PARTICLE SIZE VERSUS TIME 

3000 - 

\ 

1000 I I I 

0 50 100 150 200 

lime (hours) 

1997 NASA Aerospace Battery Workshop, Huntsville Hilton, AL 

Novc~nber 18-20, 1997 
+ 

\O 

2 

5 

52 

b 

8 I $ V 

1 Enm STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

3' Advanced and Industrial Battery Group 

X Defense and Space D~vis~on ON NiH2 CELLS 

3" 

$ 2- EFFECTS OF HYDROGEN PRESSURE 

5 

4 

AFTER 150 HOURS IN H2 : 

Pile: o:\SI7\VOltl>\slGoe.97.l>OC 1007 NASA Acrosp;lce B;~tlcry Worksl~op, I-Iu~~lsvillc I-lil~o~i, AL 

Novcnlb?r 18-20, 1997 
$ 1 

2 v 

3 - STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

g. Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

2- EFFECTS OF HYDROGEN PRESSURE 

File: o:\s'nwo~aslc,o6.97 DOC 1997 NASA Aerospace Eklllery Worksl~op, I-Innlsville I.lillon, AL 

Novc~nbcr 18-20, 1997 

> ON RESISTOR 
dm STUDY OR PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

2- EFFECTS OF HYDROGEN PRESSURE 

- 400 

s. 

a 200 

5 10 15 20 25 

Hydrogen Ressure (bars) 

QToIV +TOOV 

INCREASE OF SECOND PLATEAU WITH THE PRECEIARGE LEVEL WHEN DISCHARGED TO 0 V 

File: O:\sI'\\VOl~D\S1606-97.I)OC 1997 NASA Aerospace Battery Worksliop, Ht~ntsville Hillon, AL 

No\relnber 18-20, 1997 
0 10 20 30 40 50 

Pression de fin de charge (bars) 

I997 NASA Aerospace Battery Workhop -62 1 - Nickel Electrode Design Focused Session 
-2. 

0.5 0.6 0.7. 0.8 0.9 1 1.1 

Potentiel (V) 

Evolution du potentiel du second plateau 

en fonction de la pression atteinte en fin de charga 

0 20 40 60 80 100 

Pression de fin de charge (bars) 

I997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
ii 

(D Advanced and Industrial Baltery Group 

Defense and Space Division 

3 

ON NiHz CELLS 

2- EFFECTS OF HYDROGEN PRESSURE 

$ 

% SUMMARY ON HYDROGEN EFFECTS 

File: O:\S71\VORD\S1606-97.DOC 1997 NASA Aerospace Battery Worksl~op, Huntsville Hilton, AL 

Novcmber 18-20, 1997 
dm STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and lndustr~al Battery Group 

Defense, and Space Division ON NiHz CELLS 

CONCLUSION 

1907 NASA Acrospacc Battery Workshop, Huntsville Milton, AL 

Novc~ubcr 18-20, 1997 
m 

v STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

Advanced and Industrial Battery Group 

Defense and Space Dlvision ON NiHz CELLS 

CONCLUSION 

RECOMMENDATIONS FOR NIH~ - CELL MANAGEMENT 

IN ORDER TO DIMINISH SECOND PLATEAU APPEARANCE : 

AVOID OVERCHARGE AND y FORMATION : LIMIT THE RECHARGE RATIO 

USE POSITIVE PRECEIARGE INSTEAD NEGATIVE '. I 

AVOID DEEP DISCHARGE (BETWEEN 0.5 AND 0 V) UNDER RESISTOR 

DURING RECONDITIONNING 

File: o:\s'nwolu~\s 1606.97.1)oc 1997 NASA Aerospace Batlcry Workshop, Huntsville Hilloa, AL 

Novanber 18-20, 1997 
b-. 

'n 

'n 

V 

5 

2 

k 'I 

a 

d v 

8 (D m STUDY OF PARAMETERS CAUSING SECOND PLATEAU EFFECT 

R Advanced and Industrial Battery Group 

Defense and Space Division ON NiH2 CELLS 

J 

ACKNOWLEDGEMENTS 

A.DELAHAYE AND J.M TAR4SCON FOR FRUIFUL DISCIJSSIONS 

T. JAMIN FROM CNES FOR FINANCIAL SUPPORT 

File: ~:\sI~~oI~~sIG~~-~~.I,~c 1997 NASA Acrospacc Battcry Workshop, H~~ntsville Hilton, AL 

No\tember 18-20, 1997 
'\ 

PERFORMANCE COMPARISON BETWEEN 

NiH, DRY SINTER AND SLURRY ELECTRODE CELLS 

J. D. Armantrout 

Presented at 

1997 NASA Battery Workshop 

Focused Session - Design of the Nickel Electrode 

November 19,1997 I 

I 
b 

9 ,\ 

5 

% Background 2 

2 

e 

LAUNCHED IN 1990 WITH A 7-YEAR BATTERY DESIGN LIFE (5-YEAR MISSION). 

- 88 Ah NiH, CELLS, DRY SINTER PROCESS PLATES, HYDROGEN PRECHARGE. 

- NOMINAL 12 PERCENT BATTERY DEPTH OF DISHARGE (DOD). 

BATTERY DATA THRU 1995 SHOWED CAPACITY LOSSES OF 4.5 AhNEAR. 

BASELINE BATTERY REFURBISHMENT IN 2002 ON SERVICING MISSION 4. 

- UPGRADE TO SLURRY PROCESS PLATE; INCORPORATE POSITIVE PRECHARGE. 

- COULD REPLACE BATTERIES IN 1999 BASED ON CAPACITY TREND ANALYSIS. 

NEED TO KNOW OPERATING V/T CHARGE LEVEL (SLOPE AND INTERCEPT) 

AND HEAT DISSIPATION OF NEW CELL DESIGN. I 

I 
P '\ 2 

6 

a0 

(D 

HST Refurbishment Cell Design Summary 

tu 

2 

9 

q "Pineapple Slice" Cell Design with the Following Components: 

8 

$ 

% 48 Slurry Process Nickel Positive Electrodes 0.035 in. thick 

48 Platinum Negative Electrodes (0.006 in. thick) 

Back-to-back Electrode Stack Configuration 

Zirconium Oxide Cloth Separators and Gas Screens 

Polysulfone End Plates, Core and Retaining Nut 

Belleville and Whiteley Washers 

lnconel 718 Pressure Vessel 0.040 in. thick 

Zirconium Oxide Pressure Vessel Wall Wick 

0. Inconel718 Terminal Bosses and Weld Ring t 

Injection Molded Nylon Terminal Seals 

, "Rabbit Ear" Terminal Configuration 

Positive Precharge 
Offset Non-Opposing vs. Rabbit Ear Cell 

OFFSET NON-OPPOSING 

1- 3-97 REF 1 

RQBBIT ERR 

4-30 REF 1 
HST NICKEL-HYDROGEN BATTERY OVERVIEW 

CONTRACT MOD 593 DATED 2 JUNE 1987 DIRECTED LMSC TO DESIGN, 

DEVELOP AND DELIVER NICKEL-HYDROGEN (NiH,) BATTERY MODULES TO 

REPLACE NiCd BATTERIES FOR THE HUBBLE SPACE TELESCOPE (HST) LOW 

EARTH ORBIT (LEO) MISSION PER NAS 8-32697. 

A NiH, MODULE WITH THREE ELECTRICALLY DISTINCT BATTERIES IN A 

COMMON MECHANICAL STRUCTURE WAS DESIGNED AS AN ORBITAL 

REPLACEMENT UNIT (ORU). TWO MODULES WERE INSTALLED ON BAY 

DOORS 2 AND 3 OF THE HST SUPPORT SYSTEM MODULE (SSM). 

0 TWO MODULES MAKE UP THE FULL NiH, BATTERY COMPLEMENT OF SIX 

BATTERIES, EACH RATED AT 88Ah TO 26.4 VDC PER CELL AT A 15 AMP 

DISCHARGE RATE AT 0°C AND HAVING'A MINIMUM ORBITAL LIFE OF 5 YEARS. 

ONE STRUCTURAL ENGINEERING MODULE (s~M) AND FIVE MODULES WERE 

MANUFACTURED AND DELIVERED BY EAGLE PICHER INDUSTRIES (EPI). 
I\ 

d .a" 

8 

(D 

s 

HST NiH, BATTERY RECONDITIONING CAPACITY 

$ 

$ THE HST NICKEL HYDROGEN BATTERIES BUILT FROM 88 Ah CELLS b" MANUFACTURED WITH DRY SINTER PROCESS ELECTRODES HAVE 

EXCEEDED THEIR 7-YEAR DESIGN LIFE (5-YEAR MISSION). 

- PERIODIC BATTERY CAPACITY MEASUREMENTS SHOW IMPROVED PERFORMANCE 

SINCE SWITCHING FROM PRIMARY TO SECONDARY HEATERS IN NOVEMBER 1995. 

BATTERY 

POSITION 1990 - 1992 

1 94.2 

2 88.5 

3 85.6 

4 93.4 

5 94.0 

6 88.3 

BATTERY CAPACITY (Ah) 

1994 1995 - 1996 - 1997 

80.4 74.5 83.7 

79.3 72.4 78.5 

73.5 

\ 

. 74.7 74.4 

83.7 

78.3 73.7 76.4 

74.6 70.5 75.3 
Figure 1. HST Battery Capacity Trend Data 

-t- Battery 2 

+ Battery 3 

. . - - - - - - . . . . . 

Post-test Capacity 

Post-test Capacity 

--Ilt- Battery 4 Post-test Capacity 

+ Battery 5 Post-test Capacity 

+ Battery 6 Post-test Capacity 

- Least Squares Fit For Orbit 

- Revised Fit for Orbit 

Lp 

i I I 

-___S______ 

0 12 24 36 48 60 72 84 96 108 120 132 144 

Time From Launch (Months) 
1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
1997 NASA Aerospace Battery Workshop Nickel Electrode Design Focused Session 
HST CHARACTERIZATION TEST PROCEDURE 

BATTERY CELLS BUILT WITH SLURRY PROCESS ELECTRODES TESTED TO 

CHARACTERIZE PERFORMANCE FOR EXPECTED OPERATING CONDITIONS. 

- TESTED VARIOUS CHARGE VOLTAGE LIMITS OVER 23°F TO 50°F RANGE. 

TESTING PERFORMED UNDER ISOTHERMAL CONDITIONS THAT ALLOW FIRST 

LAW CALCULATION OF HEAT DISSIPATION AND RECHARGE RATIO. 

- CHARGING PERFORMED AT 14 AMPS TO VOLTAGE LIMIT FOLLOWED BY 2 AMP 

TRICKLE CHARGE FOR DURATION OF SUN PERIOD. 

- 9.6 Ah DOD ACHIEVED USING 12 AMP AND 24 AMP DISCHARGES DURING ECLIPSE. 

PERIODIC CAPACITY MEASUREMENTS MADE TO DETERMINE EFFECTS OF 

VARIOUS TEST CONDITIONS ON PERFORMANCE. 

CELLS MANUFACTURED WITH DRY SINTER PROCESS ELECTRODES , 

INCLUDED IN THE TEST FOR COMPARISON WITH ORIGINAL DESIGN. 

OPERATING CAPACITY MEASUREMENTS TO 1.2 AND 1 .I VOLTS PER CELL 

MADE AT 18 AMP AND/OR 40 AMP RATES. 
HST CHARACTERIZATION TEST RESULTS 

CELLS WlTH SLURRY ELECTRODES HAVE SLIGHTLY HIGHER ORBITAL 

AVERAGE THERMAL DISSIPATION THAN DRY SINTER CELLS. 

- DATA SUMMARIZED IN TABLE 1 SHOW PERFORMANCE FOR EACH CELL TYPE. 

- DATA SUMMARIZED IN TABLE 2 AND PRESENTED IN TABLE 3 SHOW CORRELATION 

OF HEAT DISSIPATIONS AND RECHARGE RATIOS. 

- DATA ILLUSTRATED IN FIGURES 4 THROUGH 10 SUMMARIZE TEST RESULTS. 

SLURRY PROCESS ELECTRODE CELLS HAVE LOWER RELATIVE CAPACITY 

AT LOWER TEMPERATURES (FIGURE 11 AND TABLE 1). 

- VOLTAGE DATA IN FIGURES 12 THROUGH 15 SHOW LOWER RELATIVE DISCHARGE 

VOLTAGE PLATEAU FOR SLURRY CELLS VERSUS DRY SINTER CELLS. 

- RECONDITIONING CELL VOLTAGE DRAINS ARE SHOWN IN FIGURES 16 AND 17. 

- CYCLING PERFORMANCE DATA IN FIGURES 18 THROUGH 21 ILLUSTRATE AN 

INDEPENDENT TEST RUN UNDER ISOTHERMAL FONDITIONS. 

THERE WAS A DRAMATIC EFFECT OF CHARGE CUTOFF VOLTAGE ON PIEAT 

DISSIPATION (FIGURES 5 AND 6). 

CAPACITY INCREASED ONLY SLIGHTLY WlTH HIGHER CHARGE CUTOFF 

VOLTAGES (FIGURES 7 THROUGH 10). 
Table 1. HST CHARACTERIZATION TEST RESULTS 

................................................................... I VT2.5 1 11 1.93% 1 120.20% / 124.96%1 128.48% 

Note: Recharge Ratios for Dry S inter Cells not measured. 

S lumy Recharge Ratio 

Dissipation [watts) 

. 8. 

- .............. Slurry Cells Dry Sinter Cells 

23°F i 32°F i 41°F i 50°F 

YT5 

VT4.5 

VT4 

YT3.5 ........................... VT3 

VT2.5 

100.88% 

101.79% 

I 03.02% 

105.30% 

107.65% 

100.55% .......................................................... 101.76% 

I 04.77% ........................................................... 108.59% ........................................................... I1 3.14% 

23°F i 32°F i 41°F i 50°F 

11.8 

14.0 

16.4 

20.9 

25.9 

34.7 

23°F i 32°F i 41°F i 50°F 

101.81% ......................................................... I 05.51 % 

109.50% 

11 5.83% 

10.2 

121 

14.5 

19.2 

24.1 

33.2 

1 04.63% 

109.74% 

I1 6.56% 

10.8 

13.3 

19.3 ..................................................................... 26.9 

36.1 

50.7 

9.1 

11.5 

17.4 

24.8 

34.0 

48.8 

13.1 

20.4 

28.5 ....................................... 41.4 

60.0 

............................................................................................................................. 

........................................................................................................................... I 8.4 

28.5 

42.2 ............................................................................................................................. 66.2 

11.3 

18.5 ......................................................... 26.5 .....................................(....,......,...... 39.1 

57.6 

I 

C. 

16.5 

26.6 

40.3 

64.2 
'I 

Table 1 (Cont). HST CHARACTERIZATION TEST RESULTS 

Operating Capacity (Ah) to 1 -2 Yolts 

Slurry Cells Dry S inter Cells 

Operating Capacity (Ah) to 1 -1 Yolts 

............... 

. Slurry Cells Dry Sinter Cells $ 23°F i 32°F i 41°F i 50°F , 23°F i 32°F i 41°F i 50°F 

23°F i 32°F 41°F i 50°F 

i? 

5 

(D 

C: 

a 

R 

P 3. 

3 

9 0 

2 

46.1 9 

46.79 

47.23 

47.64 

48.05 

47.02 

23°F i 32°F i 4I0F i 50°F 

70.56 

70.87 

71.29 

72.42 

73.45 .............................................................................................................. 70.72 

X 

EL 

3 

GI g. 

3 

VT5 

VT4.5 .............. VT4 

VT3.5 

VT3 

VT2.5 

47.60 

50.43 

51.71 

52.1 6 

52.68 

53.1 9 

69.35 

70.51 

72.55 

74.64 

75.23 

72.50 

50.25 

54.1 I 

55.85 .................................................................... 57.03 

58.26 

81.53 

82.29 

82.64 

82.84 

83.1 5 

82.6 

............................................................................................................... 

............................................................................................................ 51.77 ............................................................................................................... 55.77 

59.12 

61.45 

62.05 

69.80 

70.47 I............l............................. 

71.31 

70.46 

94.1 2 

95.31' 

93.90 

93.61 

94.63 

90.02 

58.87 

62.59 

66.30 

68.1 3 

81.35 

84.14 

84.97 

85.55 

85.90 

85.1 9 

85.73 

86.1 9 

88.26 

90.41 

90.72 

86.93 

79.68 

82.88 

84.08 

85.12 

84.98 

.............................................................................................................................. 

.........................................................................

73.91 ............................................................................................................................... 77.44 ............................................................................................................................... 80.52 ............................................................................................................................... 82.86 

71.38 

83.40 

83.86 

84.86 

84.32 

l 

.............. I 

74.07 

77.41 

80.60 

82.44 
Table 2. CHARACTERIZATION TEST RESULTS SUMMARY 

Cond VIT Level 

1 I VT 4 

VT 5 

v-f 4.5 

VT 4 

VT 4 

Recharge Dissipation Operating 

(watts) ' 

16.2 

20.9 

25.9 

34.7 

19.5 

26.9 

36.1 

50.7 

20.4 

28.5 

41.4 

60 

18.4 

28.5 

42.2 

66.2 

11.8 

14 

16.9 1 16.2 

I 10.8 

1 13.3 

19.1 

, 19.4 

13.1 

Capacity (Ah) 

82.56 

82.84 

83.1 5 

82.6 

84.97 

85.55 

85.9 

85.19 

82.88 

84.08 

85.12 

84.98 

73.91 

77.44 

80.52 

82.86 

81.53 

82.29 

82.72 

99.52 

81.35 

84.14 

85.67 

96.89 

79.68 
s % B Table 2 (Cont). CHARACTERIZATION TEST RESULTS SUMMARY 

% 

d Points from LMMS heat dissi ation test 

F 

8 9 x40ba x40 0"

VT3.5 50°FF 32.71V40 11 0 29.1 4 .f x46 50°F 32.71V 11 2 34.2 

. . . - 

.. (*) Note: Operating capacity for condtions 20 and 24 

measured at 18A to 1.2V and 9A to 1 .I V. 

x46 should be vt3.5, 41°F (40A to 1.1V for all other conditions) 

70 

h 

60 - 

V) 

CI 

2 50 

3 

Y 

c 

40 - 

0 

I- 2 30 

P .- 

U) .!? 20 

n 

lo 

0 

,,4 

.4-- 

-4- . 

4*' *. 2' 1, 

4 #' ** 

Correlation =0.9995 

Fl I I I I I 

100 105 110 115 120 125 130 

Recharge Ratio (%) 
Table 3. LMMS HEAT DISSIPATION TEST CONDITION x46 

I 

1 

Date 

29-Mar-97 

30-Mar-97 

31 -Mar-97 

1 -Apt--97 

2-Apr-97 

3-Apr-97 

4-Apr-97 

5-Apr-97 

6-Apt-97 

7-Apr-97 

8-Apr-97 

9-Apr-97 

10-Apr-97 

1 1 -Apr-97 

12-Apr-97 

13-Apr-97 

14-Apr-97 

15-Apr-97 

16-Apr-97 

17-Apr-97 

18-Apr-97 

19-Apr-97 

20-Apr-97 

21 -Apr-97 

RNH-90-3 

Rechg Ratio(%) 

104.5 

105.1 

106.4 

107.2 

108.2 

109.0 

109.2 

109.9 

110.2 

110.4 

110.4 

110.6 

110.7 

111.0 

111.2 

11 1.3 

11 1.5 

111.5 

11 1.3 

111.5 

11 1.6 

111.5 

111.8 

111.7 

RNH-90-3 

Dissipation (W) 

18.8 

19.9 

22.5 

24.2 

26.2 

27.8 

28.3 

29.7 

30.4 

30.8 

30.9 

31.4 

31.7 

32.3 

32.6 

32.9 

33.4 

33.4 

33.2 

33.7 

33.9 

33.6 

34.2 

34.2 

RNH-90-9 

Rechg Ratio(%) 

104.3 

104.9 

106.3 

106.7 

107.6 

108.3 

108.4 

108.8 

109.3 

109.5 

109.5 

109.9 

109.9 

110.0 

110.1 

1J0.2 

110.3 

110.3 

110.5 

110.3 

110.4 

11 0.6 

110.5 

1 10.6 

RNH-90-9 

Dissipation (W) 

18.3 

19.5 

22.3 

23.2 

25.1 

26.4 

26.7 

27.7 

28.6 

29.1 

29.2 

30.1 

30.1 

30.3 

30.6 

30.9 

31.1 

31.1 

31.7 

31.3 

31.5 

31.8 

31.8 

32.0 
Table 3 (Cont). LMMS HEAT DISSIPATION TEST CONDITION x46 

+ Reclig Ratio 90-9 

- Dissipation 90-3 
Figure 4. HST Characterization Test Slurry Recharge Ratios 
Figure. 5. HST Characterization Test Dissipations-Slurry Cells 
Figure 6. Characterization Test Dissipations for Dry Sinter Cells 
Figure 7. Operating Capacity to 1.2V for Slurry Cells 
Figure 8. Operating Capacity to 1 .I V for Slurry Cells 
4 

R 

(D Figure 9. Operating Capacity to 1.2V for Dry Sinter Cells 

F 

3 
Figure 1 0. Operating Capacity to 1.1 V for Dry Sinter Cells 
Figure 11. HST Cell Impedance and Capacity Performance 

95 

, ......... ..... 90 

8 5 

80 

.... . . 7 5 

70 

6 5 

........ 6 0 

5 5 

Transition from 12A to 24A 

5 0 

I 

I 45 

4 2 3 3 2 4 1 5 0 ' 59 

Temperature (OF) 

t- Slurry Imp VT4 --&- Dry Sinter Imp VT4 -6- Slurry Imp VT3 -8- Dry Sinter Imp VT3 

-4- Slurry Cap VT4 -.U- Dry Sinter Cap VT4 --Q- Slurry Cap VT3 43- Dry Sinter Cap VT3 - 
'I Figure 12. Slurry Cell Voltage vs. Discharge Rate 

I 

I 

, I 

t I 

, 

, . .. -1------- -2- - - - - - '. 6- '. I- - - 

40A discharge after Cond #19 

. . , . - - - - - - - , . - . - . - . - 8 .. ... 

18A discharge after Cond #20 

,----- --0 ---- 12At24A discharges (Cond #20) - " :- ' . I. .- . 

I- 8 - -I -. 8 . ., . . 

-. < .. --- .... . !. . - .-- - --? .- - -- -- ---- I -_ -- ( - . - . . - - - , . . . . . . -- 4.- ' / a I- 

,- ---- . , 1--- .. -_ . -. , . -- , - . , --- $ 

---___ - -- I 

' ---. . .. . .,- . , - - - - - . - . , - . . . . -. .. 7 -- 

. .'a- . ,------- "I. - ' - - .,. --.--- -- 

I I I I I I I I I 
I 

Figure 13. Dry Sinter Cell Voltage vs. Discharge Rate 

VIT 4 23°F 

I 1 

I 

I I 

'8 '....' --,--- r I' 

12Al24A discharges (Cond #20) : 

1 '-3.. 40A discharge after Cond #19 , I .. 

, 18A discharge after Cond #20 -. --I- - - -. - .I- ,- -.. 

-- ,z-- J,: - - - 3. .,' - ' - ., . . 8 -. -'. - . 8- . , 3 ..-. - - I 

-....... J--- 

------,-- . . . . ' a. : . - - - - - . . ,- - ---.;_. 

. , 

... . '-, - --- ( . . - - - - - . , - - - - - - . -7--.---- -r--. -a_------- 

. .... '! ' .- ,------- , . .. ..,. ----- --, -1------ -- 

, 'i 

I I I I I I I I 

I 

i 

40 5 0 6 0 

Ampere-Hours 
Figure 14. Slurry Cell Voltage vs. Discharge Rate 

- - 30.8 

0 10 2 0 3 0 40 . 50 6 0 7 0 80 90 

Ampere-Hours 
Figure 15. Dry Sinter Cell Voltage vs. Discharge Rate 

I 

, \ 

1.10 I I I I I I I I I 

0 10 20 3 0 40 5 0 

24.2 

6 0 70 8 0 90 

i Ampere-Hours 
Figure 16. Slurry Cell Drains 

+After Cond#20 

+After Cond#23 

+After Cond #24 

6 

Time (Hours) 
'I 

Figure 17. Dry Sinter Cell Drains 
Figure 18. RNH-90-3 Condition X46 

Filename PDAT5CCAa00350 Plot Start Time 4/21 197 05:36:16 

1.60 I I I I I I 1000 

I I I I 

............... ....... I 1 55 .............. -. -. - -- I I -. ..... 

I I I 950 

I I I Cell Pressure 1 

I ..... ... 

I 

1.50 I.. 

I I 900 

I n - .... U) 

I 

850 p 

I Y 

0 I . . ............ ........ ........ > 1.40 I . 800 ! 

I I 3 

I I I U) 

- - -.- -- .I- ............ - I-.. ...... U) 01.35 -- 750 

I n 

I .... ....-.-. ....... ....... I .. .. .............. 

.. 

1-30 700 

I I 

......... 1.. ......... 

I I 650 

I I I I I I 

I 

I I 

1.20 I I I I I 6 0-0 

0 0.5 1 1.5 2 2.5 i 3 3.5 

Cycle Time (Hours) 
4 

a 

n Figure 19. RNH-90-3 Condition X46 

ba 

f? 

d Filename PDAT5CCAa00350 Plot Start Time 4/21 I97 05:36:16 

10 25.0 

Cycle Time (Hours) 

I CellTop1 ---- Cell Bottom 1 - -- --- Cell Stack 1 - - - - - Alt Stack 1 I 

CellTop2 ---- Cell Bottom 2 - - -- -- Cell Stack 2 Current I 
Figure 20. RNH-90-9 Condition X46 

Filename PDAT6CCAa00351 Plot Start Time 4/21/97 05:29:20 

1.60 I I I I I I I 35.2 

I I I I 

I 

I 

I 

I 

I 

I 

1.55 . - . .. I I I . - .- -- -- .- -- --- - -. _-._. -_ ._ __ _ . _ ._ .. __ .. __ ... _. . I ...-. . ... 

I I I I I I I 34.1 

al 

0 

1-50 -- - - -- I . - m 

33.0 

> 

31.9 p 

al 

.ICI 

30.8 

CI 

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1.25 . 27.5 

I I I I 

I I I 

I 

I 

I I 

1.20 I I I 26.4 

0 0.5 1 1.5 2 2.5 3 3.5 4 

Cycle Time (Hours) 
Figure 21. RNH-90-9 Condition X46 

Filename PDAT6CCAa00351 Plot Start Time 4/21/97 05:29:20 

10 I I I I I I I 25.0 

9 

I ........ I ..... I .. I -.I. I. . .- ... __I _-. .. I I' 

I 

I-. - 20.0 

8 

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, - 

h 1 

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I 

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Y I . - I .... 10.0 

I I I I ...... -. -- )_. - -- --- .- ...I - -.. -- ... I .. .......... -- I._ .. -- . - -- -. r- -. -_ . . I 

I .- 5.0 ; 

I I a, 

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3 

..... ...... ... .. I -- -- .- .- - ... -5.0 0 

I I .. .. ..... .. ... f , 

- -- -. - - - - - -- - -10.0 ........ 

2 I I I .-..... ,-- .. .- -- -I- -- -- - - .-I.. -. ... .- .- I -- -- ...... ......... - . -. -15.0 

1 -. I I I ...... ..... .... - _ I , -20.0 

I 

0 

I . I I , -25.0 

0 0.5 1 1.5 2 2.5 3 3.5 ' 4 

Cycle Time (Hours) 

Cell Top 1 - - - - Cell Bottom 1 - ----- Cell Stack 1 CellTop2 ---- Cell Bottom 2 

------ Cell Stack 2 - - - - - Air Top 2 ----- Air Bottom 2 Current 
CONCLUSIONS 

CELLS WITH SLURRY ELECTRODES HAVE SLIGHTLY HIGHER THERMAL 

DISSIPATION THAN DRY SINTER CELLS FOR CONDITIONS TESTED. 

SLURRY ELECTRODE CELLS HAVE LOWER RELATIVE OPERATING CAPACITY 

TO 1.2 V AND 1 .I V AT LOWER TEMPERATURES FOR V/T LEVELS 3 AND 4. 

* THERE WAS A DRAMATIC EFFECT OF CHARGE CUTOFF VOLTAGE ON HEAT 

DISSIPATION AT HIGHER OPERATING TEMPERATURES. 

HST ONBOARD BATTERY MANAGEMENT SINCE NOVEMBER 1995 SHOWS 

IMPROVEMENT IN CAPACITY REQUIRING FURTHER TREND ANALYSIS. 

- LONG TERM CAPACITY LOSSES ATTRIBUTED To PREMATURE CHARGE CUTOFF AT 

LOWER VT LEVEL (KIL4) TO LIMIT BATTERY HEAT DISSIPATION. 

- OPERATION AT LOWER TEMPERATURES AFTER SWITCHING FROM PRIMARY TO 

SECONDARY HEATERS HAVE INDICATED IMPROVEMENTS IN BATTERY CHARGE 

EFFICIENCIES AND PERFORMANCE. 
Advanced Ni-H, / Nickel Electrode Session 

1997 NASA Aerospace Battery Workshop 
Page intentionally left blank 
LANDER AND ORBITER BATTERY 

PRESENTED BY 

WILLIAM COOK 

EAGLE PICHER INC. 

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BATTERY PRESENTATION 'I 

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PRESENTATION OUTLINE 

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*PROGRAM OVERVIEW 

.TWO CELL CPV DESIGN SUMMARY 

*CPV PERFORMANCE SUMMARY 

@BATTERY DESIGN SUMMARY "LANDER" 

*BATTERY DESIGN SUMMARY "ORBITER" 

*BATTERY PERFORMANCE SUMMARY 

@QUESTIONS AND ANSWERS I, 

SARI 008511 0083 

BATTERY PRESENTATION I 

I' 

NASA BATTERY WORKSHOP 1997 
MSP PROGRAM OVERVIEW 

EAGLE~PICHER 

.EXTREMELY TIGHT COST CONSTRAINT 

*USE PROVEN DESIGN WHEREVER POSSIBLE 

*TWO SATELLITE DEVELOPMENT 

.LANDER (1 1) CPV/ (1) IPV DESIGN 

*ORBITER (11) CPV DESIGN 

SARI 008511 0083 

BATTERY PRESENTATION 

ref. AVIATION WEEK & SPACE TECHNOLOGY 12/9/96 
Common Pressure Vessel Flight History 

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.Salt ( lnterspace Navy) 

.Apex (Orbital Sciences) 

.Seastar (Orbital Sciences) 

.Orbcomm (Orbital Sciences) 

0MSTI-2 (Spectrum Astro) 

OMSTIS (Spectrum Astro) 

.Tubsat-B (Technical Univ. of Berlin) ! 

*Tubsat-C (Technical Univ. of Berlin) , 

BATTERY PRESENTATION 

SARI 008511 0083 

1 

NASA BATTERY WORKSHOP 1997 
RNHC-I 6-1 

L, 

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' NASA BATTERY WORKSHOP 1997 

SARI 008511 0083 

BATTERY PRESENTATION 
RNHC 16-1 SPECIFICATION 

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*Terminal Type( 10132 thread) 

*Nominal Capacity 10°C 

*Pressure Vessel Length 

*Pressure Vessel Diameter 

*Pressure Vessel Material 

*Separator Material 

*Cell Weight (Nominal) 

*Electrolyte Concentration 

*MEOP 

*Specific Energy Whllb 

*Safety Factor 

SARI 008511 0083 

BATTERY PRESENTATION 'I 

Rabbit Ear 

17.5 AH 

9.00in (22.86) 

2.53in (6.43) 

Inconel) 

Zircar 

950 g 

31% 

800 PSlG , 

2 8. 

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' NASA BATTERY WORKSHOP 1997 
RNH 16-3 CELL DESIGN 

LC. 

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NASA BATTERY WORKSHOP 1997 

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BATTERY PRESENTATION 'I 

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*Terminal Type( 10132 thread) 

*Nominal Capacity 1 O°C 

*Pressure Vessel Length 

*Pressure Vessel Diameter 

*Pressure Vessel Material 

*Separator Material 

.Cell Weight (Nominal) 

*Electrolyte Concentration 

*MEOP 

*Specific Energy Whllb 

@Safety, Factor 

SARI 008511 0083 

BATTERY PRESENTATION 

b I 

9 II 

Rabbit Ear 

17.5 AH 

9.00 in (22.86) 

2.53 in (6.43) 

4 

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lnconel 

Zircar 

606 g 

31% 

I, 400 PSlG , 

16.38 (36.09) 

3.0:l 

Cell Performance Specification RNH-16-3 

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X 

NASA BATTERY WORKSHOP 1997 
RNHC 16-1 TEST RESULTS . 

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AVERAGE 

SARI 008511 0083 

BATTERY PRESENTATION 'I 

II -CHARGE RET 82% 86% 85% 

SINE VIBRATION, RANDOM VIBRATION, PYROSHOCK (all passed) 

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NASA BATTERY WORKSHOP 1997 

IMPEDANCE mOHMS 7 8 7 
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SAR 10085 Lander Battew 

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NASA BA"B"TERY WORKSHOP 1997 
LANDER UNIQUE FEATURES 

EAGLE @PICHER 

*UTILIZES IPV AND CPV COMBINATION 

*SIMPLE CONSTRUCTION 

*FOUR TRANSDUCER TEMPERATURE CONTROL 

*0.080 THICK 7075-T6 ALUM BASE PLATE 

*COFFEE CAN SLEEVE DESIGN 

*VOLTAGE REGULATOR FOR ELECTRONIC 

*STRAIN GAGE AMPLIFICATION I -5VDC I 

I 

SARI 008511 0083 

BATTERY PRESENTATION 'I 

I' IJ 

NASA BATTERY WORKSHOP 1997 

LANDER I SLEEVE ASSEMBLY 
SARl0085/10083 

BATTERYPRESENTAT 'I 

SARI0083 ORBITER BATTERY 

IASA BATTERY WORKSHOP 4997 
1997 NASA Aerospace Battery Workshop -679- Advanced Ni-H, / Nickel Electrode Session 
ORBITER UNIQUE FEATURES 

4 'n 

'0 

5 '.I EAGLE @PICHER 

2 

.HONEYCOMB BASEPLATE DESIGN 

.SIMPLE CONSTRUCTION 

.FOUR TRANSDUCER TEMPERATURE CONTROL 

*HORIZONTAL MOUNT SLEEVE DESIGN 

.VOLTAGE REGULATOR FOR ELECTRONIC 

.STRAIN GAGE AMPLIFICATION IdVDC 

SARI 008511 0083 

BATTERY PRESENTATION 

2 

LY 

2. 0 

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' NASA BATTERY WORKSHOP 1997 
ORBITER CELLISLEEVE ASSEMBLY 

EAGLE @PICHER SARI 008511 0083 

BATTERY PRESENTATION 
SARI 008511 0083 

BATTERYPRESENTAT 

SARi0083 SLEEVE ASSEMBLY 
SARI 0085/10083 

BATTERYPRESENTAT 

ELECTRONICS ASSEMBLY 

"ERV WORKSHOP 1997 
BATTERY PERFORMANCE 

SUMMARY 

EAGLE@PICHER 

LANDER 

BATTERY 

SARI 008511 0083 

BATTERY PRESENTATION 

IASA BATTERY WORKSHOP 1997 
SARI0085 CHARGE VOLTAGE VS TEMP 

EAGLE~PICHER 

TIME MIN 

SARI 008511 0083 

BATTERY PRESENTATION 

I1 

NASA BATTERY WORKSHOP 1997 

1 
SARI0085 DISCHARGE VS TEMP 

C12 DISCHARGE TO 23 VOLTS 

36 00 

' 1 1 

'r 

'0 

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TIME MIN 

SARI 008511 0083 

BATTERY PRESENTATION 

El -M- ZOC 

I1 

NASA BATTERY WORKSHOP 1997 
SARI 008511 0083 

BATTERY PRESENTATION 

I 

SARI0085 BAlTERY HIGH RATE TEST 

-- 5.0000 

E 

s ,s,---------------------------------- 

TIME IN MIN 

NASA BATTERY WORKSHOP 1997 

'1 

IRIDIUMB Satellite Batteries: Flight 

Observations "By the Dozen" 

PRESENTED TO: 

1997 NASA BATTERY WORKSHOP 

NOVEMBER 18 - 20,1997 

BY: 

MARK R. TOFT 

MOTOROLA SPACE SYSTEMS(TECHNOL0GY ; 

GROUP 

IRIDIUM is a registered trademark and service mark of IRIDIUM LLC O 1997 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
AS OF NOVEMBER 15,1997 - 39 IRIDIUM SPACE VEHICLES 

HAVE BEEN LAUNCHED: 

- LOCKHEED BUS DESIGN 

- EACH SPACE VEHICLE CONTAINS A SINGLE PRESSURE 

VESSEL NICKEL-HYDROGEN BATTERY 

BATTERIES MANUFACTURED BY EAGLE-PICHER, IND. 

JOPLIN, MO 

- SOLAR ARRAY PROVIDES ENERGY BY SWITCHING 

INDIVIDUAL STRINGS ON AND OFF 

SPACE VEHICLES IN 86.4 DEGREE ORBIT INCLINATION WITH 

THE BATTERIES MAINTAINED AT 0 "C WITH AN AVERAGE 

DEPTH-OF-DISCHARGE OF -10% (WITHOUT TELEPHONY) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
OVERVIEW 

I 

ON-ORBIT BATTERIES MAINTAINED AT PRESSURE LEVEL 

CORRESPONDING TO 85% OF FULL-CHARGE PRESSURE FROM 

VENDOR'S 10 DEG. C CAPACITY TEST 

BATTERY CHARGE MODE (FULL & TRICKLE) CAN BE 

CONTROLLED BY COMBINATIONS OF PRESSURE, PRESSURE 

SLOPE, VOLTAGE, CID RATIO AND TEMPERATURE LIMITS 

- CHARGE BACK TO BATTERY-SPECIFIC PRESSURE LIMIT, 

AND RECHARGE RATIO (1.02) OR VOLTAGE LEVEL (-33.8V) 

- BATTERY REMAINS IN TRICKLE UNTIL DISCHARGE 

REACHES A CERTAIN THRESHOLD I 

I 

PARTIALLY REDUCE STRESS BY PREVENTING 

I FREQUENT RECHARGES AT SHALLOW DOD'S 

(ENTERING AND EXITING FULL-SUN) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 

BATTERY CHARGE CONTROL I 
FIRST FIVE SV'S LAUNCHED INTO FULL-SUN 

INITIAL TRICKLE RATE (C1100) CAUSED PRESSURE AND STATE 

OF CHARGE TO SLOWLY INCREASE (ABOVE DESIRED LEVELS) 

- TEMPERATURE PERFORMANCE STABLE UNDER THlS 

INITIAL TRICKLE RATE 

WHEN ECLIPSING BEGAN AFTER -4 WEEKS, THlS "EFFICIENT99 

TRICKLE CHARGING CORRUPTED CID RATIO CALCULATIONS 

AS WELL AS THE LOGIC THRESHOLDS FOR RETURNING TO 

FULL CHARGE 

TRICKLE CHARGE RATE WAS LOWERED I 

- PRESSURE DECAYS UNDER TRICKLE AT 100% STATE-OFCHARGE 

- PRESSURE MAINTAINED AT LOWER STATES-OF-CHARGE 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
'1 

IRIDIUM SV005 BATTERY PRESSURE PERFORMANCE WITH Cll 00 

TRICKLE CHARGE 

TIME (MM:DD:YY HH:MM) - . .. 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
TIME (MM:DD:YY HH:MM) - 

- - - .. 

IRIDIUM SV005 BATTERY PRESSURE PERFORMANCE WITH Cll 00 

TRICKLE CHARGE 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 

5.0 

0.0 

-5.0 

-1 0.0 

-1 5.0 

-20.0 

-25.0 

-30.0 

-35.0 

40.0 

45.0 

-50.0 

-55.0 

8/31/97 9:36 8/31/97 11:16 8/31/97 1257 8/31/97 1438 8/31/97 16:19 8/31/97 

.. - . -. 

I 

I 

- " 

AMP-HoL 6 

INTEGRATOR CANNOT l+ -+L7 - 4 

v 

-- 

SENSE RCHARGES THESE "MINI" 

AMP-HOUR 

INTEGRATOR COUNTS 

THESE DISCHARGES 

-. -. 

-- / 

/ I 

-- - 

._. __ \ 

\, 

1 

\ 

\ 

-- _- ____, 

1 

b 

\ t \ I 
MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 

IRIDIUM SV005 BATTERY PRESSURE PERFORMANCE WITH C1100 

TRICKLE CHARGE 

5.0 

0.0 

-5.0 

-1 0.0 

-1 5.0 

-20.0 

-25.0 

-30.0 

-35.0 

-40.0 

-45.0 

-50.0 

-55.0 

. .- 

RECHARGE BECOMES 

- MORE EFFICIENT AT 

EACH LOWER STATE 

OF CHARGE --. - -. . -. . 

-- 

I I u t I I 

8/31 197 9:36 8/31/97 11:16 8/31/97 12:57 8/31/97 14:38 8/31/97 16:19 8/31/97 

TIME (MM:DD:YY HH:MM) , .- - 
IRIDIUM SV005 BATTERY PRESSURE PERFORMANCE WITH Cl250 

TRICKLE CHARGE 

5.0 

0.0 

-5.0 

DECREASES SLIGHTLY AT -1 0.0 

100% STATE OF CHARGE, 

BUT IS STABLE AT LOWER I 

-1 5.0 

STATE OF CHARGE 

-20.0 - . - - - .- . . 

-25.0 

-30.0 

I 

1, 

-35.0 1 

10115197 22:33 10115197 23:02 I0145197 23:31 10116197 0:OO 10116197 0:28 10/16197 0:57 10116197 I :26 

TIME (MM:DD:YY HH:MM) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
LOWER TRICKLE RATE DOES NOT ALWAYS a 

OVERCOME BATTERY SELF-DISCHARGE ,I 

SUBSEQUENT SV's LAUNCHED INTO, OR THAT HAVE 

SINCE TRANSITIONED TO FULL-SUN OPERATIONS, HAVE 

BEEN AFFECTED IN DIFFERENT WAYS BY THE LOWER 

TRICKLE RATE 

VOLTAGE AND PRESSURE PERFORMANCE AFFECTED 

WHETHER IN FULL-SUN OR IN ECLIPSING 

- BATTERIES GAIN OR LOSE CHARGE AT VARIOUS 

RATES WHILE TRICKLING IN FULL-SUN 

- WILDLY VARYING CID RATIOS FOR THOSE BATTERIES 

IN ECLIPSING OPERATIONS (TRICKLING BETWEEN 

THE END OF FULL-CHARGE AND THE BEGINNING OF 

ECLIPSE) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
Comparison of Battery Recharge Ratios for First Launch SVs 

-. ............... -- ... 

CID RATIOS < 1 .OO WHEN 

CM00 TRICKLE RATE USED. 

NOTE SLIGHT CHANGE IN 

SV008 PERFORMANCE 

AFTER FULLSUN PERIOD. ------- --. . --- A-- ----. 4 

Day of Year 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
Comparison of Battery Recharge Ratios for Second Launch SVs 

-0- SVOIO 

+ SVOI 1 

- SVOI 2 

j+ SVOI 3 

+ SV014 

+ SV016 

Day of Year 

. . 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
Comparison of Battery Recharge Ratios for Third Launch SVs 

....- ...-...... I -- --- --------.--. - 

NOTE DIVERGENCE +SVOIS 

+ i SV017 I 

BETWEEN SV017 AND - SV018 ' 

OTHER VEHICLES, AND -svozo I 

CHANGE IN - 

PERFORMANCE OF 

SV020 OVER THE .- - . .- - - PE-RIOD;---- - -. 

1 ~~'8t~4b~ f~~lll'L9 3 *~c~sa!t tt8483#84, *4',!3#44 tP,,,~,,,,,4,,,,,,,,,,,,,,,,,,,,,,,,,,,,,,,,~~,,,,,,, ,,,, ,,,,, , 

""I" ' ' OWWN~~f~NWOW~NWOWWNWOWCONWOWWN3~WW 

Q)Q)Q)O NNNNNNNNNN~NNNN~NNNNNNN(YN~~~C~CNNC3C) WbW~~~WWwhhhWWmmmo 7- 7- 

EEEhh~~hhhhhhhh~hhhhhhhhhhhhbbhhb mmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmmm 

Day of Year 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
I I Comparison of Battery Recharge Ratios for Fourth Launch SV's 

--- 

NOTECHANGE IN - ' 

PERFORMANCE FOR SV025 

OVER THE PERIOD, AND I -sv025 

CHANGE IN PERFORMANCE ' \L SV026 

OF SV's 026 AND 024 

RELATIVE TO EACH OTHER. ........................ 

Day of Year . . ........ . . ... I!. ......... . . ....-.... - . ... - -. ..... 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
Comparison of Battery Recharge Ratios for Fifth Launch SYs 

. -- - .- 

-- --- -- - + SV027 

+ SV028 

- -- ---- . - -- 

- 

+ SV029 

SV030 

- - - - -- - - - .- - - 

+ SV031 -- - - -- - - -. . --- .- + SV032 

+ SV033 

- - - - - .- - - - - -. ---- 

Day of Year 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
Comparison of Battery Recharge Ratios for Sixth Launch SVs 

TRICKLE RATE i 

INCREASED / 

FROM Cl250 TO 

Cll60 ON SV035 i 

Day of Year (YYDDD) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
I .. .- '1 

IRIDIUM 2ND LAUNCH VEHICLES: VOLTAGE PERFORMANCE IN FIRST FULLSUN PERIOD UNDER C1250 TRICKLE-CHARGE RATE 

9/27/97 0:OO 9127197 12:OO 9/28/97 0:OO 9/28/97 12:OO 9/29/97 0:OO 9129197 12:OO 9130197 0:OO 9130197 12:OO 

DATE -TIME (MMIDDNY HH:MM) .- 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
IRIDIUM 2ND LAUNCH VEHICLES: PRESSURE PERFORMANCE IN FIRST FULLSUN PERIOD UNDER Cl250 TRICKLE-CHARGE RATE 

I DATE -TIME (MMIDDNY HH:MM) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
I i 

I I 

'I 

IRIDIUM 6TH LAUNCH VEHICLES: VOLTAGE PERFORMANCE IN FIRST FULL- 1: 

I i SUN PERIOD UNDER Cl250 TRICKLE-CHARGE RATE 

- SV034 

SV035 

. . .. ... SV036 - - 

'--*--- ...-.. *-- 

- 'C -.A. .,. *, 

9130197 0:oo 1 OM 197 0:00 1012l97 0: 00 10/3/97 0: 00 10/4/97 0: 00 ' 10/5/97 

DATE -TIME (MMIDDIYY HH:MM) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
. -- - - '1 

IRIDIUM 6TH LAUNCH VEHICLES: PRESSURE PERFORMANCE IN FIRST FULLSUN PERIOD UNDER C1250 TRICKLE-CHARGE RATE 

10.0 

, I 1 

DATE -TIME (MMIDDNY HH:MM) 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
LOWER TRICKLE RATE CAN RESULT IN SIGNIFICANT SELF ( DISCHARGE WlTH ASSOCIATED HIGHER CID RATIOS 

IRIDIUM SV026 BATTERY PRESSURE PERFORMANCE WlTH Cl250 

TRICKLE CHARGE 

INTEGRATOR CANNOT 

SENSE THIS "MINI" 

DISCHARGE --I-- =-=--.=-= - ---;_ -.=-. _ 

INTEGRATOR COUNTS 

- - 

, 

I 

9/2/97 21:07 912197 2136 9/2/97 22:04 9/2/97 22:33 

3 /I TIME (MM:DD:YY HH:MM) 3 

to 

B 

6' 

3 MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
ANALYSIS 'I 

PROBABLE CAUSE(S): 

- SELF-DISCHARGE RATE VARIES FROM ONE BATTERY TO ANOTHER 

- CURRENT SENSOR ERROR (ACCURACY f 0.1 AMPS) 

SENSOR READS PROPER CURRENT BUT IS IN FACT TOO LOW 

INVESTIGATION: 

- CURRENT SENSOR ERROR SHOULD PROVIDE STATIC 

DIFFERENCES IN PERFORMANCE (e.g. CID RATIO ALWAYS 5% 

HIGHER) 

- THE FACT THAT PERFORMANCE CHANGES WITH TlME ANDIOR 

RELATIVE TO ANOTHER BATTERY UNDER SIMILAR CONDITIONS 

SUGGESTS THAT BATTERY IS SOURCE OF PHENOMENON 

* SOME BATTERIES WILL NEED TO HAVE TRICKLE RATE 

INCREASED REGARDLESS OF THE SOURCE OF THE 

PHENOMENON 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
ADDITIONAL INVESTIGATION 1 

AS A TRIAL, THE TRICKLE RATE WAS INCREASED FROM C1250 

TO C1160 ON SVO18 (FULL-SUN OPERATIONS) AND SV035 

(ECLIPSING) 

- SELF-DISCHARGE WAS REVERSED ON SV018; CID RATIO WAS 

REDUCED BY 4.5% ON SV035 

ANY INCREASES IN TRICKLE CHARGE MUST BE CAREFULLY 

CONSIDERED AND CLOSELY MONITORED 

- SINCE BATTERY PERFORMANCE SEEMS TO BE CHANGING, 

TRICKLE RATES MAY HAVE TO KEEP CHANGING TO KEEP UP 

- E.g. FIRST-LAUNCH SV TOLERANCE OF HIGHER TRICKLE RATE HAS 

CHANGED 1, I 

RECENT CHANGE BACK TO HIGHER TRICKLE RATE INCREASED 

I BATTERY TEMPERATURE -2 DEG. C 

REMOVAL OF HIGHER RATE REDUCED BATTERY 

TEMPERATURE 

MOTOR0L.A SPACE SYSTEMS TECHNOLOGY GROUP 
FIRST-LAUNCH SV TEMPERATURE RESPONSE TO 

REDUCTION IN TRICKLE CHARGE #I 

SV006: BATTERY TEMPERATURE RESPONSE TO DECREASE IN TRICKLE 

CHARGE RATE FROM CIIOO TO Cl250 

TIME (MMIDDMY HH:MM) - - - - - . . 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
CONCLUSIONS AND FUTURE STRATEGY. 'I 

BATTERIES NEED DIFFERENT TRICKLE CHARGE RATES 

EITHER BECAUSE OF CURRENT SENSOR ERROR OR 

VARIABLE SELF-DISCHARGE RATES 

HIGHER TRICKLE RATES WlLL OFFSET THIS EFFECT BUT 

MAY CREATE UNDESIRABLE WASTE HEAT 

AVOIDING EXCESSIVE WASTE HEAT SHOULD 

CONTRIBUTE TO BATTERY LONGEVITY, WHICH IN TURN 

CONTRIBUTES TO IRIDIUM SUCCESS 

BATTERY OVERCHARGE AND THERMAL STRESS WlLL 

CONTINUE TO BE MINIMIZED 

MOTOROLA SPACE SYSTEMS TECHNOLOGY GROUP 
BEHAVIOR OF 2-CELL CPV NI-H2 BATTERY DURING PULSE DISCHARGE 

- 

Hari Vaidyanathan 

COMSAT Laboratories 

Clarksburg, MD 

AND 

Gopalakrishna Rao 

NASA-GODDARD SPACE FLIGHT CENTER 

Greenbelt, MD 

ABSTRACT 

A study was carried out to determine the transient voltage behavior of 

the 2-cell CPV nickel-hydrogen battery with the objective of using the results 

as a basis for mathematical modeling . The 2- cell CPV battery which is 

manufactured by Eagle Picher, Inc. for the GOES program yields 18.5 Ah at C/2 

rate of discharge at 10°C with a mid-discharge voltage of 2.514 V. The capacity 

increased with decrease of temperature and a maximum capacity of 22 Ah was 

obtained at - 5OC. The pulse tests consisted of obtaining the voltage profile in 

the first 20 milliseconds of the one minute pulse discharge at 37 A and pulse 

discharge was repeated as a function of state- of- charge. The pulse test at 10°C 

and 20°C provided voltage profiles with the expected decrease in voltage as 

the pulse was applied. The end of pulse voltage decreased with the state-ofcharge. The battery voltage was above 2V at the end of the one-minute pulse 

at 8 % state-of-charge at 10°C. The voltage profile during the 37 A pulse 

discharge consists of an initial drop in voltage which was independent of the 

state of charge. The invariability in the value for the initial drop in voltage 

with state of charge is a very important observation. The results show that 

towards the end of discharge the dominant resistance is not ohmic in nature. 

It could be hypothesized from the nature of the voltage transients that the 

dominant mechanism towards the end of discharge is proton diffusion. The 

study also shows that the dominant resistance in the voltage plateau during 

discharge is activation polarization. 

INTRODUCTION 

The 2-Cell CPV nickel-hydrogen battery manufactured by Eagle Picher 

is a variation of the aerospace individual pressure vessel cell (IPV) with the 

same elctrochemical performance features but with an'operating voltage of 

2.5 V and a specific energy of 58 Wh/Kg. The purpose of this work is to 

examine the characteristics of the 2-cell CPV battery including the behavior 

during pulse discharge. The purpose of the pulse test was to determine 

lowest state of charge at which the battery can sustain a 37 A pulse discharge 

I997 NASA Aerospace Battery Workshop -713- Advanced Ni-Hz/ Nickel Electrode Session 

I 
with minimum voltage degradation and assess the effect of high burst 

currents on the usable capacity. Another objective is to analyze the transient 

voltage response to determine the dominant mechanism for the abrupt 

decrease- in voltage at low states-of- charge. 

CAPACITY AND SELF DISCHARGE - - 

The capacity of the four 2-cell CPV batteries at 10°C was determined to 

be 18.5 Ah ( to 2V) and the mid-discharge voltage to be 2.514 V. The capacity 

remaining after 72-hours of open-circuiting at 10°C was 86.7%. Thus, the selfdischarge rate is similar to that of single cells. Figures 1 and 2 show the 

voltage profiles during charge and discharge, respectively and the features are 

such as voltage roll-over during charge and plateau region during discharge 

are identical to that of IPV cells. The mid-discharge voltage during discharge 

at C/2 after the 72-hour stand test was 2.485 as shown in Figure 3. Thus, the 

initial capacity tests indicate performance features of two properly connected 

cells in series. 

- 

RATE AND TEMPERATURE DEPENDENCY 

The variation of capacity with the rate of charge and discharge and 

temperature was determined. The capacity was the highest with a value of 

22 Ah at -5°C at a charge rate of C/10 and discharge rate of C/10. The capacity 

increased from 19.58 Ah to 20.35 Ah when the charge rate was increased from 

C/20 to C/2 at 10°C. The capacity at a charge rate of 0.8 C and discharge rate of 

0.8 C was 20.05 Ah. The capacity increased with decrease of temperature and 

at -5 "C, a charge rate of C/2 yielded a capacity of 20.97 Ah. Figure 4 shows the 

voltage profiles at different rates of discharge at 10°C for one of the CPVs. As 

expected the mid-discharge voltage decreased with increase in the rate. The 

capacity decreased when the discharge rate was increased from 0.5 C to 0.8 C. 

Figure 5 shows the voltage profiles at C/2 rate of discharge at different 

temperatures. At -5°C the discharge voltage is high in the beginning and low 

towards the end. In the plateau region the mid-discharge voltage is almost the 

same at all temperatures studied. Figure 6 shows the voltage profiles at C/10 

rate of discharge at different temperatures. The profiles at C/10 differ from 

that obtained at C/2 rate with respect to the appearance of a shift in the middischarge region and the observation of higher discharge voltages at -5°C 

towards the end of discharge. It is to be noted that the negative electrode 

polarization is higher at - 5OC and this characteristic results in a slightly lower 

discharge voltage at -5°C The reason for the higher voltage beyond the middischarge region at C/10 rate compared to C/2 is due to the overwhelming 

influence of the polarization at the nickel electrode. The variation of capacity 

with temperature when charged and discharged at 0.8 C was also determined 

as shown in Table 1. The mid-discharge voltage and capacity showed 

marginal changes with the temperature. 

I99 7 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
PULSE TEST 

-The pulse test was performed using a Hewlett-Packard 6050 A 

programmable electronic load which has a slew rate of 400,000 A/s to reach 

the pulse current of 37 A in <I50 ps. The CPV battery voltage was measured 

using a Nicolet 460 storage oscilloscope. The procedure consisted of obtaining 

the transient voltage profile in the first 20 milliseconds of a one-minute pulse 

at 37 A. The application of the 37 A pulse was done by increasing the 

discharge rate from 9.3 A to 37 A during discharge at regular intervals such as 

90% state-of -charge, 80%, 70 %, etc. The application of pulses were concluded 

when the voltage reached 2 V during the 9.3 A discharge. The pulse discharge 

was first done with the CPV battery stabilized at 10°C and then the test was 

repeated at 20°C. 

Figure 7 shows the voltage transient obtained during the first 2 

milliseconds of pulse discharge. The voltage of the battery decreases by 0.71 V 

in first few microseconds and then the voltage recovers to a value 200 mV 

less than that obtained during the 9.3 A discharge. This is an expected 

behavior and the initial drop of 710 mV translates to an ohmic resistance of 

25.6 milliohms. 

Figure 8 shows the variation of end of pulse battery voltage as a 

function of state-of -charge at 20°C. The features of the curve are similar to a 

discharge voltage profile with an initial decline, a plateau region and final 

decrease in voltage. Thus, the battery is able to sustain pulse discharges at 

states-of-charge as low as 12%. A very similar behavior was observed at 10°C 

and the lowest state-of-charge at which the battery would support a pulse 

current was 8 %. 

Figure 9 shows the variation of initial drop in voltage with the state-of 

charge. The trend in the data showed that the initial drop was - 670 mV and 

it did not change with state-of-charge. The initial drop in voltage is related to 

the ohmic resistance at the nickel electrode surface. It is known that, the 

discharged active material, Ni(OH)2 grows on the surface of the charged 

active material. The electronic conductivity is expected to decrease with the 

growth of the discharged active material on the nickel electrode with the 

consequent increase in polarization. The pulse behavior data indicates that 

the ohmic resistance calculated from the initial drop in voltage amounts to - 

25.6 mohms and it does not increase with the accumulation of discharged 

active material. 

. - 

DOMINANT RESISTANCE 

During constant current discharge the reaction rates at the nickel and 

hydrogen electrode proceed at a constant rate, the reaction at the nickel 

electrode being the limiting factor. - The hydrogen electrode functions with an 

average polarization loss of 25 mV at 15 m~/cm~ at a discharge rate of C/2. 

1997 NASA Aerospace Battety Workhop -715- Advanced Ni-H, / Nickel Electrode Session 
Thus the shape of the voltage - time curve depends primarily on the reactions 

of Ni(OH)2 which consists of processes like the diffusion of OH- ions in the 

bulk electrolyte, diffusion of protons into the bulk of the solid phase, increase 

of OH- concentration on the nickel electrode surface, and build-up of Ni(OH)2 

on the surface of NiOOH active material. The voltage of the nickel electrode 

- - and the cell is therefore governed by:l 

Concentration polarization due to a rise in the KOH concentration 

Activation polarization or charge transfer resistance of the nickel 

electrode reaction 

Diffusion polarization related to proton 

Ohmic effects: resistance polarization at the solid phase 

The dominant resistance at different segments of the discharge curve is 

not known with certainty. The dominant resistance in the plateau region of 

the discharge curve is said to be activation polarization. The dominant 

resistance towards the end of discharge (knee of the curve) during constant 

current discharges is said to be ohmic in nature. The invariability in the 

initial voltage drop obtained in the pulse test indicates that the dominant 

resistance is not ohmic. The other possibilities such as concentration 

polarization and activation polarization can be excluded based on published 

information in the literature. The other alternative is polarization loss 

attributable slow proton diffusion. A previous study by Salkind et a12 

attributed the suppressed discharge voltages at 1C rate of discharge at 25% 

state of charge by sintered impregnated electrodes to a diffusion problem since 

an ohmic effect would have magnified the effect in a fiber mat electrode. The 

results of this study also suggests that the voltage drop is attributable 

diffusion and specifically proton diffusion. Weidner et a13. pointed out that 

the ohmic effect is not significant during a major portion of discharge and 

contributes very little to the overall polarization loss, but conductivity at the 

end of discharge causes significant polarization losses. The results obtained in 

our study are contrary to this and the reason for the large polarization loss 

towards the end of discharge is not ohmic in nature. 

CONCLUSIONS 

The results of this study suggest the following conclusions: 

The CPV battery yields a capacity of 22 Ah and retained 86.7% of the capacity 

in the 3-day open-circuit stand test. 

The battery can sustain a 3C rate pulse for one minute at low states-ofcharge (8%). The initial decline in the voltage during the 3C rate pulse is 

about 670 mV and the voltage recovers in less than one millisecond. 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Analysis of the initial decline in voltage during the pulse test indicates that 

the resistance polarization is not significant and the polarization loss at very 

low states-of-charge is attributable to slow proton diffusion. 

REFERENCES 

- 

1. Z. Mao and R.E.White and John Newman, J.Electrochem. Soc., Vol 141, 

1994,p.54 

2. Salkind, A.J., Kelly, J.J., and Ockerman, J.B. Proc. Symp.on the Nickel 

Electrode, 176th Meeting of the Electrochemical Society, Hollywood, F1, 

Pennington, N.J. Vo1.90-4, pp247-260 

3.J.W.Weidner, and P. Timmerman, J. Electrochem. Soc ,Vol 141,1994,p 346 

I997 NASA Aerospace Batteiy Workshop -717- Advanced Ni-H, / Nickel Electrode Session 
FIGURE I: VOLTAGE PROFILES DURING CHARGE AT C110 AT 10°C 

TIME (Hrs.) 
FIGURE 2: VOLTAGE PROFILES DURING DISCHARGE AT C12 AT 10°C 

3.0 

2.0 

0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 I5 16 17 18 19 20 

CAPACITY (Ah) 
s '4 

5 

2 FIGURE 3: VOLTAGE PROFILES AFTER OPEN-CIRCUIT STAND AT 10°C 

P 

y 3.0 

f 

Q 

Ep 2.9 3 

P 8 2.8 !z 

4 

2.7 

9 2.6 Y 

W 

j(3 h, 8 4 2.5 

Id 

> 2.4 

b 

R 2.3 8 

ii 

3 2.2 

-& 

\ 

3 2.1 ii 

8 

2 a 2.0 

R 

o, 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 

8 

b' 

3 CAPACITY (Ah) 
FIGURE 4: VOLTAGE PROFILES AT DIFFERENT RATES OF DISCHARGE ATIO°C 

- 0.1 C Rate (1.86 A) 

--- 0.5 C Rate (9.26 A) 

... -.. 0.8 C Rate (1 4.86 A) 

h $ 2.3 

8 

Q 

3 

-& 2.2 

\ 

!T 2.1 

2 

2 a 2.0 

% 

%= 0 4 2 3 4 5 6 7 8 9 1011 12 131415 1617 18192021 22 2324 25 

z I 6' 3 CAPACITY (Ah) 
FIGURE 5: VOLTAGE PROFILES AT DIFFERENT TEMPERATURES DURING 

DISCHARGE AT CI2 I. 

-5'C 

--- 1 O'C 

...... 2Oec 
FIGURE 6: VOLTAGE PROFILES DURING C110 DISCHARGE AT DIFFERENT 

,. TEMPERATURES I. 

10 11 12 13 

CAPACIN (Ah) 
FIGURE 7:VOLTAGE TRANSIENT OBTAINED AT 21% STATE OF CHARGE 

-0.00005 0.00000 0.00005 O.0001 0 0.00015 0.00020 

TIME (S) 
FIGURE 8. VARIATION OF END-OF-PULSE VOLTAGE WITH THE STATE-OFCHARGE 

STATE-OF-CHARGE (Oh), 
FIGURE 9. VARIATION OF INITIAL DROP IN VOLTAGE WITH STATE-OF-CHARGE 

DISCHARGED CAPACITY (AH) 
TABLE 1. 

!? 

a 3 72-hoi~r open circuit stand 

R 

Current L1-007 L1-013 L1-027 L1-028 

cycle Temp Charge Discharge EOCV Mid-Dch V Capacity EOCV Mid-Dch V Capacity EOCV Mid-Dch V Capacity EOCV Mid-Dch V Capacity 

('C) (A) (A) M 0 (Ah) (v) (v) (Ah) M M (Ah) M 0/) (Ah) 

1 10 0.8 8 2.976 2.495 19.57 2.970 2.498 19.42 2.976 2.498 19.57 2.998 2.495 19.77 

2 0 1.6 8 3.082 2.491 20.27 3.076 2.498 20.17 3.080 2.498 20.23 3.093 2.499 19.45 

3 10 1.6 8 2.992 2.513 18.64 2.990 2.516 18.57 2.993 2.510 18.68 3.005 2.519 17.80 

4* I0 1.6 8 3.010 2.482 16.36 3.008 2.488 16.32 3.010 2.480 16.36 3.022 2.492 14.90 

5 20 1.6 8 

6 -5 1.86 1.86 

7 10 1.86 1.86 8; 

8 20 1.86 1.86 

9 -5 9.26 9.26 

10 I0 9.26 9.26 

11 20 9.26 9.26 

12 10 14.81 14.86 

13 15 14.81 14.86 

14 20 14.81 14.86 

2.941 2.548 16.45 

3.044 2.567 22.61 

2.992 2.576 20.13 

2.944 2.607 17.59 

3.258 2.470 21.16 

3.135 2.483 20.53 

3.019 2.489 18.96 

3.220 2.429 20.21 

3.190 2.428 19.98 

3.143 2.429 19.55 

2.941 2.552 16.34 

3.043 2.570 22.51 

2.991 2.576 20.08 

2.945 2.610 17.48 

3.252 2.478 21.10 

3.129 2.487 20.53 

3.020 2.491 19.03 

3.213 2.434 20.21 

3.184 2.433 19.89 

3.135 2.431 19.48 

2.943 2.545 16.58 

3.042 2.566 22.69 

2.992 2.576 20.27 

2.945 2.609 17.75 

3.265 2.471 21.21 

3.144 2.482 20.61 

3.032 2.487 19.17 

3.235 2.426 20.28 

3.204 2.425 20.05 

3.158 2.425 19.66 

2.944 2.558 15.67 

3.048 2.573 21.71 

2.999 2.577 19.79 

2.949 2.611 16.96 

3.233 2.482 20.43 

3.115 2.492 19.74 

3.004 2.494 18.54 

3.193 2.437 19.52 

3.160 2.436 19.20 

3.114 2.433 18.91 
Page intentionally left blank 
NICKEL-HYDROGEN (NiH,) - - 

SINGLE PRESSURE VESSEL (SPV) 

BATTERY DEVELOPMENT UPDATE 

Steve Sterz, Beth Parmley, Dwight Caldwell and John Bennett 

Technologies Division, Eagle-Picher Industries, Inc. 

Joplin, MO 64802 

(41 7) 623-8333 

ABSTRACT 

The NiH, battery system is the dominant rechargeable technologyfor communications, scientific and military satellites and is recognized for its long life, reliability and durability. Eagle-Picher 

Industries, Inc. (EPI) has developed the Single Pressure Vessel 

(SPV) NiH, battery, which houses all cells within a single vessel, 

as an evolutionary step in NiH, technology. The SPV configuration was developed to improve the performance, producibility and 

reliability of the NiK, system at the battery level and ultimately to 

reduce overall battery cost and weight. 

Recognition of the benefits and advantages of the SPV resulted 

in its choice as the battery design for a large-scale communications 

satellite program. After instituting high-volume production procedures and undergoing a significant life test effort at EPI, this program is proceeding very smoothly, on time and on cost This particular application requires a 22 cell, 60 Ah Low Earth Orbit (LEO) 

battery, however, 50 Ah, 30 Ah and 15 Ah battery configurations 

have been produced and capacities up to 120 Ah are possible with 

a short development time. An accelerated (70% depth-of-discharge 

for 4612 cycles) life test of an early prototype and an independent 

design review have both shown the design to be robust. The advanced SPV battery has the flexibility to fulfill the requirements of 

any mission needing a reliable, durable, space-qualified battery with 

a high duty cycle capability. This presentation discusses SPV design features and presents life test data Also addressed are the production features of the SPV battery which have resulted in cost and 

performance benefits for the IRIDIUM@ global communications 

system. 

1997 NASA Aerospace B 

INTRODUCTION 

The nickel-hydrogen (NiH,) battery system has been extensively 

developed over the past twenty years, primarily for space applications such as communications satellites. The battery combines a 

traditional nickelhickel-hydroxide electrode, as is used in nickelcadmium and nickel-iron batteries, with a fuel-cell type catalytic 

hydrogen electrode. Since hydrogen is a very lightweight, efficient 

energy storage material, this allows the prospect of higher energy 

density than typical chemical-type electrodes such as cadmium or 

iron. 

Several distinct NiK, cell and battery designs are currently in 

production and under development for a wide variety of applications. These include traditional aerospace applications such as earthorbital communications satellites as well as terrestrial uses such as 

telecommunications equipment, utility load leveling and remote 

location power systems. Existing Individual Pressure Vessel (IPV) 

Niq technology has been supplemented by newer Common Pressure Vessel (CPV) cell designs and other cell and battery designs 

such as the Dependent Pressure Vessel (DPV) and Low Pressure 

Vessel (LPV) battery. The SPV battery design is the next step in 

the continued development and evolution of the NiH2 battery system, and with the launch of the first five IRIDIUM@ satellites in 

May, 1997, EPI has the distinction of being the first battery manufacturer to orbit SPV NiH, battery technology used in a telecommunications application. 

NIH, BAlTERY DESIGN CONCEPTS 

Traditional 'N~H, battery design (Figure 1) consists of multiple 

IPV cells connected in series to develop the required battery voltage. The cylindrical cells are packaged together in a rectangular 

battery footprint. The cells are sleeve mounted perpendicular to the 

'attev Workshop, Huntsville, AL, Nov. 18-20, DC041097.PM6 

1997 NASA Aerospace Battely Workshop -729- Advanced Ni-Hz/ Nickel Electrode Session 
SINGLE PRESSURE VESSEL DESIGN FEATURES 

FIGURE 1 

baseplate. Volumetric packaging efficiency is low. Each individual 

cylindrical pressure vessel contains a single cell. The individual 

cells are sized to deliver the rated ampere-hour (Ah) capacity of the 

battery and delivers 1.25 VDC, which is characteristic of the nickelhydiogen electrochemical couple. In larger sizes, typically above 

60 Ah, each cell contains two discrete electrode stacks, connected 

electrically in parallel, within the pressure vessel. This type of internal dual-stack cell construction is shown in Figure 2. The next 

step for the dual-stack IPV is the common pressure vessel (CPV) 

cell design. This design incorporates two discrete electrode stacks, 

connected electrically in series, within a common pressure vessel. 

The design is mechanically identical to a dual-stack IPV cell with 

the only difference being the electrical connection of the two electrode stacks. The CPV cell has a terminal voltage of 2.5 VDC, which 

is the additive voltage of the two series-connected electrode stacks. 

The electrical capacity of the cell is the rated capacity of either 

electrode stack (the two stacks are of equal capacity). 

w 

FIGURE 2 

1997 NASA Aerospace Bt 

The next logical step after the CPV design, which contains only 

two cell stacks, is to incorporate an entire multicell battery within a 

single pressure vessel. The single pressure vessel NiH: battery design concept is a full battery of series-connected cells within a single 

hermetically sealed contaim(Figure 3). The number of cells within 

the battery is a function of the desired voltage. Each SPV cell delivers the 1.25 VDC typical of the NiH, couple. In this design, each 

cell shares a common hydrogen atmosphere within the battery pressure vessel. Therefore each cell is operating at the same hydrogen 

pressure throughout charge and discharge. This provides a pressure 

balance between cells. The individual cell doesn't contain a net 

pressure differential with the internal battery atmosphere. The advantages of this design concept are increased volumetric energy 

density and specific energy, improved performance, reduced complexity of construction and decreased overall battery cost. For a 

typical 28 VDC spacecraft battery, a single container of twentytwo series-connected cells is more efficient than twenty-two individual, smaller pressure vessels. (1) 

FIGURE 3 

SINGLE PRESSURE VESSEL CELL DESIGN 

Cell design is a critical aspect of the SPV battery system. The 

cell design must conform to a variety of mechanical, electrical and 

thermal interface requirements at the battery level. The semi-circular lightweight plastic cells use a dual-layer plastic bag and conform to a ten inch diameter battery pressure vessel. Cell geometry 

is designed to fit a cylindrical pressure vessel, which is discussed in 

the next section as part of the overall battery design. The overall 

cell design concept is to maintain electrplyte integrity and isolation 

for each cell while allowing the free exchange of common hydrogen gas throughout the battery. This is accomplished by venting 

the plastic cell case with a microporous plug which selectively allows the diffusion of hydrogen gas while preventing the passage of 

water or electrolyte. The holes in the material are small enough so 

that the surface tension of the electrolyte blocks its passage but 

large enough to allow sufficient hydrogen flow into and out of the 

zttery Workshop, Huntsville, AL, Nov. 18-20, DC041097.PMd 

1997 NASA Aerospace Battery Workhop -73 10- Advanced Ni-Hz/ Nickel Electrode Session 
cell. The hydrogen gas pressure inside the SPV cell remains equal- -. zed with the ambiegt pressure within the battery pressure vessel. 

This eliminates the requirement of the individual cell case to contain any pressure. i'he SPV cell operates in a starved electrolyte 

mode, as with other NiH, cell designs, so there is virtually no free 

electrolyte. Nearly all of the electrolyte contained by the cell is 

absorbed within the electrodes and separators. 

Cell mechanical design is oriented primarily towards electrolyte 

containment. The physical structure of the cell is simply the interleaved electrode stack containing the electrodes and separators. 

Electrical design of the cell consists of connecting electrodes of the 

same type electrically in parallel. The capacity of the cell is determined by the number of nickel electrodes connected in parallel. 

(The hydrogen electrodes are likewise connected electrically in 

parallel.) The electrical connections are made by welding electrical 

tabs to a comb-type bus bar. The hydrogen electrode contains an 

integral tab and the nickel electrode has a welded tab. Because of 

the cell geometry, the tabs in an SPV cell are much shorter than 

those in an IPV cell. This results in an overall low impedance cell 

design (typically less than half the impedance of an IPV cell). The 

bus bar design also provides some additional mechanical support 

to the electrodes and helps preclude electrical short-circuit between 

opposing electrodes. 

SINGLE PRESSURE VESSEL BATTERY DESIGN 

The SPV battery design is similar to standard 3.5 inch Nq spaceflight cell technology. The Inconel alloy pressure vessel is a cylinder with hemispherical domes on either end. This configuration 

allows the maximum strength-to-weight ratio for containing gas 

pressure with the minimum material weight. The pressure vessel 

maintains a 2.5: 1 safety margin, based on the maximum operating 

pressure of the battery, and is classified as a leak before burst container under MIL-STD- 1522. Normal battery operating pressure may 

range up to 800 psi, depending on the specific battery design and 

the free volume available to contain the hydrogen gas required for 

operation. 

The electrical feedthroughs for the battery use a mechanical compression seal. The battery electrical terminal passes through, and is 

insulated from, an Inconel terminal boss. A machined polymeric 

material is mechanically crimped into place and forms the actual 

seal around the battery electrical terminal. The Inconel pressure 

vessel is closed by laser welding the two dome ends to the cylindrical body of the battery. This provides a hermetic seal. A weld ring 

is incorporated into each of the two weld joints. 

The internal design and construction of the battery is shown in 

Figure 4. The semi-circular cells are arranged in pairs to form a 

circle and those circular pairs are arranged along, and perpendicular to, the central longitudinal axis of the battery. The number of 

cells included determines the battery output voltage. Each pair of 

cells is contained in a flat thermal cooling plate with radial fins 

which contact the battery pressure vessel wall. The fins on each 

cooling plate are sprung outward slightly in order to make positive 

FIGURE 4. SPV BAITERY. 

mechanical contact with the pressure vessel wall when inserted 

during battery assembly. This provides a maximum, direct thermally conductive path to remove heat from the cell to the pressure 

vessel where it can be dissipated into an exterior heat sink, such as 

a battery baseplate or thermal radiator in the spacecraft. The thermal plate is constructed of lightweight aluminum and has a nickel 

coating. The plate is electrically isolated from the adjacent cells by 

the plastic cell case. 

The cell electrical terminals are arranged along the central axis 

of the battery to minimize resistance within the plates. A mechanical connection is used between the terminals of the adjacent cells; 

all of the connections within the cell are welded. Large conductor 

cross-sections are used to fi.uther minimize IR voltage drop. 

TESTING 

The EPI SPV design has been optimized for electrical performance. The impedance of the 50 Ah SPV battery, SAR 10065, is 

less than half that of a representative 50 Ah IPV battery. Beyond 

the immediately apparent benefits of lower charge voltage and higher 

discharge voltage, this also increases the current carrying capability of the battery. As part of its acceptance test procedure, the SAR10065 undergoes current rates in excess of C/2 for charge and 2C 

for discharge. In addition, heat generation increases directly with 

impedance, thus the SPV allows the use of a smaller, lighter weight 

cooling system. 

In order to prove the durability of the design, an accelerated life 

test and three full length life tests were initiated. The recently completed accelerated life test, run at 10°C, subjected a 43 Ah actual 

capacity battery to successive 100 minute LEO cycles. The cycles 

consisted of a 36.7 A charge for 51 minutes followed by 36.7 A 

discharge for 49 minutes, for a depth-of-discharge (DOD) of approximately 70%. Over the course of the test, the charge rate was 

increased as necessary to keep the end of discharge voltage (EODV) 

1997 NASA Aerospace Battery Workshop, Huntsville, AL, Nov.18-20, DC041097.PMd 

1997 NASA Aerospace Battery Workshop -73 1 - Advanced Ni-H, / Nickel Electrode Session 
above 1 Vlcell. An EODV plot for the 46 12 cycle test is shown in 

-Graph 1. Wherenot~therwise noted, discontinuities represent short 

(four hours or less) periods of open circuit. 

After the high rate test, a low rate test was performed. In this 

test, also a 100 minute LEO cycle, the battery was charged at 16.1 

A for 51 minutes and discharged at 15.6 A for 49 minutes, for a 

DOD of approximately 31%. The 4812 cycle EODV plot of this 

test is shown in Graph 2. 

27 

Restart at& I month open circuit. 

Restart Mer I day open circuit. 

B 

37.8 A 

1 21 

0 1000 2000 30M) 4WO 5000 

(Zbn Plrrnkrr 

GRAPH 1. HIGH RATE EODV VS ORBIT NUMBER 

GRAPH 2. LOW RATE EODV VS ORBIT NUMBER 

The ongoing 111 length life tests are 100 minute, 31% DOD 

LEO cycles. Both the charge and discharge waveforms are pulsed 

(the details of the profile are proprietary). After more than 15,900 

cycles, a 50 Ah nameplate test battery, operating at 10°C, has exhibited excellent electrical characteristics and is expected to exceed 

the 25,000 cycle specification requirement. Two other full-length 

life test batteries, operating at -5OC and S°C, have completed over 

7500 and 4100 cycles respectively, and have shown the same positive results. 

MANUFACTURING CAPABILITY 

By virtue of its design, the SPV battery is better suited to high 

volume production than the IPV battery. The use of one pressure 

vessel requiring only 2 girth welds is much easier produced than a 

22 cell IPV battery requiring 22 welds. Proof pressure testing of a 

single pressure vessel is also much faster and easier than repeating 

the same process on 22 IPV cells. In addition, gas management is 

simplified by reducing the number of hydrogen seals from 66 to 4 

I997 NASA Aerospace B 

I997 NASA Aerospace Battery Worhhop 

and all cells are at the same pressure, leaving only one pressure to 

be monitored and set. Efficiency is also gained by activating and 

testing the cells in the final battery configuration. With an IPV 

battery, there is a cell acceptance test procedure (ATP) and a battery ATP, so the cells are subjected to essentially the same test procedure twice. - - 

The internal design of the SPV is also well suited to production 

speed and efficiency. The plastic cell cases require less fabrication 

time and fewer high tolerance parts than the metal IPV enclosures. 

There are less than one third as many plates in each cell (the SPV 

plates are larger). This reduces the amount of plate preparation, 

speeds cell assembly and helps to decrease the total part count to 

less than half that of an equivalent IPV battery. 

These factors, along with intensive process development, continuous refmement, capital investment and an IS0 9001 style quality system (certification anticipated June, 1998), have helped EPI 

achieve a single shift manufacturing capability of one 22 cell, 60 

Ah SPV battery per week, as well as a testing capacity of 12 finished batteries. EPI has produced 15 Ah, 30 Ah, 50 Ah and 60 Ah, 

27 volt (22 cell) SPV battgies, and the modular nature of the design allows rapid development of capacities up to 120 Ah as well as 

higher or lower voltages. As of this writing, over 60 SPV's have 

been delivered. 

ACKNOWLEDGMENTS 

The authors would like to acknowledge the contributions of 

Lockheed Martin-Sunnyvale, and the assistance of Sharon Wecker 

in the editing, manuscript preparation and publication of this paper. 

CONCLUSIONS 

The Eagle-Picher NiH, SPV battery provides lighter weight, 

smaller volume and better electrical performance than a comparable IPV system. This, combined with high rate production capability, short development time for different configurations and low 

cost make the Eagle-Picher SPV well-suited for satellite energy 

storage. 

REFERENCES 

Coates, D.K., Ansley R.E., Cooper, S.C., Morgan, J.L. and Fox, 

C.L., 1995, "The Design and Manufacture of Nickel-Hydrogen 

Batteries for Spacecraft Power Sytems, " 16th ICSSC, Washington 

D.C. 

Coates, D.K., Grindstaff, B.S. and Fox, C.L., 1995, "Single Pressure Vessel (SPV) Nickel-Hydrogen Battery Design," in Proc. 10th 

Annual Battery Conference on Applications and Advances, California State University, Long Beach, CA. 

'attely Workshop, Huntsville, AL, Nov. 18-20, DC041097.PMd 

Advanced Ni-H, / Nickel Electrode Session .,< 
MECHANICAL BEHAVIOR OF NICKEL OXIDE ELECTRODE UNDER 

SHALLOW CYCLES AND OVERCHARGE 

G. Davolio, E. Soragni, P. Zannini and P. Baraldi 

Department of Chemistry, University of Modena 

183, Via G. Campi, I41 100 Modena, Italy. 

1. Introduction - 

Mechanical effects of cycling on Nickel Oxide Electrode (NOE) have been 

observed for a long time: swelling and bulging, with shedding of active mass are the 

main physical defects, often associated with capacity fading. Studies of the mechanical 

aspects related to the electrochemical reactions on NOE are rather new. Such studies 

were not possible on pocket electrodes, since the external perforated metal strip 

constrains the active mass, so hindering any measurement of that phenomena. They 

can be done, on the other hand, on NOES with an interior current collector, such as 

sintered, metal felt or metal foam supported and plastic bonded electrodes. Some 

physical properties of these electrodes (expansion, elongation, Young's elastic 

modulus), are related to volume changes, associated to phase transitions caused by the 

electrochemical reactions. These studies open a possible way for a better understanding 

of the failure mechanisms, by evaluating the effects of fabrication variables and 

cycling conditions on the electrode internal stresses, therefore giving experimental 

bases for a reliable prediction of the electrode life. 

Several works have been done on this line by various authors; the following 

physical properties have been studied: bending of asymmetric electrodes (1, 2), 

elongation (3), thickness changes (43, electrical resistivity (6), Young's flexural 

elastic modulus (7,8). In Figures 1-4 are reported drawns of some experimental set-up 

of measurements. The main results of these studies are that on cycling of NOE the 

active mass expands during the discharge half-cycle and contracts during recharging 

half-cycle; the stress exerted on the supporting structure (sinter) brings about the 

expansion of the electrode. The expansion-contraction process is only partially 

reversible; the irreversible part being due to the plastic deformations of the supporting 

structure and to some degree of irreversibility of the phase transitions of the active 

material. Some effects of manufacture and cycling variables on NOE expansion have 

1997 NASA Aerospace Battely Workshop Advanced Ni-H, / Nickel Electrode Session 
also been established, confirming the experimental findings that both the increase of 

the level of impregnation or the depth of discharge enhance electrode expansion with a 

negative effect on cycle life (9,lO). - - Sintered NOES show often a more or less evident asymmetry with respect to net 

or perforated sheet acting as the current collector (Fig. 5). In this cases a volume 

expansion of active mass originates a bending moment, which acts against the stiffness 

of the structure and is related to the asymmetry of the electrode. In Fig. 5a is sketched 

the longitudinal section of an asymmetric uncycled electrode, and in Fig. 5b, the same 

section of a discharged Nickel electrode, bent as a result of the active mass expansion. 

The opposite behavior is shown, for example, by an asyrnmemc metal-hydride 

electrode, since its active mass expands during charge and contracts during discharge 

(1 1, 12). In the case of an asymmetric iron electrode we observe a bending during the 

first plateau , corresponding to the reaction: Fe ---> Fe(OH)2. The bending decreases 

during the second plateau, owing to the contraction of the active mass by the reaction 

Fe(OH)2 ---> FeOOH (13). 

By applying the Mechanics of deformable bodies and the theory of cantileverd 

beam (14) to a bent asymmetric electrode, clamped at one end, we have that the 

bending moment M, is balanced by a Force F, applied at a distance L. This force 

cancels the displacement, z, of the free end, brought aboutby the changes of the active 

mass: the following relation holds (Fig. 6): 

z= FL313YI (1) 

where I is the momentum of inertia of the cross section (1=bh3/12), relating to a 

beam cross section of length b and eigth h. 

and Y is the Youngflexicral elastic modulus . If we consider the model of sintered NOE as a composite material, constituted by 

sinter, active mass and voids, (Fig. 7), and apply the principles of Micromechanics , holding for composite materials (14), we obtain: 

Yelectrode = (~s~am/(ys~am+~ds)) (varn+vs) (2) 

where : 

Y is Young's flexural elastic modulus 

V is the Volume fraction of a component 

and the subscripts s and am refer to sinter and active mass respectively. 

If the displacement z is kept constant, the force F, balancing the bending moment, 

depends only on the properties of the active mass (Volume fraction and Modulus), 

which in turn depend on the state of charge, then by combining the relations (1) and (2) 

we have: 

It follows from the equation (3) hat phase transitions andor volume changes of 

the active mass influence the balancing force F, whose changes can be measured by a 

balance in the experimental set-up as reported in the draw of Fig. 8. The electrode E is 

I997 NASA Aerospace Battery Workshop -734- Advanced Ni-Hz/ Nickel Electrode Session 
clamped at one end to the cell and the free end is mechanically connected to a balance 

through a rigid rod and a weigh,^. This weight allows the maintaining a constant 

deflection, z of the electrode and the possibility of recording fie values of the force F 

during a charge-discharge cycle, and measuring the changes in the bending moment of 

the electrode. 

The aim of this work is to find a correlation between mechanical behavior and 

structural changes of the active mass. Another task is to collect data in order to 

establish efficient conditions of accelerated tests for reliable life predictions. The latter 

problem is related to the stability of the sinter structure under the stress produced by 

the active mass changes; in effect the force we measure is a direct measure of the stress 

suffered by the supporting structure during cycling. 

Experimental. 

Owing to the irreversible expansion, the shape of an asymmetric electrode, 

working as a cantilered beam, bound at one end, shows after cycling a permanent 

bending, as sketched in Fig. 9a; the permanent expansion leads also the electrode to 

take the shape represented in Fig. 9b. In order to maintain the original shape of the 

electrode, such as normally occurs in practical cells, we have applied the principle of 

Fig. 9c, accomplished as sketched in Fig. 9d. The electrode is clamped both at the 

extremities: the end with the current collector is clamped at C, fixed to the electrolytic 

cell, the other end is clamped to the frame A, free of pivoting around the axis X: the 

bending moment is equal to F.b. 

By using the electrode assemblage of Fig. 9d, placed in the cell of Fig. 8 we have 

studied the electrochemical and mechanical behavior of Nickel oxide electrodes, with 

these characteristics: 

Type A: chemical impregnation, Co content: 2,9%. 

Type B: electrochemical impregnation, Co content 4,395 

Type C: electrochemical impregnation , Co content 8,295. 

All the tested electrodes had been manufactured on a sintered nickel support. 

The testing procedure was as follows: a smp, 5 cm long and 0.5 cm wide, is cut 

from the asymmemc plate. After the conducting tab is welded, the two extremities are 

masked with teflon tape, leaving it uncovered for a length of 2cm. The asymmemc 

electrode is then placed in the cell, the masked areas are clamped within the frames, as 

in Fig. 9d. The cell is filled with the solution and the level of the balance is lowered in 

order to have a force of some grams with a displacement, z, in the order of tenth of 

rnm. These values depend upon the characteristics of the electrode, the only important 

condition is to maintain the electrode in the elastic range of the supporting structure. 

On the other end one must remember that AF, and not the asbsolute value of F, is of 

interest. 

Once the experiment was started no further change in the dimensional parameters 

of the apparatus was done. The change of the solution, for example, is performed by 

rotating taps without touching the cell. Each cycle of one test, consisting of various 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
runs, was numbered and the cycle number (#) is reported in the leyendu of the 

following figures. 

Three runs of experiments have been performed: i). regular cycles, ii) shallow 

cycles, iii) overcharge. KOH solutions of different concentrations (6 N and 10 N in 

deionized water) and different temperatures (O°C, 20°C and 40°C) have been used in 

the tests. The electrodes were cycled at constant current, with an AMEL galvanostat 

(model. 545) controlled by a timer (in charge and in shallow cycles) or a voltage 

sensor. A Hg/HgO reference electrode and a cadmium counter electrode were used. 

The regular cycle was performed by charging 75 minutes at 1C current ; shallow 

cycles were performed by discharging to 30%-50% DoD, recharging with a 20% 

overcharge of the discharged capacity. The overcharge lasted normally two hours at 

in 

IL. 

All the data refer to discharges following said cycling schedule (regular, shallow 

or-overcharge) and a regular recharge. 

Results and discussion. 

Electrode voltage (Vvs. Hg/HgO ) and F (N-2)values were recorded vs. Time 

during the discharge-charge cycles. The records vs. Time were changed in vs. Capacity 

records, a representation giving a more immediate view of the related physical 

phenomena. For example in Fig. 10a is shown a typical record of Force, Voltage and ' 

Current vs. Time, and in Fig lob the plots of Force and Voltage vs. Capacity. In the 

latter plot the capacity has positive values during charge and negative in discharge. The 

arrows indicate the temporal sequence of experimental points. The origin of the plots, 

zero value Capacity, is represented by the charged electrode, so in Fig lob the 

discharge of the electrode begins at point 0 and proceeds to be completed at point P, 

where the current is reversed and starts the recharge. A first comment on the 

mechanical behavior concerns a "negative hysteresis" in the ForceJCapacity plot: the 

stress slowly increases during the discharge and it is quickly relieved during the frrst 

10-15% of recharging. On the far left of the same plot we observe some negative value 

of Capacity, corresponding to overcharge. In this case overcharge does not bring about 

increase of Force. We will see later that an increase of Force occurs during overcharge 

in some particular conditions (y-phase formation). 

We have studied the NOE behavior during shallow cycling and overcharging, 

owing to the related practical interest, since these cycling conditions often occurs in the 

normal use of nickel-based storage batteries.and produce in discharge the well known 

40-60 mV Voltage depressions, the so called 'memory effect' (15), related to phase 

changes in the active mass (16). 

Several are the variables tested, such as type of impregnation, Co content, KOH 

concentration and temperature. The electrochemical and mechanical parameters one 

can take into account for a comparison are the amount and shape of Voltage depression 

and the dependence of the Force upon the Capacity delivered or accepted by the 

electrode. 

1997 NASA Aerospace Battely Workshop -736- Advanced Ni-Hz/ Nickel Electrode Session 
Some of the most significant results are presented in Figures 11-13, relating to the 

three electrodes in 6N KOH at 20°C. In all the conditions the full discharge following 

shallow cycles shows a voltage depression at the same capacity as that of the repeated 

shallow discharges. The voltage depression is accompanied by an increase in Force, 

with respect to the behavior after regular cycle. This means that higher internal stress is 

suffered by the electrode during the shallow cycling. Two hours overcharge at 1C do 

not produce notable changes on the electrochemical and mechanical behavior of A and 

C type electrodes. A remarkable effect is observed on B type electrode (Fig. 12): it 

appear greatly stressed after overcharge. In discharge the stress remains high for about 

the 30% of the capacity, then it decreases at the end of the discharge. The voltage 

depression appears in discharge and the voltagelcapacity slope is lower than after 

normal charge. 

All these results are in agreement with the hypothesis that y-NiOOH is formed 

during overcharge, as suggested by the well known Bode cycle (Fig.14), and by the 

physical parameters of the Nickel compounds, reported in Table 1. 

Table 1 

Crystallographic parameters of exagonal unit cell-and densities of Nickel 

oxhydroxides, related to NOE. 

alcm-8 clcm-8 Density/g.cm-3 

(reference) (18, 19) (1-8, 19) (20) 

a-Ni(OH)2, 3.13 8-9 2.82 

P-Ni(OH)2 3.13 4.6 3.97 

P-NiOOH 2.82 4.7 4.68 

y-NiOOH 2.82 72 3,79 

The discharge of y-phase gives rise to a-phase of high molar volume, so the stress 

remains unchanged for some time, then the stress goes down due to the transition aNi(OH)2 ----> PNi(OH)2, with molar volume decrease. 

The comparison with the various types of electrodes shows that the factors 

(temperature and KOH concentration) enhancing y-NiOOH formation bring about 

stressed electrodes during overcharge. The electrode type C, for example, did not show 

voltage depression nor mechanical stress in the discharge after overcharge in 6N KOH 

at 20°C (Fig. 13). This electrode becomes susceptible to y-NiOOH formation when 

cycled in 10 N KOH at 20°C, under both shallow cycling and overcharge (Fig. 15). 

The electrode type A, when it is overcharged in 10N KOH at 40°C.during 3.5 

hours shows to be susceptible to y-NiOOH formation, whereas it did not do it when 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
had been overcharged 5 hours at 0°C nor 3.5 hours at 20°C in the same solution: 10 N 

KOH (Fin. 16) 

Rkgsrding to the effect of the method of impregnation-one observes a different 

behavior of chemically impregnated electrode (type A), with respect to the 

electrochemically impregnated one (types B and C): in the regular cycling the former 

shows a force increase as soon as the discharge begins (Fig. 11, 16), the latter shows 

the force increase after the 20-30% DoD has been reached (Figures 12, 13, 15). One 

must note that shallow cycling reduces that DoD (Fig. 12) and the same effect has the 

overcharge in the case of low susceptibility to y-phase formation (Fig. 13, type C, 6N 

KOH, 20°C) The model of sintered NOE of fig 7 can help in explaining the different 

mechanical behaviors: the electrochemical impregnation deposits the active material on 

the walls of the sinter network, so the voids are regularly distributed in the porous 

structure. At the beginning of the discharge these voids con accomodate part of the 

expanded mass; only after that, the further expansion brings about stress on the 

structure and the bending moment arises. In chemically impregnated electrodes the 

active mass is distributed at random into the structure, no well distributed voids are 

present, so that the stress arises as soon as starts.the discharge. The different aspects of 

the surfaces of the electrodes type A and C are shown-on the SEM micrographs 

reported in Figures 17 and 18. 

Conclusions 

- Mechanical stress on the supporting structure; related to phase changes, can be 

detected, in situ during cycling, with the aid of the technique described. 

-- Overcharge originates y-phase, with active mass expansion; further expansion 

may occur in discharge, due to reduction y-NiOOH---->aNi(OH)2: the 

following transition aNi(OH)2----->PNi(OH)2 occurs spontaneously with 

volume contraction. 

-- Repeated shallow discharges also originates y-phase, the stress on the structure, 

developed when the discharge occurs is higher than after regular cycles. This 

stress arises at the same time of the voltage depression and occurs'at the same 

DoD of the shallow cycles. 

--An increase of temperature or concentration enhances the susceptibility to yphase formation in overcharge. 

Aknowledgments. 

This work has been funded by MURST (Minister0 dell'Universith e della Ricerca 

Scientifica), Rome. 

Test electrodes have been supplied by : 

- Eagle-Picher Industries, Joplin (MO) and Colorado springs (CO). 

- Office of Research and Development, Washington PC) 

1997 NASA Aerospace Battery Workshop -738- Advanced Ni-H, / Nickel Electrode Session 
References; 

1) H. S. Lim and S. A. Werzwyvelt, NASA Conference Publication no. 2177, 

175 (1980) - - 2) G. Davolio, E. Soragni, "Power Sources I I ", L.J. Pearce editor, International 

Power Sources Symposium Committee, Leatherhead (U.K.), 1987, p.227 

3) D.H. Fritts, J. Electrochem. Soc., 129 , 118 (1982). 

4) M. Oshitani, T. Takayarna, K. Takashima and S. Tsuji, JAppl.Electrochem. , 16,403 (1986) . 5) K. Micka, J. Mrha and B. Klapste, J. Power Sources, 5,207 (1980) 

6) D. H. Fritts, "Power Sources 7", J. Thompson, editor., Academic Press, 

London, 1981, p.105. 

7) G. Davolio, A. Da Pieve and E. Soragni, Power Sources 12, T. Keily and B.W. 

Baxter, editors, International Power Sources Symposium Committee, 

Leatherhead (U.K.), 1989, p. 253. 

8) L. H. Thaller and T. P. Barrera, Proceedings E.SA. Conference, Florence, 2-6 

Sept., 1991. 

9) H. S. Lim and S. A. Werzwyvelt, Proceedings of 20th ZECEC Conference, Miami Beach, FL, August 1985, p.164. 

10) G. Davolio, A. Da Pieve and E. Soragni, Nickel Hydroxide Electrodes , D.A. 

Comgan and A. H. Zirnrnerman editors, The Electrochemical Society Inc. , Proceedings Volume 90-4 , p. 28 1. 

11) G. Davolio, E. Soragni and C. Fontanesi, in Ext. Abst. of the Spring 1986 

Meeting of the Electrochem. Soc., The Electrochemical Society Inc., 

Pennington, New Jersey, Vol. 96-1, p. 91 

12) A. Pecheron-Guegan, M. Latroche, A.J. Achard, Y.Chabre, J. Bouet, 

Hydrogen and Metal-Hydride Batteries, P. D. Bennet and T. Sakai eds, The 

Electrochem. Soc. Inc., Pennington, New Jersey, Proc. Vol; 94-27, 1994, p. 

196 

13) G. Davolio , A. Da Pieve, P. Baraldi and E. Soragni, Journkes DElectrochimie 

89 , Monpellier (France), Resume des Communications, A 7-12 

14) M. P. Feyman, R. B. Leighton and M. Sands, Lectures in Physics, Addison 

Wesley, Reding, Mass. 1966, Vol. 2, sect. 18-9 

15) G.Weidmann, P. Lewis and N. Reid, editors. "Structural Materials", 

Butterworths, London, 1990, p. 332. 

16) R.N. Thomas, Modern Battery Technology, C.D.S. Tuck, editor, Ellis 

Horwood Ltd, New York 1991, p. 276. 

17) Y. Sato, K. Ito, T. Arakawa, and K. Kobayakawa, J. Electrochem. Soc., 143, 

L225 (1996). 

18) C. Faure, C. Delmas, M. Fouassier, J. Power Sources, 35 (1991) 279. 

19) H. Bode, K. Dehemelt and J. Witte, Zeit. An. All. Chemie, 366 (1969) 1. 

20) H. Bode, K. Dehemelt and J. Witte, Electrochim. Acta 11 (1966) 1079. 

1997 NASA Aerospace Battery Workshop -73 9- Advanced Ni-H, / Nickel Electrode Session 

- 
NICKEL SINTER r ACTIVE MITER 

TIME. min 

H. S. Lim and S. A. Wenwyvelt, NASA Conference pubbl. no 2177 

(1980) p. 175. (Bending of asymmetric electrode) 

Fig. 1 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Countrr - 

Elrctrodr 

Rrfrrrncr Elrctrodr 

Sintcrcd Hi 

G. Davolio. E. Soragni, "Power Sources II", L. J. Pearce editor. 

International Power Sources Symposium Committee, Leatherhead (U.K.), 

1987, p.227 (Bending of asymmetric electrode) 

I997 NASA Aerospace Battery Workshop 

Fig. 2 

Advanced Ni-Hz/ Nickel Electrode Session 
I 

FIXED 

WAU 

U-H I posm DiREciloN oF 

PLATFORM DJSPLACEMEM 

- - 

L. H. Thaller and T. P. Barrera, Proceedings E.SA. Conference, Florence, 

2-6 Sept., 1991.( Elastic modulus of electrodes) - - .- - . . - 

screw J 

1 balance 

G. Davolio. A. Da Pieve and E. Soragni, Power Sources 12, T. Keily and B.W. 

Baxter, editors, International Power Sources Symposium Committee, Leatherhead 

(U.K.), 1989, p. 253. (Elastic modulus of electrodes) 

Fig. 4 

1997 NASA Aerospace Battely Workshop Advanced Ni-H, / Nickel Electrode Session 
Fig. 5. Sketch of the cross-section of asymmetric NOE. 

5a) charged (or uncycled) -- 

5b) discharged. 

Fig. 6 Asymmetric electrode as a Cantikvered beam fixed at one end 

1997 NASA Aerospace Battev Workshop Advanced Ni-Hz/ Nickel Electrode Session 
- 

Fig. 7 Sketch of a model of sintered NOE, as a composite material, 

made by sinter, active mass and voids 

/ BALANCE 1 

Fig.8. Draft of the cell for electrochemical and mechanical tests. 

1997 NASA Aerospace Battery Workshop Advanced Ni-Hz/ Nickel Electrode Session 
Fig. 9 System for maintaining the electrode planarity during cycling. 

I997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Fig. 10 a) Typical record of Force (F) , Vohige (V) and Current (I) vs. 

Time , during a discharge- charge cycle 

Fig. lob) Plot of the data of Fig. 10a vs. Capacity 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Fig. 11. Force and Potential vs. Capacity. type A, 6N KOH, 20°C. 

0 cycle # 135, regular charge, : cycle # 1 14, 19 shallow cycles, 

disch. 30% DOD, 0 : cycle # 132, 11 shallow cycles, disch. 36% 

DOD, A : cycle #116, overcharge 1C 2 hours. 

1997 NASA Aerospace Battery Workshop -747- Advanced Ni-H, / Nickel Electrode Session 
Fig.12. Force and Potential vs. Capacity. type B, 6N KOH, 29°C. 

o : cycle # 118, regular charge, o : cycle # 114,48 shallow cycles, 

disch. 50% DOD, A : cycle #122, overcharge 1C 2 hours. 

1997 NASA Aerospace Battev Workshop Advanced Ni-H, / Nickel Electrode Session 
Fig. 13. Force and Potential vs. Capacity. type C, 6N KOH, 20°C. 

0 : cycle # 190, regular charge, a : cycle # 183, 22 shallow cycles, 

disch. 50% DOD, n : cycle #191, overcharge 1C 2 hours. 

1997 NASA Aerospace Battery Workshop Advanced Ni-Hz/ Nickel Electrode Session 
1997 NASA Aerospace Battely Workshop 

Fig. 14 The Bode's Cycle 

A L A L 

Advanced Ni-H, / Nickel Electrode Session 

discharge charge 

overcharge i 

P-NiOOH c y-NiGOH 

charge discharge 
Fig. 15. Force and Potential vs. Capacity. type C, 10 N KOH, 20°C 

0 : cycle # 213, regular charge, : cycle # 210, 16 shallow cycles, 

disch. 40% DOD, A : cycle #212, overcharge 1C 2-hours. 

1997 NASA Aerospace Battely Workshop Advanced Ni-Hz/ Nickel Electrode Session 
Fig. 16 Force and Potential vs. Capacity. type A, 10 N KOH;: 

0 cycle # 567, O°C., 2hrs. overcharge, V : cycle # 568, O°C., 5 

hrs. overcharge, O. : cycle # 520,20°C. 3.5 hrs. overcharge, 

A : cycle # 558,40°C. 3.5 hrs. overcharge. 

I997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
P 20 pm. 

FIG. 17. Aspect of the surface of Type A electrode. 

- 20 pm. 

FIG. 1 Y. Aspect of the suxface of Type C electrode. 

1997 NASA Aerospace Battely Workshop Advanced Ni-H, / Nickel Electrode Session 
Page intentionally left blank 
Characterization of Electrochemically 

Impregnated Fiber Nickel Electrodes 

D. Coates, J. Francisco, D. Chiappetti and J. Brill 

Eagle-Picher Technologies - Power Subsystems 

Joplin, Missouri 
Possible Approaches to Improve the 

Specific Energy of the Nickel Electrode I) 

High porosity sintered nickel powder 

Heavy loading 

Nickel metal fiber 

Carbon fiber 

Nickel coated graphite fiber 

Carbon composite fiberlpowder 

Pasted electrodes 

Nanostructured materials 

Additives 

Increased electron transfer 
Metal 1 Carbon Fiber Composite 

Nickel Electrode Structures 

W. Ferrando, W. Lee, A. Salkind, D. Hall, B. 

Tatarchuk, DAUG / Acme, et. al. 

1984 - present 

Nickel-coated carbon fibers 

Carbon / stainless steel 

Cellulose mat-type preforms , 

Sintered structures 

Pasted 1 electrochemical impregnation 
Electroplated Nickel on Carbon 

Florida Atlantic University 

- Electroplating process using a proprietary 

plating cell technology 

- Process for the continuous through-plating of 

carbon fiber tows 

- Batch process for electroplating carbon fiber 

mats \ I 

- Very thin, pure contiguous coating of the fiber 
1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
I997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Fiber Electrode Substrates 

Ribbon Technology 

- Melt overflow fiber process 

- Dry air-layered mat process 

Memtec America 

- Fibers produced by drawing process 

- Wet-layering process for making mats 1, l 

National Standard I 

- Fiber 1 powder substrates 
Electrochemical Impregnation Process 

NO3- + 6 H20 + 8 e- = NH, + 9 OHNH, + H20 = NH,+ + OHNi2+ + 2 OH- = Ni(OH)2 

Co2+ + 2 OH- = CO(OH)~ 
SPECIFIC ENERGY vs POROSITY 

SlNTER POROSITY (%) 

LOADING (glcc) 
ACTIVE MATERIAL LOADING (glcc) 
3 

2 ACTIVE MATERIAL LOADING (glcc) 

6' 

3 

ACTIVE MATERIAL LOADING (glcc) 

70 - 

60 - 

50 - 

40 - 

30 

20 - 

10 - 

0 

0 

0 0 

0 

0 0 

0 

0 0 0 

0 

O 0 0 0 00 

0 

00 O 0 

o 

0 00 0 

0 0 

Oo O 0 

0 0 5 0 

0 0 0 000 0 

0 0 0 0 

00 o o 

0 

0 

0 0 0 Q0 0 

0 0 0 

8 0 Q Oo 0 

0 O 0 0 

0 O% 0 : O 

0 0 

0 \ o FORMATION 

o IMPREGNATION 

I I I I 1 I I I I I t 

0.6 0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 
SAMPLE NUMBER 
30 40 50 60 70 

SAMPLE NUMBER 
o RIBTEC 

o MEMTEC 

30 40 

SAMPLE NUMBER 
SAMPLE NUMBER 
SAMPLE NUMBER 
a 

* * 

a 

a 

a a 

a l a l 

a a * 

0 

l a. ** a a 

0 *+o 0 

0 

a 

** 0 

0 oO 0 0 0 

O 0 00 0 

@O a 

0 ** 8 O + 0 

Oo l 0 0.0 00 0 0 

0 

0 l 0 0 ** 0 0.0 0 

0 0 0 O0 

0 0 00 0 0 0 

0 0 

0 04 * 0 

O* + 

0 0 O 0 <: ORMATI% MA TI% a RIBTEC IMPREGNATION @ o o o o o 

l 

0 MEMTEC FORMATION I 

+ MEMTEC IMPREGNATION 

*. - -- 

1'3 I #I! I I III I I I I I ,,,,I I,II,,(, 

SAMPLE NUMBER 
Conclusions 

0 

Sintered powder electrode designs are limited to 

approximately 140 mAh/g maximum 

Fiber / graphite composite electrodes are limited 

to approximately 1 8 5 mAWg maximum 

A variety of trends have been identified relating 

electrode design parameters such as fiber 

construction type, porosity and loading to 

electrode performance parameters such as active I 

material utilization, specific energy, energy 

density and thiclsness growth 
- 

Nickel Electrode Corrosion During 

Nominal On-Orbit Operation 

Chuck Lurie and Bob Tobias 

TRW Space and Electronics Group 

Redondo Beach, California 90278 

The 1997 NASA Aerospace Battery Workshop 

The Huntsville Hilton 

Huntsville, Alabama 

November, 17 - 20,1997 

1997 NASA Aerospace Battery Workshop -777- Advanced Ni-H, / Nickel Electrode Session 
Background 

The increase in capacity often observed during early 

nickel-cadmium battery mission life is attributed to 

corrosion of nickel electrode sinter forming active nickel 

hydroxide. 

* An example of this phenomenon is examined using the 

20-year, on-orbit history of the FLTSATCOM F1 spacecraft. 

At 20 years into the mission: 

- Battery eclipse discharge performance continues to 

be nominal. - 

- Reconditioning discharge capacity increased during 

the first 6 - 7 years of operation, plateaued, and then 

started a continuing decrease. 

The FLTSATCOM F1 spacecraft provides an excellent opportunity to examine the 

phenomenon of nickel electrode sinter corrosion using on-orbit data. Extensive eclipse 

season and reconditioning operating data has been collected throughout its 20-year 

mission life. 

1997 NASA Aerospace Battery Workshop Advanced Ni-Hz/ Nickel Electrode Session 
. The Spacecraft and Mission 

~eos~nchronous orbit communications satellite 

Eclipse power provided by three 24-cell, %&Ah, NiCd 

batteries 

Cell characteristics: 

- Rated capacity 24 Ah 

- Measured capacity 28.5 Ah 

- Number of plates 1 1 positive, 12 negative 

- Plate loading Positive 13.4 + 0.6 gmsldfi 

Cd treatment 

Negative %.I f 0.6 gmsldrh 

Ag treatment 

- Neg lPos ratio I .7:1 

- Electrolyte quantity 68 ml, 34% KOH 

The mission of FLTSATCOM is to provide a satellite communication capability that is 

shared by the United States Navy and Air Force. 

During eclipse, spacecraft power is provided by three 24-cell nickel-cadmium batteries. 

The batteries are connected to the main spacecraft bus, during charge, through a diode 

path. When the spacecraft enters into eclipse, relays are switched to connect the 

batteries directly to the bus, which is powered by the spacecraft solar array. 

Each battery weighs approximately 65 pounds and has envelope dimensions of 15.5 by 

1 1.7 by 8.5 inches. The cells were manufactured by General Electric Battery Division in 

Gainesville, Florida. 

1997 NASA Aerospace Battery Workshop -779- Advanced Ni-H, / Nickel Electrode Session 
Battery Operation 

* DOD 

- 75% of rated capacity - - 

- 63% of measured capacity 

Charge 

- Charge at Cll5 - CIIO rate to a recharge ratio of - 1.0 

- Trickle charge at -Cl100 rate untii following discharge 

Reconditioning 

- Immediately prior to each eclipse season 

- Discharge into 85.7 ohm load to 18 volts (-100 hours) 

- Provision for manual termination - 

Temperature 

- 0°C to 10°C 

Battery depth-of-discharge (DOD) has remained close to 63% of measured capacity (75% 

of rated capacity) during on-orbit eclipse season operations. 

Battery charging consists of a full rate charge period followed by trickle charge. During 

full charge, current is provided by the battery charge section of the solar array. Full 

charge current is limited to approximately 2.1 amperes by the current limiting 

characteristics of the charge section of the solar array. Full charge is terminated when the 

battery voltage reaches a preselected temperature compensated voltage limit. The 

charge current is then reduced to a trickle charge rate of 0.24 amperes and maintained on 

trickle charge for the remainder of the sunlight period. 

Reconditioning is performed prior to each eclipse season. The batteries are discharged 

through 85.7 ohm resistors. Discharged is terminated automatically when the battery 

reaches 18 volts. The reconditioning discharge can also be terminated manually at lower 

voltages. 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
History: Eclipse Performance 

Eclipse Day 

The data presented here summarizes the eclipse season performance of the spacecraft. 

Performance is tracked by observing the trends in end of discharge voltage from season 

to season. The data shows a gradual decrease in end of discharge voltage during the 

mission. However, the end of discharge voltages remain at satisfactory levels. 

199 7 NASA Aerospace Battery Workshop Advanced Ni-Hz/ Nickel Electrode Session 
History: - Reconditioning Discharge 

\ : I -Battery 1 Capacity (Ah) ----- \:I 

?' - - Battery 2 Capacity (Ah) :t i ........... : ------.------: ---------.--- L ------------- L -------------: ---- I IIIIII-; 

it i 

- - Battery 3 Capacity (Ah) -I: b b b i b i 

8 10 12 

Time On Orbit (years) 

This chart shows the reconditioning discharge capacity history for FLTSATCOM F1 

mission from inception. During the first eight years capacity is increasing and individual 

battery capacities are very close. After a short plateau the capacities begin to decrease 

and dispersion begins to increase. It is interesting to note that, while the batteries 

continue to support the eclipse season requirements with margin, the reconditioning 

discharge data suggests that the batteries are near end of life. 

I997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Reconditioning - Discharge Profiles 

Discharge Duration (hours) 

As the mission progresses the reconditioning discharge profiles continue to have smooth 

knees indicating that cell capacities, within a battery, are not diverging significantly. This 

is an interesting observation because cell capacity dispersion, due to the presence of soft 

shorts, is frequently observed as cells age. 

1997 NASA Aerospace Battely Workshop Advanced Ni-H, / Nickel Electrode Session 
Reconditioning - Capacity Dispersion 

0 2 4 6 8 10 12 14 16 18 20 

TIME (years) 

Capacity dispersion is defined as the standard deviation of the capacities of the three 

batteries. As can be seen from the chart, dispersion is small during the first six to seven 

years, when the capacity is increasing, and becomes increasingly larger, during the 

remaining years as the capacity decreases. The small dispersion, initially, reflects careful 

cell matching. The trajectory of the dispersion growth suggests a mechanism that is 

common to all cells in all batteries and is well behaved. 

1997 NASA Aerospace Battely Workshop Advanced Ni-Hz/ Nickel Electrode Session 
Capacity Growth - Ni Corrosion at the Current Collector 

The rate of capacity increase is stable at approximately 

0.34 Ahlyear, during the first 6 -7 years of operation. 

Capacity growth is assumed to be due to the formation of 

NiOOH by corrosion of Ni electrode sinter during 

overcharge. 

0.34 Ah of additional NiOOH per year suggests that 0.75 

gms of the Ni sinter is consumed per cell per year. 

This rate of Ni loss is equivalent to 8% of the Ni electrode 

sinter mass, at year 10. 

Corrosion of the nickel electrode sinter, during overcharge, can occur uniformly during 

early life because of the inherent uniformity of the cell plate stacks. As more and more 

sinter is corroded the plates become less uniform and increased capacity dispersion can 

occur. The amount of sinter corroded, approximately 8% at year 10, appears to be 

sufficient to significantly alter the interface between the active material and the current 

collector . 

I997 NASA Aerospace Battery Workshop -785- Advanced Ni-H, / Nickel Electrode Session 
Prediction of Remaining Life 

The reconditioning capacity trend curves provide a signature 

describing processes which limit battery life. 

This signature may be useful in determining remaining life, at 

some point in the mission. 

This hypothesis has been tested. 

- Assuming a Weibull reliability model 

- Shape parameter, P, assumed to be 8 

- Scale parameter, 6, estimated from the orbital -. data 

Spacecraft and electrical power system designers are usually concerned with battery 

system performance and life. Most of the related effort reported has focused on 

determining battery life as a function of battery and mission characteristics. The 

FLTSATCOM F1 battery reconditioning performance data base provides an opportunity to 

explore a different aspect of life prediction. The hypothesis that the reconditioning 

discharge history may be useful in determining remaining life was tested using a twoparameter Weibull reliability model. The Weibull distribution has been used by many 

investigators and is generally considered to be appropriate for battery reliability modeling. 

1997 NASA Aerospace Battely Workshop Advanced Ni-H, / Nickel Electrode Session 
Weibull Reliability Model 

0 5 10 15 20 25 30 35 40 45 

Mission Duration (years) 

The two-parameter Weibull distribution was chosen as being the distribution most 

representative of the aerospace battery data and experience base. 

The most familiar distribution is the normal (Gaussian) distribution. Unfortunately battery 

lifetime data is markedly skewed and not bell shaped. Also the normal distribution, 

applied to aerospace battery lifetime data, is mathematically difficult complicating 

determination of failurelhazard rates. The log-normal distribution is more convenient and 

has been used by others. However its hazard rate increases to a maximum and then 

decreases in the long-life range, whereas the normal distribution has a continuously 

increasing hazard rate. The Weibull distribution is convenient mathematically and has an 

increasing hazard rate that, with the appropriate choice of Weibull parameters, faithfully 

models the historical aerospace battery data base. 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Battery Life, Predicted at Year 12 

RB = RC~~ + 24RcZ3(1- RC) 

0 5 10 15 20 25 30 

Mission Duration (years) 

The hypothesis that remaining life can be predicted with the Weibull model was tested at 

the twelve-year point, in the FLTSATCOM F1 mission. The $ variable was assumed to 

be 8. This value assumes that failures are the result of wearout and that random failures 

contribute insignificantly to the hazard rate. Values of 6 equivalent to individual battery 

reliabilities of 0.97 and 0.99, at year 12, were calculated. These are considered 

conservative in that the three batteries were functioning nominally at year 12. The 

resulting PDF's are shown on the previous chart. Battery reliability is calculated, from cell 

reliabilities, using a binomial model, assuming one cell is redundant. As can be seen from 

the curves above the model predicts that reliability is high for a twenty year mission, and 

falls rapidly thereafter. This prediction is in good agreement with conclusions drawn, 

using engineering judgment, from observation of on-orbit battery performance and 

reconditioning data. 

1997 NASA Aerospace Battery Workshop Advanced Ni-H, / Nickel Electrode Session 
Summary 

Data has been presented which tracks NiCd battery 

reconditioning discharge capacity during the continuing 

20-year FLTSATCOM F1 mission. 

The data trend describes a reconditioning discharge capacity 

growthldecay signature. The capacity signature is postulated 

to be the result of corrosion of the sinter, in the Ni electrode, 

forming NiOOH. 

The hypothesis that this reconditioning discharge signature 

can be used to predict remaining life was tested using a 

Weibull reliability model. 

- Predictions are consistent with observations. 

- A larger data base is needed before this approach can 

be validated. 

I997 NASA Aerospace Battery Workshop -789- 

-- - 

Advanced Ni-Hz/ Nickel Electrode Session 
Page intentionally left blank 
199 7 NASA Aerospace Battery Workshop Attendance List 

- Walter Alsbach 

KRIS Engineering Inc. / J & T 

GSFC Code 734 

Greenbelt, MD 20771 

Ph: (301) 286-4822 Fa: 

Chris Baker 

Acme Electric Corporation 

528 W. 21st 

Tempe, AZ 85282 

Ph: (602) 894-6864 Fax: (602) 921-0470 

E-mail: cbaker@acmepower.com 

Bob Bechtel 

Marshall Space Flight Center 

EBIl 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3294 Fa: (205) 544-5841 

Samuel Birken 

The Aerospace Corporation 

MS M41986 

POB 92957 

Los Angeles, CA 90009-2957 

Ph: (310) 336-6080 Fax: (3 10) 336-5581 

E-mail: samiam@lafn.org 

Rita Brazier 

Marshall Space Flight Center 

EBll 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3295 Fa: (205) 544-5841 

Harry Brown 

Naval Surface Warfare Center - Crane Div. 

Code 6095 B2949 

300 Highway 361 

Crane, IN 47522 

Ph: (812) 854-6149 Fax: (812) 854-3589 

E-mail: brown-hany@crane.navy.mil 

Shirley Butler 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3298 Fax: 

1997 NASA Aerospace Battery Workshop 

Jon D. Armantrout 

Lockheed Martin 

OIE201 -B/154 

11 11 Lockheed Way 

Sunnyvale, CA 94089 

Ph: (408) 742-1800 Fax: (408) 742-273 1 

Doug Bearden 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Charles W. Bennett 

Lockheed Martin Astro Space 

230 E. Mall Blvd. 

King of Prussia, PA 19406-2995 

Ph: (610)354-6155 Fa: (610)354-3434 

E-mail: ebennett@voicenet.com 

Yannick Borthomieu 

SAFT Advanced Batteries 

Rue G. Leclanche 

BP 1029 

86060 Poitiers Cedex 

France 

Ph: 33-549-55-4014 Fax: 33-549-55-5620 

E-mail: yannick.borthomieu@saft.alcatel-alsthorn.fr 

Jeff?ey C. Brewer 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3345 Fax: (205) 544-5841 

E-mail: jeff.brewer@msfc.nasa.gov 

Ramakumar Bugga 

Jet Propulsion Laboratory 

MS 277-207 

4800 Oak Grove Dr. - Pasadena, CA 9 1 109 

Ph: (818) 354-01 10 Fax: (818) 383-6951 

E-mail: kumar@hyperfine.caltech.edu 

Dwight Caldwell 

Eagle-Picher Industries, Inc. 

1216 W. C St. 

Joplin, MO 64801 

Ph: (41 7) 623-8000 Fax: (417) 623-53 19 

E-mail: pwrsubsys@epi-tech.com 

Attendance List 
199 7 NASA Aerospace Battery Workshop Attendance List fl-= 

- John E. Casey 

Lockheed Martin Space Mission Sys & Services 

2400 NASA Rd. 1, EP5 

Houston, TX 77058-3799 

Ph: (281) 483-0446 Fax: (281) 483-3096 

E-mail: john.e.casey I@jsc.nasa.gov 

Dwaine Coates 

Eagle-Picher Industries 

"C" and Porter Streets 

Joplin, MO 64801 

Ph: (417) 623-8000 Fax: 

Dr. Marsha Daman 

Lockheed Martin Missiles and Space 

Org. E6-01; Bldg 149 

11 11 Lockheed Martin Way 

Sunnyvale, CA 94089 

Ph: (408) 756-0749 Fax: (408) 756-0645 

E-mail: marsha.daman@lmco.com 

Jeff Dermott 

Eagle-Picher Industries, Inc. 

Joplin, MO 64802 

Ph: (417) 623-8333 Fax: (417) 623-0233 

Juergen Eckert 

Los Alamos National Laboratory 

Lujan Center, MS H805 

Los Alamos, NM 87545 

Ph: (505) 665-2374 Fax: (505) 665-2676 

E-mail: juergen@lanl.gov 

Beth Fleming 

The Boeing Company 

689 Discovery Drive 

Huntsville, AL 35806 

Ph: (205) 922-7359 Fax: (205) 922-7525 

E-mail: beth.fleming@boeing.com 

Donald Georgi, P.E. 

Batteries Digest Newsletter 

1261 Town Line Road 

Maple Plain, MN 55359 

Ph: (612) 479-6190 Fax: (612) 479-3657 

E-mail: teksym@aol.com 

I997 NASA Aerospace Battery Workshop 

Marie-Pierre Cassagne 

ALC ATEL 

av champollion 

Toulouse Cedex 

France? l.037 

Ph: (33) 5-61-19-62-58 Fax: (33) 5-61-19-61-69 

William D. Cook 

Eagle-Picher Industries, Inc. 

C and Porter Streets 

Joplin, MO 64801 

Ph: (417) 623-8000 Fax: 

Giovanni Davolio 

University of Modena 

Department of Chemistry 

183, Via G. Campi 

Modena, Italy 1-41 100 

Ph: 39-59-378461 Fax: 39-59-373543 

E-mail: davolio@c220.unimo.it 

Sal Di Stefano 

Jet Propulsion Laboratory 

MS 277-212 

4800 Oak Grove Drive 

, Pasadena, CA 9 1 109 

Ph: (818) 354-6320 Fax: (818) 393-6951 

Ted Edge 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3381 Fax: 

Pete George 

Marshall Space Flight Center 

EB12 

Marshall Space Fliglit Center, AL 35812 

Ph: (205) 544-3331 Fax: 

Shirley Georgi 

Batteries Digest Newsletter 

1261 Town Line Road 

Maple Plain, MN 55359 

Ph: (612) 479-6190 Fax: (612) 479-3657 

E-mail: teksym@aol.com 

Attendance List 
1997 NASA Aerospace Battery Workshop Attendance List 

- Charles Hall 

MarshallSpace Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3330 Fax: 

Steve Hall 

Naval Surface Warfare Center -Crane Div. 

Code 6095 

300 Highway 361 

Crane, IN 47522 

Ph: (812) 854-4172 Fax: (812) 854-3589 

David Hall 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) -544-42 15 Fax: 

Gerald Halpert 

Jet Propulsion Laboratory 

MS 277-207 

4800 Oak Grove Dr. 

Pasadena, CA 91 109 

Ph: (818) 354-5474 Fax: (818) 393-6951 

E-mail: gerald.halpert@jpl.nasa.gov 

Jeff Hayden Susan M. Herczeg 

Eagle-Picher Industries, Inc. National-Standard Co. 

3820 South Hancock Expressway 240 1 M Home St 

Colorado Springs, CO 8091 1 Mishawaka, IN 46545 

Ph: (719) 392-4266 Fax: (719) 392-5103 Ph: (219) 259-8505 x257 Fax: (219) 259-8507 

Albert Himy 

Navy / JJMA 

2341 Jefferson Davis Highway 

Suite 715 

Arlington, VA 22202 

Ph: (703) 41 8-4257 Fax: (703) 41 8-4269 

Charles Huang 

Lockheed Martin 

3700 Bay Area Blvd. 

Houston, TX 77058 

Ph: (281) 335-6036 Fax: (281) 335-6284 

E-mail: huang@ips.jsc.nasa.gov 

Nathan Isaacs 

Mine Safety Appliances Company 

38 Loveton Circle 

Sparks, MD 21 152 

Ph: (410) 628-5432 Fax: (410) 628-0613 

E-mail: nedizakmsa@aol.com 

Jason E. Jenkins 

Johns Hopkins University / APL 

MS 23-2 14 

Johns Hopkins Rd. 

Laurel, MD 20723-6099 

Ph: 41 0 792-5 106 Fax: 

1997 NASA Aerospace Battery Workshop 

Scott Hoover 

Orion Satellite Corporation 

2440 Research Blvd. 

Suite 160 

Rockville, MD 20850 

Ph: (301) 258-3325 Fax: (301) 258-3319 

E-mail: shoover@onsi.com 

Scott Hutchison 

Lockheed Martin SMS&S 

16100 Space Center Blvd. # 1 12 

Houston, TX 77062 

Ph: (281) 335-6464 Fax: (281) 335-6400 

E-mail: hutchison@mscips.jsc.nasa.gov 

Lorna Jackson 

Marshall Space Flight Center 

EB12 

Marshall Space Flight-Center, AL 35812 

Ph: (205) 544-33 15 Fax: 

Grady S. Jobe 

Marshall Space Flight Center 

EAOl 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-0027 Fax: 

Attendance List 

- 
1997 NASA Aerospace Battery Workshop Attendance List 

- Walter Kaden 

Lockheed Martin Missiles and Space 

Mail Stop 419-3A 

POB 800 

Princeton, NJ 08543-0800 

Ph: (609) 490-23 10 Fax: (609) 490-2856 

E-mail: walter.kaden@lmco.com 

Marcie Kennedy 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3724 Fax: 

Randolph Leising 

Wilson Greatbatch Ltd. 

10000 Wehrle Dr. 

Clarence, NY 1403 1 

Eric Lowery 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-0080 Fax: 

Chuck Lurie 

TRW 

MS R411082 

One Space Park 

Redondo Beach, CA 90278 

Ph: (310) 813-4888 Fax: (310) 812-4978 

E-mail: chuck.lurie@tnv.com 

Dr. Tyler X. Mahy 

U.S. Government 

c/o OTS, NHB 

Washington, DC 20505 

Ph: (703) 874-0739 Fax: (703) 641-9830 

Lynn Marcoux 

BlueStar Advanced Technology Corporation 

1155 West 15th Street 

North Vancouver, B.C. 

Canada V7P 1M7 

Ph: (604) 986-4020 Fax: (604) 986-1048 

1997 NASA Aerospace Battev Workshop 

Chad 0. Kelly 

Eagle-Picher Industries 

C & Porter 

Joplin, MO 64801 

Ph: (417)623-8000 ~436 Fax: (411) 623-0850 

E-mail: eplion@aol.com 

Donald R. Kleis 

ENSER 

227 Silverwood Ln. 

Hazel Green, AL 35750 

Ph: (205) 828-0569 Fax: (205) 890-6707 ~247 

E-mail: slvrstreek@aol.com 

Danny Liu 

Intelsat 

Intelsat-WI Program Office 

C/O Lockheed Martin Astrospace 

POB 800 / MS 93 

Princeton, NJ 08543-0800 

Ph: (609) 426-9225 Fax: (609) 426-8968 

E-mail: danny.liu@intelsat.int 

Dave Lucero 

Eagle-Picher Industries, Inc. 

3820 South Hancock Expressway 

Colorado Springs, CO 8091 1 

Ph: (719) 392-4266 Fax: (719) 392-5103 

Michael Mackowski 

The Boeing Co. 

1022 W. Juanita Ave. 

Gilbert, AZ 85233 

Ph: (602) 732-3064 Fax: (602) 732-3885 

E-mail: 71571.33O@compuserve.com 

Michelle Manzo 

Lewis Research Center 

MS 309-1 

2 1000 Brookpark Rd:- 

Cleveland, OH 44135 

Ph: (216) 433-5261 Fa: (216) 433-6160 

Dr. Nehemiah Margalit 

Tracor Applied Sciences Inc. 

1601 Research Blvd. 

Rockville, MD 20850-3 176 

Ph: (301) 838-6223 Fax: (301) 838-6222 

E-mail: nmargalit@tracor.com 

Attendance List 
199 7 NASA Aerospace Battery Workshop Attendance List 

Richard harsh 

AFRL /.PRPB 

1950 Fifth St. 

Wright-Patterson AFB, OH 45433 

Ph: (937) 255-7770 Fax: (513) 255-1759 

E-mail: marshra@wl.wpafb.af.mil 

Dean W. Maurer 

Loral Skynet 

292 Johnston Dr. 

Watchung, NJ 07060 

Ph: (908) 753-4457 Fax: (732) 949-8082 

E-mail: dwrnaurer@attmail.com 

Kurt McCall 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 961-4501 Fax: 

Martin Milden 

2212 So. Bevenvil Dr. 

Los Angeles, CA 90034-1 034 

Ph: (310) 836-7794 Fax: (310) 836-4494 

E-mail: mmilden@worldnet.att.net 

Thomas R. Newbauer 

SMC / AXEN 

SMC / AXE 

160 Skynet St. 

Los Angeles, CA 90245-4683 

Ph: (310) 363-5176 Fax: (310) 363-2532 

E-mail: newbauertr@post6.laafb.af.mil 

Burton Otzinger 

Jet Propulsion Laboratory 

4800 Oak Grove Drive 

Pasadena, CA 9 1 109 

Ph: (626) 354-235 1 Fax: 

Rick Parmley 

LaBarge Electronics 

1505 Maiden Lane 

POB 1210 

Joplin, MO 64802-121 0 

Ph: (417) 781-3200 x258 Fax: (417) 781-2175 

Jeff Martin 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Louis C. Maus 

Marshall Space Flight Center 

PDll 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-0484 Fax: (205) 544-4225 

George Methlie 

2120 Natahoa Ct. 

Falls aurch, VA 22043 

Ph: (202) 965-3420 Fax: (703) 641-9830 

Thomas L. Miller 

Barrios Technolou, Inc. 

MS DF75 

NASA Johnson Space Center 

Houston, TX 77058 

Ph: (281) 244-0048 Fax: (281) 244-003 1 

E-mail: thomas.l.millerl@jsc.nasa.gov 

Pat O'Donnell 

Lewis Research Center 

MS 309-1 

21000 Brookpark Rd. 

Cleveland, OH 44135 

Ph: (216) 433-5248 Fax: (216) 433-6160 

E-mail: seodo@lims02.lerc.nasagov 

John Pajak 

ZAE Research, Inc. 

555 Sparkman Dr. 

Suite 450 

Huntsville, AL 35516 

Ph: (205) 890-0507 Fax: (205) 890-0857 

Dave Perrone 

Jet Propulsion Laboratory 

MS 277-212 

4800 Oak Grove Dr. 

Pasadena, CA 91 109 

Ph: 818 354-9044 Fax: 818 393-6951 

1997 NASA Aerospace Battery Workshop -795- Attendance List 
1997 NASA Aerospace Battery Workshop Attendance List 

- David F. Pickett 

Eagle-Piher Industries 

Power Systems Department 

3820 S. Hancock Expressway 

Colorado Springs, CO 8091 1 

Ph: (719) 392-4266 Fax: (719) 392-5103 

E-mail: aaaa-energy-enterprises@msn.com 

Allen R. Powers 

Hughes Space & Communications 

POB 2999 

3 100 Lomita Blvd. 

Bldg 231 1 MS 1918 

Torrance, CA 90509-2999 

Ph: (3 10) 5 17-5435 Fax: (310) 5 17-7676 

E-mail: arpowers@ccgate.hac.com 

David L. Randall 

Johnson Space Center 

Mail Code DF6 

Houston, TX 77058 

Ph: (281) 483-0746 Fax: (281) 483-3997 

E-mail: david.l.randalll@jsc.nasagov 

Paul Repiogle 

USBI 

POB 21212 

Mail Stop USB-EE 

Kennedy Space Center, FL 32815-0212 

Ph: (407) 867-7871 Fax: (407) 867-9829 

E-mail: reploglep@usbi.com 

Ricky J. Roberson 

Huntsville Integated Voltaic Equipment 

21 15 Britain Ave. 

Huntsville, AL 35803 

Ph: (205) 512-3001 Beeper Fax: 

E-mail: ricky.roberson@rnsfc.nasa.gov 

Frank Scalici 

INTELSAT 

3400 International Dr. NW 

Box 34 

Washington, DC 20008 

Ph: (202) 944-7353 Fax: (202) 944-7333 

E-mail: frank.scalici@intelsat.int 

Darren L. Scoles 

Lockheed Martin Missiles & Space 

GPS IIR Flight Operations 

440 Discoverer St. 

Suite 79 

Falcon AFB, CO 80912-4479 

1997 NASA Aerospace Battery Workshop 

Darren Pivarnik 

Orion Satellite Corporation 

2440 Research Blvd 

Suite 160 

RockviHe, MD 20850 

Ph: (301) 258-3219 Fax: (301) 258-3319 

E-mail: 695-4848@rncimail.wm 

Dan T. Radzykewycz 

AFRL / VSDVP 

3550 Aberdeen Ave. SE 

Kirtland AFB, XM 871 17 

Ph: 505 846-5703 Fax: 

Paul L. Reinstein 

The Aerospace Corporation 

2350 E. El Segundo Blvd. 

MS M41179 

El Segundo, CA 90245-4691 

Ph: (310) 336-1909 Fax: (310) 336-5581 

E-mail: paul.l.reinstein@aero.org 

Ron Repplinger 

Eagle-Picher Industries, Inc. 

POB 47 

Joplin, MO 64802 

Ph: (417) 623-8000 Fax: (417) 623-53 19 

Paul S. Rodriguez 

ENSER 

5 12 Bain Dr. 

Huntsville, AL 35803 

Ph: (205) 882-3465 Fax: 

Charles Scogins 

ACM Inc. 

3054 D Leeman Feny Rd. 

Huntsville, AL 3580T 

Ph: (205) 881-5493 Fax: (205) 883-5885 

E-mail: ACM@traveller.com 

Annie Sennet 

SAFT 

156 Avenue DeMetz 

Romainville, France 93230 

France 

Ph: 33 1 49 15 36 44 Fax: 33 1 48 91 95 53 

Attendance List 
1997 NASA Aerospace Battery Workshop Attendance List 

- 

Sat P. Singhal 

Computer Sciences Corporation 

7404 Executive Place 

Seabrook, MD 20706 

Ph: 301 464-7489 ~7579 Fax: 301 464-3352 

E-mail: ssinghal@csc.com 

Yoshitsugu Sone 

National Space Development Agency of Japan 

Tsukuba Space Center 

2-1 -1 Sengen, Tsukuba-city 

Ibaraki-pref. 305 Japan 

Ph: 81-(0)298-52-2285 Fax: 81-(0)298-52-2299 

E-mail: sone.yoshitsugu@nasda.go.jp 

James A. Stepro 

National-Standard Co. 

2401 N. Home St. 

Mishawaka, IN 46545 

Ph: (219) 259-8505 Fax: (219) 259-8507 

Joe Stockel 

Office of Research & Development 

LF-7 Rm 61 18 

Washington, DC 20505 

Ph: (703) 613-8833 Fax: (703) 613-8954 

Rao Surampudi 

Jet Propulsion Laboratory 

4800 Oak Grove Dr. 

Pasadena, CA 91 109 

Ph: (818) 354-0352 Fax: (818) 393-6951 

E-mail: rao.surampudi@pl.nasa.gov 

Lawrence Thaller 

The Aerospace Corporation 

MS M21275 

POB 92957 

Los Angeles, CA 90009 

Ph: (310) 336-5180 Fax: (310) 336-1636 

E-mail: lawrence.h.thaller@aero.org 

Mark R. Toft 

Motorola 

44330 Woodridge Parkway 

Leesburg, VA 201 76 

Ph: (703) 724-8188 Fax: (703) 724-8156 

E-mail: mark_toft-F77760@email.mot.com 

Gerald S. Smith 

INTELSAT 

3400 Intemational Dr., NW 

Box 34 

Washingten, DC 20008-3098 

Ph: (202) 944-8194 Fax: (202) 944-7333 

E-mail: gerald.smith@intelsat.int 

David M. Spillman 

Wilson Greatbatch Ltd. 

10000 Wehrle Drive 

Clarence, NY 1403 1 

Ph: (716) 759-5378 Fax: (716) 759-8579 

E-mail: dspillman@greatbatch.com 

Tina D. Stewart 

United Space Alliance 

MS Dm5 

NASA JSC 

Houston, TX 77058 

Ph: (281) 483-8900 Fax: (281) 244-003 1 

E-mail: tina.d.stewartl@jsc.nasa.gov 

Bradley Strangways 

Symmetry Resources, Inc. 

POB 785 

Arab, AL 35016 

Ph: (205) 586-891 1 Fax: (205) 586-9443 

E-mail: symm@bmtc.mindspring.com 

Vijay V. Thakur 

Defence Evaluation and Research Agency 

Space Power Laboratory 

XI07 Bldg. 

DERA Farnborough 

Hampshire. GU140LX.U.K. 

Ph: 01 1-44-1252 393669 Fax: 01 1-44-1252 392664 

Robert F. Tobias 

TRW 

MS R411074 

One Space Park - 

Redondo Beach, CA 90278 

Ph: (310) 813-5784 Fax: 

Greta Tracinski 

Applied Power International 

1236 N. Columbus Ave., #41 

Glendale, CA 91202-1 672 

Ph: (81 8) 243-3 127 Fax: (81 8) 243-3 127 

E-mail: 104024.1452@compuserve.com 

I997 NASA Aerospace Battery Workshop Attendance List 
199 7 NASA Aerospace Battery Workshop Attendance List 

- Walter A. Tracinski 

Applied Power International 

1236 N. Columbus Ave., #41 

Glendale, CA 91202-1 672 

Ph: (818) 243-3 127 Fax: (818) 243-3127 

E-mail: 104024.1452@compuse~e.com 

Daniel J. Vonderwell 

The Boeing Company 

MS AHSV-33B1 

689 Discovery Drive 

Huntsville, AL 35806-2804 

Ph: (205) 922-6739 Fax: (2050 922-7525 

E-mail: daniel.j.vondenvell@boeing.com 

Steven C. Walker, P.E. 

The ENSER Corporation 

UAH Johnson Research Center 

Huntsville, AL 35899 

Ph: (205) 890-6707 ~227 Fax: (205) 890-6668 

E-mail: royjacxon@aol.com 

Harry Wannemacher 

Jackson & Tull 

7375 Executive Place 

Suite 200 

Seabrook, MD 20706 

Ph: (301) 805-6090 Fax: (301) 805-6099 

E-mail: harry.wan@jnt.com 

Steve Westfall 

LaBarge Electronics 

1505 Maiden Lane 

Joplin, MO 64802 

Ph: (417) 781-3200 Fax: (417) 781-2175 

Brett Williams 

EUTELSAT 

70 rue Balard 

75502 PARIS CEDEX 15 

Ph: 33.1.53.98.46.43 Fax: 33.1.53.98.49.99 

Doug Willowby 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3334 Fax: 

I997 NASA Aerospace Battery Workshop 

Dr. Hari Vaidyanathan 

CQMSAT Laboratories 

22300 Comsat Dr. 

Clarksburg, MD 20871 

Ph: (304)428-4507 Fax: (301) 428-3686 

E-mail: hari.vaidyanathan@comsat.com 

Sammy Waldrop 

Symmetry Resources, Inc. 

POB 785 

Arab, AL 35016 

Ph: (205) 586-891 1 Fax: (205) 586-9443 

E-mail: symm@bmtc.mindspring.com 

Bryan Walls 

Marshall Space Flight Center 

EB12 - 

Marshall Space Flight Center, AL 35812 

Marvin Warshay 

Lewis Research Center 

MS 301-5 

21000 Brookpark Rd. 

Cleveland, OH 44135 

Ph: (216) 433-5261 Fax: 

Tom Whitt 

Marshall Space Flight Center 

EB12 

Marshall Space Flight Center, AL 35812 

Ph: (205) 544-3313 Fax: 

Floyd L. Williamson 

ZAE Research, Inc. 

555 Sparkman Dr. 

Suite 450 

Huntsville, AL 35816 

Ph: (205) 890-0507 Fax: (205) 890-0857 

E-mail: flw@hiwaay.net 

Doug Wright 

Eagle-Picher Industries 

PQB 47 

Joplin, MO 64802 

Ph: (417) 623-8000 Fm: (417) 623-0850 

E-mail: dwrightaepi-tech.com 

Attendance List 
I997 NASA Aerospace Bnttery PVorkshop Attendance List 

Thomas 1.i 

Goddard Space Flight Center 

Code 734 

Greenbelt. hlD 2077 1 

Ph. (301) 286-305 I Fay: (301) 286-9214 

.Albert H. Zimmernian 

The .Aerospace Corporation 

hlS h121275 

POB 92957 

Los Angeles. CA 90009-2957 

Ph: (3 10) 336-7415 Fas: (3 10) 336-1636 
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I. AGENCY USE ONLY (Leave Blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED 

July 1998 Conference Publication 

4. TITLE AND SUBTITLE ( 5. FUNDING NUMBERS 

I The 1997 NASA Aerospace Battery Workshop 

I 6. AUTHORS 

J.C. Brewer, Compiler 

7. PERFORMING ORGANIZATION NAMES(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION 

REPORT NUMBER George C. Marshall Space Flight Center 

I Marshall Space Flight Center, AL 35812 

9. SPONSORlNGlMONlTORlNG AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORlNGlMONITORING 

National Aeronautics and Space Administration AGENCY REPORT NUMBER 

I Washington, DC 20546-0001 

I 

11. SUPPLEMENTARY NOTES 

Proceedings of a workshop sponsored by the NASA Aerospace Flight Battery Systems Program, hoste 

by the Marshall Space Flight Center, and held at the Huntsville Hilton on November 18-20, 1997. 

12a. DISTRlBUTlONlAVAlLABlLlTY STATEMENT 12b. DISTRIBUTION CODE 

Unclassified-Unlimited 

Subject Category 44 

Standard Distribution 

1 13. ABSTRACT (Maximum 200 words) 

This document contains the proceedings of the 30th annual-NASA Aerospace Battery Workshop, 

hosted by the Marshall Space Flight Center on November 18-20, 1997. The workshop was 

attended by scientists and engineers from various agencies of the U.S. Government, aerospace 

contractors, and battery manufacturers, as well as international participation in like kind from a 

number of countries around the world. 

The subjects covered included nickel-cadmium, nickel-hydrogen, nickel-metal hydride, .lithium, 

lithium-ion, and silver-zinc technologies, as well as various aspects of nickel electrode design. 

14. SUBJECT TERMS 15. NUMBER OF PAGES 

battery, nickel-cadmium, nickel-hydrogen, nickel-metal hydnde, lithium, 810 

lithium-ion, silver-zinc, nickel electrode, electrode, battery tests 16. PRICE CODE 

A99 

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIFICATION 20. LIMITATION OF ABSTRACl 

OF REPORT OF THIS PAGE OF ABSTRACT 

Unclassified I Unclassified I Unclassified I Unlimited 

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) Prescribed by ANSI Std. 239-18 

298-102 
National Aeronautics and 

Space Administration 

ATOlS 

George C. Marshall Space Flight Center 

Marshall Space Flight Center, Alabama 

35812