The heritage of the Ni-H2 technology that makes the space station application feasible is discussed. It also describes a design for a potential space station Ni-H2 battery system. Specific design values presented here were developed by Ford Aerospace as part of the Rocketdyne team effort on the Phase B Definition and Preliminary Design of the Space Station Power System in support of NASA Lewis Research Center

Applewhite, A. Z.

Vanommering, G.

English

NASA Technical Reports Server

\N87-III01
NICKEL-HYDROGEN BATTERIES FROM INTELSAT V TO SPACE STATION
G. van Ommering and A.Z. Applewhite
Ford Aerospace & Communications Corporation
Palo Alto, California
INTRODUCTION
The Ni-H_ battery system has been under development for about 14 years
and has _een flying on several geosynchronous orbit (GEO) spacecraft
since 1983, in configurations such as the Intelsat V battery assembly
shown in Figure I. It has been qualified as well for low earth orbit
(LEO) applications but is not as yet flying in LEO. An application
now being studied in detail is the Space Station, which may require
very large Ni-H 2 batteries to meet the 75 kW power requirement cost-
effectively.
This paper discusses the heritage of Ni-H 2 technology that makes the
Space Station application feasible. It aIso describes a design for a
potential Space Station Ni-H 2 battery system. Specific design values
presented here were developed by Ford Aerospace as part of the
Rocketdyne team effort on the Phase B Definition and Preliminary
Design of the Space Station Power System in support of NASA Lewis
Research Center.
SPACE STATION Ni-H 2 BATTERY SYSTEM OVERVIEW
The Ni-H_ battery system is a current option for the Space Station
Initial Operating Capability (IOC). The system consists of four
batteries of 105 individual pressure vessel (IPV) cells. Each cell
has a nominal 275 Ah capacity. The four batteries each consist of
five battery assemblies with 21 cells. The assemblies contain heat
pipes for heat tranport to a fluid loop interface. The total system
consists of 20 battery assemblies held in two racks, one in each Power
System utility center. Design details will be provided following some
background discussion.
CELL DESIGN HERITAGE
While Ni-H_ cells are currently flying in GEO only, there is a large
body of work and data that provides confidence in the readiness of the
system for a large-scale LEO application in the early 1990s. The key
features of the design of the 275 Ah Ni-Hp cell required for the Space
Station have already been individually demonstrated. Figure 2
illustrates some of these important efforts, some completed and some
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still in process, along with their major background contribution to
this program.
Intelsat V provides important background as the first and longest
operational flight of Ni-Hp batteries. Spacenet, G-Star, and Satcom K
(RCA programs, not shown) _emonstrate cell scale-up feasibility. The
Air Force LEO cell and MANTECH programs provide LEO design and
manufacturing data bases. The dual electrode stack design concept is
being qualified on the MILSTAR program. While all these efforts
involved 3.5-inch diameter cells, the 4.5-inch diameter required for
high-capacity cells has been developed under Air Force sponsorship.
Component-level developments and life test programs are supporting
these efforts. NASA-LeRC is funding electrode optimization studies
for LEO, and has pursued other innovations such as vessel-wall-mounted
oxygen recombination technology. The Air Force is initiating life
tests at Naval Weapons Support Center in Crane, IN.
Ford Aerospace and Yardney are co-funding development of a 220 Ah
Ni-H 2 cell which incorporates the necessary and best features of these
other efforts. The cell is a 4.5-inch diameter, dual-stack LEO
design, incorporating wall-mounted recombination sites, LEO-optimlzed
components, and several new and upgraded features. Development of all
components has been completed, and tests of the first cell will com-
mence in December 1985. The cell is shown in Figure 3 next to a
typical Intelsat V flight cell. This large-cell demonstration will
establish readiness for future development of a specific Space Station
size cell, in any capacity ranging up to 300 Ah. Current Space
Station battery design calls for a 275 Ah cell, on which the discus-
sion below is based.
CELL DESIGN
The 275 Ah Ni-Hp cell is a 4.5-in diameter, tandem-stack LEO cell,
based on a combination of proven features of already developed lower
capacity cells.
The nickel electrode design for the 275-Ah LEO cell is based on design
parameters developed by the space nickel battery industry over the
last decade for long life electrodes, including sinter porosity, pore
size distribution, and loading levels consistent with those derived in
the NASA-LeRC funded research at Hughes Research Laboratories, as well
as in U.S. Air Force development efforts. The hydrogen electrodes
are based on a proven design that currently is flying on several
spacecraft. The baseline separator system combines features of demon-
strated separator materials to provide the necessary electrolyte
reservoir and barrier characteristics.
The mechanical design of the cell is derived from the demonstrated Air
Force 4.5-in cell technology. It also incorporates scaled-up features
388
i
developed under the Air Force/Yardney MANTECH program and additional
improvements to provide more uniform stack support. The stack com-
ponents are supported on a central core which attaches to the weld
ring. Each stack is held between two support/end plates, one of which
can move with respect to the core against a Belleville washer to
maintain constant compressive force over the life of the cell.
Electrode tabs are fed through the central core.
The pressure vessel is made of Inconel 718 with a 0.035-in design
thickness. The two hydroformed and age hardened shells are joined by
electron-beam welding to the Inconel 718 weld ring. The vessel has
demonstrated a 3900 psi burst pressure. Maximum operating pressure is
expected to be 1100-1200 psi. The electrical feedthroughs incorporate
hydraulic cold-flow teflon seals.
Oxygen management is achieved by recombining oxygen generated on
overcharge on the vessel wall which is coated with porous zirconia
(wall wick), on which platinum catalyst is deposited based on a design
pioneered by NASA-LeRC. Heat generated during overcharge thus is
removed very effectively without thermal burden on the stack. The
water formed is returned to the stack by the wall wick via separator
edges in contact with it. The wall wick also serves as electrolyte
concentration and inventory equilibrator, and as a reservoir.
BATTERY SYSTEM DESIGN DISCUSSION
Electrical Design
The 75-kW power requirement of the station, plus allocations for user
converter inefficiency and PMAD processors, is provided by the battery
through a 0.90 efficiency chain for a total battery system output of
95.8 kW (see Table I). Power peaks are supported during sunlight
periods by reducing the charge current if required and during eclipse
by the battery at 125.8 kW. With 105 cells in series per battery and
an average EOL discharge voltage per cell of 1.25 V, delivered
capacity per cell is about 110 Ah for a typical eclipse. Table 2
provides additional electrical design data.
Nominal battery DOD is 40%, which assures the capability for contin-
gency support following a peak eclipse, and is consistent with a
5-year life expectancy for the battery system. Cell capacity required
to meet this requirement is 275 Ah, well within the estimated minimum
300-Ah capability of the tandem stack 4.5-in cell design. The 275 Ah
capacity is achieved by adding six electrode modules to each of the
two stacks of the 220-Ah cell discussed above.
Any non-wearout failures observed in Ni-H 2 cells have typically been
short circuits. Because of this and the maintainability of the sta-
tion hardware, no individual cell bypass hardware is included in the
design. Outage of a single battery during maintenance or recondition-
389
ing represents a temporary increase in DODto 53%for the remaining
batteries, which represents no life risk.
Charge managementinvolves microprocessor-based coulometry during
charge and discharge. Charge current and time are determined based on
a programmablerecharge ratio and end-of-charge current taper profile.
Mechanical Design
The overall battery rack concept for one Utility Center is shown in
Figure 4. It has space for 10 battery assemblies which slide into the
rack. On one side cold plates are provided which interface with the
heat exchanger of the corresponding battery assembly. Cable harnesses
are incorporated in the rack with connectors at each battery shelf.
Each battery assembly, as shown in Figure 5, contains 21 cells and is
an independent unit interchangeable with any other. It consists of a
graphite/epoxy honeycomb panel on which graphite/epoxy support beams
are bonded that carry heat pipes on their top surface. Cells are
contained in aluminum sleeves, which provide mechanical support and
transport heat away from the cell as well. A resilient insulator
layer electrically isolates the cell and sleeve and provides good
thermal contact. Flanges on the sleeve are mounted to the heat pipe
saddles forming both mechanical and thermal interfaces. The cell
mounting design is shown in cross-section in Figure 6.
Battery physical data are shown in Table 3. Based on an individual
battery assembly mass of 220 kg, and a rack mass of 75 kg, the total
system mass is 4550 kg. This is not necessarily the lowest-mass
battery design, but represents the overall most cost-effective
approach.
Thermal Design
Battery thermal design relies on the cell sleeve, primary and secon-
dary heat pipes, and the utility center coolant loop as major elements
in the heat rejection path. The cell sleeve surrounds the cylindrical
portion of the cell over the length of the cell stack and is insulated
from it by a conductive layer. The sleeves conduct heat to two sets
of flanges which contact the primary heat pipes as shown in Figures 5
and 6. The primary heat pipes carry heat to one side of the battery
panel where their condensers interface with the evaporators of secon-
dary heat pipes. The latter terminate on part of the long side of the
panel where they form a heat exchanger which contacts a coolant loop
cold plate. Instead of secondary heat pipes a simple coolant conduit
fitted with quick-disconnect couplings can be used, into which an
external coolant loop is plugged.
Average heat dissipations of each battery assembly during discharge
and charge are 995 W and 270 W, respectively. Nominal temperatures
o
are 10 to 20 C. The area requirement for an AI/NH_ h_at pipe radiator
system to support the battery heat load would be 9_ m_, taking
390
advantage of heat load averaging by the battery heat capacity of 60
Wh/°C per assembly.
Llfe and Reliability
Prediction of cycle llfe capability of Ni-H^ batteries for LEO
applications is more difficult than for NI-6d batteries. The latter
have been tested extensively at Naval Weapons Support Center, Crane,
IN (NWSC-Crane). A thorough analysis of thls data base Is represented
by the model derived by McDermott and reported in various proceedings
of the NASA Goddard Battery Workshop over the last four years.
Reported cycle capabilities for early developmental 50 Ah NI-H_ cells
at 80% DOD equates to 33,000 to 40,000 cycles at 40% DOD, base_ on
applying the McDermott Ni-Cd model, which should be conservative for
the NI-H_ system. Sufficient progress has been made in NI-H_ cells
since th_ mid-1970s, that a mean cycle llfe of 40,000 to 45,D00 at 40%
DOD in LEO appears to be a realistic projection for Ni-Hp cells,
particularly in vlew of the fact that even some NI-Cd ceils have
achieved this (packs IH, IJ, 8G at NWSC-Crane).
The large In-orblt data base shows that random failure probability for
spacecraft batteries is extremely low on the many Ni-Cd and Ni-Hp
batteries that have been flown. Based on Ford Aerospace's in-orSit
experience, Ni-Cd cells exceed 40,000,000 cell hours without random
failures; industry-wide on Ni-H 2 the total is 7,000,000. Analysis
yields an expected 0.6 random cell failures for the Space Station
battery system over 5 years or an mean-time-to-failure of 68,000
hours.
Reliability analysis based on the Weibull distribution with a shape
factor of 12 and a mean cycle life of 45,000 cycles was performed.
The shape factor value is somewhat optimistic, but based on the
smaller number of wearout mechanisms in Ni-H 2 cells compared to Ni-Cd
(typical shape factor 8), and the tight control of operating
conditions, not unrealistic. Table 4 shows the estimated system
reliabilty based on these figures, an assumption that 3 shorted cell
failures per battery string are allowed, and for shorted/open failure
distributions of 98/2 and 80/20. Another variable is the possible
presence of a spare assembly for every two batteries or 10 assemblies,
which can be switched in on demand. Conclusions are:
o There is high probability (97%) that all four batteries In the
system wlll be available for use at all times during a 5 year
operational period
o Probability of uninterrupted power support for 5 years is 99.9%
391
o A single spare battery assembly per set of two batteries
provides sufficient open cell protection over a wide range of
open cell failure possibilities
The reliability analysis results indicate that it may be not be neces-
sary to plan for replacement of individual assemblies, but to replace
instead the entire battery as a package upon wearout.
SUMMARY AND CONCLUSIONS
The heritage of space Ni-H 2 batteries from Intelsat V, through many
LEO-oriented cell and compGnent development efforts and culminating in
the Ford Aerospace/Yardney development of a 220-Ah LEO cell, has
prepared the technology to a point of readiness where application on
the Space Station can be seriously considered. Practical battery
system designs have been derived that are compatible with the require-
ments of the Station. While these designs do not necessarily have the
lowest possible mass, they are configured to provide a 5-year battery
system capability with maximal cost-effectiveness.
392
’ OktGlNAL PAGE IS 
OF POOR QUALITY 
Figure 1 .  INTELSAT V NICKEL-HYDROGEN FLIGHT BATTERY 
3.5-1N. 25-50 AH 
3.5-lN. 50 AH 
ELECTRON DESIGN 
Figure 2. INTEGRATION OF DEMONSTRATED NICKEL-HYDROGEN TECHNOLOGIES FOR THE SPACE 
STATION 
393 
ORtGlNAL PRGE E3 
OF POOR QUALTPI 
Figure 3. 220-Ah LEO NiH, CELL WITH 30 Ah GEO FLIGHT CELL 
TABLE 1 SPACE STATION N I - H ~  BATTERY SYSTEM OPERATING REQUIREMENTS 
o NOMINAL DISCHARGE POWER 
o DISCHARGE DURATION 
o PEAK POWER 
o PEAK DURATION 
o RECHARGE DURATION - MAX 
o DISCHARGE VOLTAGE 
o CHARGE VOLTAGE 
o CONTINGENCY CAPABIL ITY  
= 86,25  K W  a 0.90 E F F  = 95,8 K W  
= 35.8 M I N  
= 113,25 K W  a 0,90 EFF = 1 2 5 . 8  K W  
= 7 .5  M I N  
= 58 M I N  
= COMPATIBLE WITH 160 V SOURCE BUS 
= COMPATIBLE WITH 160 V SOURCE BUS 
= 50% OF LOAD FOR 1 ORBIT  AFTER ECLIPSE 
394 
TABLE2. SPACESTATIONNI-H2 BATTERYSYSTEMDESIGNAND OPERATINGCHARACTERISTICS
o NUMBEROF BATTERIES 4
o NUMBEROF IDENTICALASSEMBLIESPERBAI-IERY 5
o CAPACITYPER BATTERY 275 AH
o CELLSPER BATTERY 105
ECLIPSE
o AVERAGEDISCHARGEVOLTAGE 131.3V
o AVERAGEDISCHARGECURRENT 182.5A
o AVERAGEDEPTHOF DISCHARGE 39%
o PEAKORBITDOD 41.6%
o NON-PEAKORBITDOD 36.4%
o WORST-CASECONTINGENCYDOD 97,9%
o AVERAGEHEATDISSIPATION 19.9KW
CHARGE
o AVERAGECHARGEVOLTAGE 154.4V
o MAXIMUMCHARGECURRENT 128.4A
o AVERAGEHEATDISSIPATION 5.4 KW
BATTERY S
RACK
Figure 4. NiH2 BATTERY RACK CONFIGURATION IN UTILITY CENTER BAY
395
BRACES (TYPICAL ALL CELLS)
HONEYCOMB PANEL
Figure 5. SPACE STATION NiH2 BATTERY ASSEMBLY LAYOUT
6.50 in.
I
I
8.50 in. m,
,ll-l.OO in..-e=, F 4.65 in.
/
7" / _<°"+'_-°+"n
i _ // /i +_o.+,o.
MIIIIINIIIIIIIIIIIMIIIIIIIIIIIIIIIIIIIIIIrlIIIIIIIIIIIIIlllllllllllllillllllllllllllllllll lllllllt[,=,n
Figure 6. SPACE STATION NiH2 CELL MOUNTING CONFIGURATION
396
TABLE3, SPACESTATIONNI-H2 BATTERYSYSTEMPHYSICA_CHARACTERISTICS
MASSPER CELLKG(LB)
MASSPER BAI-[ERYASSEMBLYKG(LB)
MASSPER BATTERYKG (LB)
TOTALESS MASSKG(LB)
CELLDIMENSIONSCM (IN)
BAI-I'ERYASSEMBLYDIMENSIONSM (FT)
BAI-I'ERYDIMENSIONSM (FT)
BATTERYSYSTEMDIMENSIONS(2 EA,)M (FT)
TOTALBATTERYSYSTEMENVELOPEVOLUMEM3 (FT 3)
6,99 (15.q2)
220 (485)
II00(2425)
4550 (10030)
55,1x 11,8DIA
(21.7x q.65DIA)
2.52 X 1.25X 0.17
(8.30X 4,10X 0.56)
2,52 X 1.25X 0.95
(8,30X 4.10X 3.12)
2.62 X 1.35X 1.90
(8.60X 4.43X 6,23)
13.4 (475)
THERMALMASS WH/°C 1210
TABLE 4. SPACE STATION NI-H2 BATTERY RELIABILITY ESTIMATES
981; Short/Z1; Open 80% Short/201_ Open
Wtthout
Spare Asseulbly
Ntsston ' '
T|m 4
(years) Batteries
1 >0.9999
Z >0.9999
3 >0.9999
4 0.9979
5 O. 9674
6 0.0839
3of4
Batteries
> 0.9999
>0.9999
> 0.9999
>0.9999
O. 9996
0.3721
With
Spare Ass_l tes
4 30f4
Batte ri es Batte ties
>0.9999 > O.9999
>0.9999 >0.9(399
>0.9999 >0.9999
>0,9999 >0,9999
0.9974 >0.9999
0.1086 0.4310
WIthout
Spare Assembly
4 3of4
Batteries Batteries
Wtth
Spare Asseulbltes
4 3of4
Batteries Batteries
> O. 9999 > 0.99(399
>0.9(399 >0.999(39
0.9993 >0.9999
0.9790 0.9998
0,7340 O. 9699
0.0165 0.1347
>0.9999 >0.9999
>0.9999 >0.9999
>0.9999 >0.9999
0.9999 >0.9999
0.9777 0.9998
0,0934 0.3955
397


Nickel-hydrogen batteries from Intelsat 5 to space station

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