India's first experimental communication satellite, APPLE, carried a 12 AH Ni-Cd battery for supplying power during eclipse. Failure to deploy one of the two solar panels resulted in the battery operating in a high temperature environment, around 40 C. This also resulted in the battery being used in diurnal cycles rather than just half yearly eclipse seasons. The management and performance of the battery during its life of two years are described. An attempt to identify the probable degradation mechanisms is also made

Agrawal, B. L.

Subrahmanyam, A.

Suresh, M. S.

English

NASA Technical Reports Server

MANAGEMENT AND PERFORMANCE CIP APPLE BATTERY 
IN HIGH TEMPERATURE EWICBLONMENT 
M. S. Smesh 
A, Sukamanyam 
B. L. A e a w d  
ISRO Satellite Centre 
Bangalore, India 
ABSTRACT 
India's first experimental communication satellite, APPLE, carried a 1 2  AH Ni-Cd 
battery for supplying power during eclipse. Failure to deploy one of the two solar 
panels resulted in the battery operating in a high temperature environment-- 
around 40 C. This also resulted in the battery being used in diurnal cycles rather 
than just half yearly eclipse seasons. This paper describes the management and 
performance of the battery during its life of two years. An attempt to identify 
the probable degradation mechanisms is also made. 
INTRODUCTION 
APPLE (Ariane Passenger Payload Experiment) spacecraft was launched by Ariane 
LO3 on June 19, 1981. This was a 380 Kg (in orbit mass) three-axis stabilized 
satellite parked at 102 E longitude and designed for a two-year life to conduct 
communication experiments. The power system was comprised of two deployable, 
sun-tracking solar panels supplying 250 watts of power to a regulated 28 volt bus. 
A 1 2  AH, 16 cell Ni-Cd battery was employed to supply eclipse, transfer orbit, and 
peak loads. Figure 1 shows the block schematic of the power system. 
The battery charge path had two chargers-one for the transfer orbit phase and 
the other for the on orbit phase. The on-orbit charger was a fixed 2 ampere char- 
ger with the 8-level voltage-limited auto and manual control for the changeover 
to a 350 mA trickle charge. The transfer orbit charger was a variable current 
charger which maintained the bus voltage at 33 volts or 28 volts as selected by 
ground command, and thereby pumping all the extra current into the battery. 
Control of the transfer orbit charger was also possible by the 8-level voltage limit 
or by ground command. However, the transfer orbit charger opened the charge 
path once the battery was completely charged. 
BATTERY DESIGN 
The battery was a single, 16 cell, 1 2  AH battery supplied by SAFT, France. The 
cell case was, in fact, identical to that of the 10 AH cell, and cells with a slightly 
higher loading were selected to achieve the capacity of 1 2  AH. The design para- 
meters and constructional details of the battery are given in Tables 1 and 2. The 
cell details are provided in Table 3. The battery was designed for an operational 
life of two years at 72% maximum depth of discharge, thereby undergoing four 
https://ntrs.nasa.gov/search.jsp?R=19840025622 2020-03-22T10:18:35+00:00Z
eclipse seasons. Complete reconditioning was not envisaged in view of its short 
life, and limited reconditioning to about 17.6 volts w a s  possible by proper choice 
of charger and control combinations. 
Table 1 
Battery Design Parameters 
Capacity 
Number of cells in series 
Operating voltage range 
Temperature range of 
operation 
Peak depth of discharge 
Operating life 
Charge current onorbit 
Transfer or bit 
Trickle charge current 
onor bi t 
Transfer orbit 
Discharge current 
Expected EOL capacity 
Expected EO L discharge 
Voltage 
Thermal control 
Battery Details 
Battery type 
Capacity 
Dim ensions 
Mass 
Thermal cutout 
Temperature sensor 
Cell type 
Vendor 
Cell bypass diodes 
Heaters 
Mounting 
12 AH 
16 
19-24 volts 
0 C-25 C (Average 20 C9 
7 2 8  
2 years - 180 cycles 
(4 eclipse seasons) 
1.8 0.2 Amps 
0 to  5 Amps 
120 10 mA 
0 (No trickle-battery open) 
6 Amps 
9.5 AH 
18.6 volts a t  the end of 
70% discharge 
Passive 
Nickel Cadmium 
12 AH (Nominal) 
225 x 135 x 225 mm 
9.3 Kg 
35 2 C  
4K thermistors (two) 
VO 10 S2 
SAFT, France 
Provided 
5 Watts (two) 
Under the  Al3M strut on the  
bottom deck 
Table 3 
Cell Details 
Cell type 
Capacity 
Mass 
Dimensions 
,Positive Electrode Number 
Thickness 
Loading 
Negative Electrode Number 
Thickness 
Loading 
Negative t o  Positive Ratio 
Case Thickness 
Negative Treatment  
Inter-electrode Distance 
Separator 
Electrolyte 
Specific Gravity 
VO 12 S2 
12 AH 
490 gms 
85 x 29.2 x 76.2 mm 
12 
- 
1 3  
0.89 mm 
17.5 gm/dm2 
1.8 
0.4 mm 
None 
0.26 mm 
Villidon 2119 
KOH 
1.306 
Figure 2 shows the l i fe  predictions by SAFT for their cells and i t  is  seen tha t  the  
design was adequate t o  meet  t he  two-year l ife goal at 20 C with 0.99 reliability. 
The thermal control of the  battery was passive control backed up by two heaters. 
ORBlT OPERATION 
The spacecraft  was put into synchronous orbit on the sixth transfer orbit, but the  
north solar panel could not be deployed resulting in covering a large area provided 
for thermal control. The battery temperature thus attained very high values. 
Another problem also arose as a consequence. As can be seen from the  power sys- 
tem schematic, certain heater loads were directly connected onto the battery 
apart  from the  emergency and power control loads. I t  was envisaged in the design 
tha t  the charger current in the  trickle charge mode be set t o  500 mA which 370 
mA was assigned t o  t he  loads and 130 mA for charging the  battery. However, due 
t o  the thermal problems, these heaters were not switched ON reqularly and about 
350 mA was flowing into the  battery, In trickle charge mode, this would be 
wasted as heat resulting in the  battery attaining a higher temperature. It  was 
thus decided t o  change over t o  t h e  transfer orbit charger in trickle mode, subse- 
quent t o  the  charging by the  on-orbit charger. Since the  transfer orbit charger in 
trickle mode opens the  charge path, a small amount of discharge ( 120 mA) was 
experienced by the  battery. During the  non-eclipse seasons, thus, the  battery was 
discharged daily t o  about 19.8 volts at 120 mA and was charged subsequently at a 
t ime when the  temperature was low, ie., around 5 UT. 
The battery was thus subjected t o  diurnal cycles apar t  from the  eclipse cycles. 
The operating temperatures were too  high from any standards for a battery. 
Reduced generated power due t o  t he  nondeployment of one solar panel did not 
cause much problem as the margin available was large but the  thermal problems 
were acute* The eclipse loads were reduced to a minimum to enhance the life of 
the battery. The maximum discharge was about 3 AH during the peak eclipse at 
2.5 A sate and during the non-eclipse season it was about 2 AH at 0.12 A rate. 
In Figure 3, battery voltage, current, and temperature are plotted For one 
complete day, Septemkr 20, 1981, a day on which the eclipse dwation was 64 
mim. It is obvious that the battery spends most of the time with low discharge in 
the transfer orbit trickle charge mode. Prior to and soon after eelipse, the bat- 
tery was charged* 
Table 4 lists the major events in the life of the battery in chronologicd order. 
The battery was not reconditioned prior to the first and second eclipses. Before 
the third edipse, limited conditioning was done. Perhaps, it is more appropriate 
to call them just deep discharges as the battery was discharged only up to 19.5 
volts at 300 mA. The problem about remnditioning was that manual control of 
charge/discharge was possible only with the redundant charger system and when 
all the main systems were working satisfactorily, changeover to the redundant 
system was not advisable. Only prior to the fourth eclipse season a meaningful 
reconditioning was possible and even then the discharge was not complete as the 
battery was required to supply peak power to the momentum wheel assembly. 
Table? 4 
Major Evenb In T k  Lge  Of The Battey 
Launch 
Orbit acquisition 
First edipse season 
Second edipse season 
Limited reconditioning 
Third eclipse 
Reconditioning after change- 
over to redundant systems 
Fourth eclipse season 
Reconditioning cycle 
Fifth eclipse season 
End of mission 
June 19,19841 
June 22, 1981 
August, September 1981 
Peb~uary, March I982 
August 11, 1982 
August, September 1982 
February 17,1983 
February, March 1983 
August 16, 1983 
August 24,1983 
Spetember 12, 1883 
In the reconditioning performed on Febrursiry 17, 1983, discharge was up to 18,1 
volts in the first cycle and up to 18.5 volts in the seeond cycle. Discharge beyond 
18.5 volts was not attempted as the eleventh cell voltage reached 0.7 volts, 
whereas the other cell voltages were still above 1.15 volts. The charge/discharge 
eharacteristim of these two cycles are shown in Figure 4. The improvement in 
the charaeteris8ie is evident. 
Before the fifth eclipse season, three reconditioning cycles were completed but 
this time reconditioning was done, rather inadvertantly, down to 17.1 volts. The 
fifteenth cell reversed to -0.16 volts for a few minutes. Figure 5 shows the 
ehargeldischarge cycles. The improvement in capacity from 5 AH to 6.5 AH and 
improvement in the voltage performance are clearly evident. 
On September 12, 1983, during the fifth eclipse season, soon after an eclipse, the 
solar panel failed to track the sun resulting in a total discharge of the battery 
down to zero volts. This was the first opportunity when the battery capacity was 
measured at normal discharge rate of 2.5 amps. During this discharge, the battery 
delivered 6.25 AH down to zero volts. The 12 AH battery had degraded by 50%: 
Considering the extremes of strain to which the battery was subjected, this was a 
remarkably good performance. 
BA'PTERY PERFORMANCE 
Figures 6 and 7 show the variation of the battery temperature. Figure 6 shows the 
diurnal variation for days which are approximately six months apart and Figure 7 
shows the variation of the maximum and minimum temperatures over its life. The 
days selected in Figure 6 correspond to the minimum temperature encountered by 
the battery during the equinoxes and the maximum temperature during the winter 
solstice. It is observed that for most of the time, the battery temepratures were 
above 30 C and for a considerable time, above 40 C, reaching a maximum of 55 C: 
on certain days (not shown in the figure). Increase in the battery temperature 
over its life was also due to the degradation of the passive thermal control ele- 
ments - OSR. Consequent to this, a crucial problem regarding battery manage- 
ment was the setting of the voltage at which the battery should change over from 
charge to trickle charge. Improper setting would result in the battery getting 
either grossly undercharged or overcharged. No data is available regarding bat- 
tery performance at temperatures above 30 C making predictions almost impos- 
sible. 
Table 5 summarizes the details of the performance of the battery during the four 
eclipse seasons. The battery temperature during charge was maintained within a 
degree. The post-eclipse end-of-charge voltage varied from 22.83 volts to 23.13 
volts. These values are the forced values rather than the values that would have 
been attained had the battery been allowed to charge completely under normal 
circumstances. In fact, it is evident that during the third and fourth eclipse sea- 
sons the battery was not charged completely. However, the battery was charged 
sufficiently before every eclipse to supply the required 3 AH of discharge. 
A comparison of the change in the charge voltage can be made by examining the 
battery charge characteristics during reconditioning before the fourth and fifth 
eclipse seasons-refer Figures 4 and 5. During reconditioning the charging was 
done in the manual mode and the battery was charged until it definitely showed 
signs of overcharge. During the first reconditioning, the battery charge voltage 
was 23.25 volts and during the second reconditioning, 24.75 volts. These two 
values correspond to temperatures of about 31 C, End-of-charge voltages on 
average cell basis correspond to 1.55 volts during the second reconditioning. This 
was rather unexpected and clearly shows a significant increase in the over poten- 
tials-increase in the internal resistance. The charge characteristic also is very 
much different from what one would normally expect at this temperature-not 
flat and sharply rising while nearing end-of-charge. The over charge factors cal- 
culated from these cycles are around 1.2. 
If one looks at the discharge characteristics--refer to Figure 8--one cannot figure 
out much, as there is hmdly any change in them from the first eclipse season to 
the fourth eclipse cycle. Our first opportunity to measure the capacity was during 
the first reconditioning cycle before the fourth eclipse season. For a discharge 
rate of 0.3 amps9 the capacity down to 18.4 volts (1.15 voltsjcell) was 6 AH. In 
the second reconditioning cycle, the battery delivered 6.25 AH at the same rate. 
The actual details of the discharge cycle are shown in Figure 9. 
Figure 10 compares the discharge characteristics with those of the fresh battery, 
as determined by ground tests prior to launch. Superposed over its characteristics 
are the discharge characteristics during the first eclipse and the last discharge 
cycle. It is observed that up to 2.5 AH discharge, there is absolutely no voltage 
degradation. 
CONCLUSIONS 
In conclusion, one can observe the following degradation modes: 
(a) Very high end-of-charge voltage-1.55 volts-an abnormal value for 
this high operating temperature. 
(b) Reduction in the battery capacity by about 50%. 
(c) No appreciable discharge voltage degradation up to 2.5 AH, but a sud- 
den fall  subsequently. 
Abnormal end-of-charge voltages indicate very high internal impedance due to the 
hydrolysis of the separator resulting from operation at high temperatures. This 
could also partly be due to electrolyte redistribution. The charge characteristics 
also indicate a significant loss of negative electrode capacity, approaching nega- 
tive capacity limitation. The loss in capacity is due to the combined effects of 
separator hydrolysis and electrode degradation. 
The surprising fact is the non-appearance of voltage degradation. Considering the 
fact that the battery has undergone 730 diurnal cycles in addition to the eclipse 
cycles and without proper reconditioning, this is remarkable. 
The authors are grateful to Mr. S. Kini for the assistance provided during all the 
tests conducted on the battery. 
The authors are also grateful to Mr. R. Ashiya, Head Electronics Division and 
Director, ISAC, for their constant encouragement and permission to publish this 
paper. 


Figure  2 .  L i f e  expectancy of Ni-Cd ce l l s  24h o r b i t .  










Management and Performance of APPLE Battery in High Temperature Environment

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