Flight data for 17,000 orbital cycles are reviewed and summarized. The nickel cadmium battery system operated without failure or abnormality. Battery trend analysis used in determining the feasibility of extending mission life is discussed. The life test data for 20% depth of discharge indicates design life requirements would be reached even at a deeper depth of discharge

Mcdermott, J. K.

English

NASA Technical Reports Server

IN FLIGHT PERFORMANCE OF A 
SIX AMPERE-HOUR NICKEL-CADMIUM BATTERY 
IN LOW EARTH ORBIT 
Joseph K. McDermott 
Martin Marietta Denver Aerospace 
Denver, Colorado 
ABSTRACT 
A low earth orbit (LEO) spacecraft has successfully completed over 17,000 
orbital cycles. The energy storage system for this spacecraft supplies eclipse 
power and power for peak loads, and is a 28 volt, 6 ampere-hour Nickel-Cadmium 
battery system. There are two batteries per spacecraft, and each battery 
consists of 21 series connected, hermetically sealed Nickel-Cadmium cells. 
The cells utilize a pellon separator, a standard positive plate and a 
silver-treated negative plate. 
This paper reviews and summarizes flight data for 17,000 orbital cycles. 
Discussion focuses on battery trend analysis used in determining the 
feasibility of extending mission life. 
SYSTEM DESCRIPTION 
The battery function on this spacecraft is power production during 
eclipse periods and power provision for power conditions in which the required 
load exceeds the solar array power. The average bus load for the completed 
17,000 orbital cycles was 40 watts. The batteries equally share the bus load, 
although the system was designed so that either battery could support the load 
during eclipse periods.. A simplified electrical power system (EPS) block 
diagram, Figure 1, shows that each battery has its own battery charger 
comprised of a primary (A) and backup (B) charger. The primary charger 
utilizes the Voltage/Temperature (V/T) charge curves shown in Figure 2. The 
lower V/T curve (A) is approximately equivalent to level 4 of the NASA Modular 
Power System V/T curves. The backup charger operates at a constant C/10 
charge rate or a commandable C/60 trickle charge rate. Only the primary 
charger has been employed for the completed 17,000 cycles with all operations 
at level A. 
https://ntrs.nasa.gov/search.jsp?R=19840025619 2020-03-22T10:18:47+00:00Z
The spacecraft solar array system comprises three body-mounted solar 
panels and four deployable solar paddles. There are a total of 53 solar cell 
strings. Thirty-five strings are permanently wired to the power bus while 18 
strings can be ground commanded IN or OUT. The power control and distribution 
unit routes and controls power to the battery chargers and from the solar 
arrays or batteries. Excess power is controlled with the 18 switchable solar 
cell strings. When the bus voltage rises above 32 volts, a shunt control unit 
switches shunt resistor banks ON to dispense excess power. 
CELL~BATTERY DESCRIPTION 
Nickel-Cadmium cell design parameters are summarized in Figure 3. Plate 
loading and the negative-to-positive capacity ratio are standard for sealed 
spacecraft cells (ref. 1). The material for the positive and negative 
terminals is Nickel 200, the separator is Pellon 2505 and the cell case is 
304L stainless steel. The cell height, including terminals, is 3.63 inches. 
Width is 2.13 inches and thickness 0.82 inches. The cell weight averages 285 
grams. 
The six ampere-hour, 28 volt Nickel-Cadmium battery, depicted in Figure 
4, consists of 21 series connected cells. Battery dimensions are provided in 
Figure 4, and the battery weight is 16.3 pounds. Three thermistor temperature 
sensors are strategically placed on each battery. Data from the two 
thermistors mounted to the top surfaces of cells number six and eleven are 
included in telemetry data. A third thermistor is mounted between cells 
number 16 and 17, and is used for the V/T charge control. The battery is 
mounted to an aluminum honeycombed deck in the spacecraft with an RTV-60 
compound for improved heat transfer. 
FLIGHT DATA 
The battery data base for this spacecraft includes battery temperature, 
recharge fraction (chargeldischarge ratio), average daily depth-of-discharge 
(DOD), minimum battery voltage and the bus load. Battery performance is 
directly related t,o solar array performance, which is influenced by percent 
illumination, alpha angle (complement of beta) and solar flux. The solar flux 
is normalized to a standard 135.3 milli-watts per square centimeter. These 
parameters are provided in Figure 5 as a function of orbit number. 
The average battery temperature for the 17,000 cycles was approximately 
20°C. Battery temperature versus revolution (orbit) is plotted in Figure 
6. Comparison of Figures 5 and 6 illustrates battery temperature is largely a 
function of the alpha angle, although percent illumination and solar flux also 
influence battery temperature. The temperature plot, Figure 6, indicates 
spacecraft battery temperature is high compared to other aerospace LEO 
vehicles (ref. 2). High battery temperatures occurred only on this 
spacecraft, the first of a series. Several subsequent spacecraft have 
maintained battery temperatures to less than 10°C. 
The average daily DOD for both batteries has been consistently below 10 
percent. The low average DOD results from the conservative design approach; 
i.e., sizing the power system for operation with one battery failure. The DOD 
largely depends on the eclipse period, with larger DOD during an eclipse 
greater than 22 percent of the orbit period. This condition is illustrated, 
Figure 7, for DOD and recharge fraction, which are closely related. The 
recharge fraction for eclipse periods greater than 22 percent of the orbit is 
from 110% to 120%. This recharge fraction is considered normal for a 
temperature of 20'~ and for the V/T charging level used for this spacecraft. 
The minimum battery voltage is the most easily observed cell degradation 
characteristic (ref. 3 and 4). Figure 8 shows minimum battery voltage as a 
function of orbit number. As seen in the figure, minimal cell voltage 
degradation has occurred during the 17,000 orbits. Minimum battery voltage 
for seven percent DOD as a function of revolution is shown in Figure 9. Data 
for the first 7000 cycles was for an Alpha 90° condition, thus was not 
plotted due to its particularity. The end-of-discharge voltage has remained 
reasonably stable during the past 10,000 orbits, ranging from 26.1 to 26.3 
volts. This corresponds to cell voltages of 1.24 to 1.25 volts. This is 
substantially above the minimum acceptable bus voltage of 23.8 volts. Barring 
any sudden changes in performance, the batteries will almost certainly meet 
and exceed the mission requirement. 
LIFE TEST DATA 
Life testing on a flight type battery has been conducted in parallel with 
the mission. Over 16,000 LEO cycles at 20 percent DOD and 25% have been 
completed to date. The minimum battery voltage degradation for 20 percent DOD 
is shown in Figure 9. The end-of-discharge battery voltage was stabilized 
around 25.7 volts up to orbit 11,500. An appreciable battery voltage 
degradation of 0.5 volts has occurred during the last 4,000 cycles. The same 
V/T charge level as the flight battery is utilized. The discharge regime 
consists of a 2.0 amperes constant current discharge. A capacity check prior 
to the start of the life test yielded 7.80 ampere-hours. The capacity check 
was performed at a ~ / 2  discharge rate to a battery voltage of 23.8 volts. 
Similar capacity checks at cycles 308 and 1264 yielded 7-35 ampere-hours and 
5.55 ampere-hours respectively. The charge and discharge regimes for cycles 
2200 and 15465 are compared in Figures 10 and 11. The end-of-charge voltage 
curves for cycles 2200 and 15465 are basically identical, thus only one charge 
voltage curve is present in Figure 10. The charge current increase from cycles 
2200 to 15465 is considered normal and has been previously reported (ref. 5 and 
6). The end-of-discharge battery voltage degradation, depicted in Figure 11, 
from cycles 2200 to 15465 is analogous to the life test minimum battery voltage 
degradation depicted in Figure 9. This voltage degradation is inherent to 
Nickel-Cadmium and is considered acceptable. 
SUMMARY 
The Nickel-Cadmium battery system on this spacecraft has successfully 
completed 17,000 low earth orbit cycles without a failure or an abnormality. 
The in-house life test data for 20 percent depth-of-discharge, Figure 9, 
indicate design life requirements would be reached, even at a deeper depth- 
of-discharge. The relatively simple, predictable and repetitive load profiles 
for this spacecraft, in conjuntion with the conservative power system design 
approach, will result in the capability to extend mission life. 
REFERENCES 
1. Scott, W. R., and D. W. Rusta. Sealed Nickel-Cadmium Battery 
Applications Manual, NASA Reference Publication 1052, December 1979. 
2. DOD-STD-1578 (USAF). "Nickel-Cadmium Battery Usage Practices for Space 
Vehicles," 27 July 1981. 
3 .  Kent, J., "Analysis and Evaluation of Spacecraft Battery Life Test Data, 
Phase 2, Summary Report," NASA CR-107119, October 1969. 
4. "Evaluation Program for Secondary Spacecraft Cells," Annual Reports of 
Cycle Life Test, Contracts W11,252B, W12,397, S234046, and S53742A6, 
Weapons Quality Engineering Center, Naval Weapons Support Center, Crane, 
Indiana, 1965-1976. 
5. Thierfelder, H., "ITOS Battery Life Cycling Test, Final ~eport," Report 
TM-SP-PO-100, RCA, ~ights town, New Jersey, March 1968. 
6. Johnson, C.R., "HEAD Battery Cell Life Test (Mod. loo), Final Report ," 
Engineering Report 77-8725.6-108, TRW Systens Group, Redondo Beach, 
California, February 22, 1977. 


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In Flight Performance of a Six Ampere-hour Nickel-cadmium Battery in Low Earth Orbit

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