The design and testing of charge control concepts for several power systems is presented. The design of the CRAM charge control system for the SAS satellites and the design of the battery charge regulator for the AMPTE satellites is reviewed. Another charge control concept examined is for the OPEN program and its use of a microprocessor. The use of the voltage limiter is also discussed

Sullivan, R. M.

English

NASA Technical Reports Server

NICKEL CADMIUM CHARGE CONTROL CONCEPTS 
Ralph M. Sullivan 
Johns Hopkins University Applied Physics Laboratory 
ABSTRACT 
APL has used several different  types of charge control systems on the i r  
spacecraft. Some have used dissipative shunts to  get rid of the excess solar 
array power; some have used non-dissipative shunts and some have used a hybrid 
system. Although they have a l l  worked reasonably we1 1, there are tradeoffs to  
be made between the impact on the thermal design caused by the dissipative 
devices and the generation of conducted emissions caused by the switching of 
non-dissipative devices. 
INTRODUCTION 
The primary function of a spacecraft battery charge control system i s  to  
provide sensible l imits of battery discharge, recharge and overcharge to  
prevent undue s t resses  within the battery. In most cases t h i s  means that the 
charging current from the solar array must be reduced af ter  the battery has 
reached ' f u l l  charge'. This leads to  some interesting thermal problems, not 
just  for the battery, b u t  for  the ent i re  spacecraft. 
SAS A I B  
Figure 1 shows one of the early attempts to  solve t h i s  problem at APL on 
the Small Astronomy Sa te l l i t e s  (SAS) A and B e  These s a t e l l i t e s  were small, 
using a single battery of eight 6 ampere hour ce l l s  t o  sustain a 30 watt load 
through a 36 minute eclipse which recurred every 96 minutes ( a  300 nmi 
equatorial o rb i t ) .  The Low Voltage Sensing Switch was se t  at  1.1 volts per 
battery cell  (8.8 vol t s )  t o  preclude excessive discharge which could resul t  in 
the reversal of a low capacity ce l l .  
For such a small battery the instantaneous variation in the array power 
was large (40 to  100 watts). This led t o  the need for  correspondingly large 
shunts driven by a sophisticated Charge Regulator and Monitor (CRAM) system. 
CRAM monitored the bus voltage, the battery current and the battery tempera- 
ture. I t  had a voltage l imiter that  limited the bus voltage in accordance 
w i t h  a voltage - temperature curve whether the battery was on the l ine or not, 
I t  also had an electronic coulometer that  both monitored and controlled the 
https://ntrs.nasa.gov/search.jsp?R=19840025617 2020-03-22T10:18:58+00:00Z
b a t t e r y  recharge i n  p a r a l l e l  w i t h  the  vol tage l i m i t e r .  There were four  
commandable percent re tu rns :  105%, 11 0%, 125%, and 'moni t o r  o n l y '  o r  i n f i n i t e  
re turn .  
F igure  2 shows a t y p i c a l  charge, discharge p r o f i l e .  The e l e c t r o n i c  
coulometer counted down t o  a value i n d i c a t i v e  o f  t h e  ampere - minutes d i s -  
charged. During recharge i t counted back a t  a s l i g h t l y  d i f f e r e n t  r a t e  depend- 
i n g  on t h e  percent r e t u r n  t h a t  was selected. A f t e r  about 90 t o  95% o f  t h e  
charge was returned, the  vol tage l i m i t e r  would s t a r t  t o  reduce t h e  b a t t e r y  
charge r a t e  i n  order  t o  l i m i t  t he  bus vol tage t o  t h e  value requ i red  by t h e  V-T 
curve. When t h e  coulometer was s a t i s f i e d  (1 10% r e t u r n  f o r  t he  case o f  f i g u r e  
2), then t h e  b a t t e r y  t r i c k l e  charge r a t e  was reduced t o  a preset  t r i c k l e  
charge r a t e  o f  C/20 (300 ma). This r e s u l t e d  i n  a corresponding drop i n  the  
bus voltage. 
Aside from these gradual changes i n  the  bus voltage, the  bus was r e l a -  
t i v e l y  qu ie t ,  w i t h  l i t t l e  o r  no noise o r  conducted emissions being generated 
by the  charge c o n t r o l  system. This was t r u e  even wi thout  the  b a t t e r y  on t h e  
1 i n e  (So lar  Only Mode) and was p r i m a r i l y  due t o  the  use o f  1 inear  shunts. 
Return ing t o  f i g u r e  1, we see two groups o f  l i n e a r  shunts; one i n t e r n a l  
and f o u r  ex te rna l .  The reason f o r  an i n t e r n a l  l i n e a r  shunt was t o  provide 
i n t e r n a l  heat i f  requested by a p roper l y  located thermis tor .  This was done 
over the  p r o t e s t s  o f  t he  thermal designer s ince the  heat was provided dur ing  a 
shor t  pe r iod  a t  t h e  end o f  sun l i gh t  a f t e r  t h e  b a t t e r y  was charged. For t h i s  
reason and because t h e  shunt d r i v e r  can be r e l a t i v e l y  heavy, t h i s  concept i s  
not  recommended and has not been used on subsequent APL spacecraft .  
The ex te rna l  shunt was d i v ided  i n t o  f o u r  p a r a l l e l  un i t s ,  one mounted on 
each o f  t h e  f o u r  so la r  panels and was designed w i t h  a t o t a l  capac i ty  o f  120 
wat ts  (30 wat ts  each) so t h a t  t he  system could func t ion  even i f  one shunt were 
l o s t .  Re fe r r i ng  t o  f i g u r e  3, we can see t h a t  t he  shunt r e s i s t o r ( s )  must be 
designed t o  accept a l l  t h e  ava i l ab le  power o f  30 wat ts  each. This was accom- 
p l  ished by p lac ing  a d i s t r i b u t e d  r e s i s t o r  under the  so la r  c e l l s  on each o f  t he  
f o u r  so la r  panels cover ing an area of 1.4 f t2 .  The shunt d r i v e  t r a n s i s -  
tors,  however, could not  be d i s t r i b u t e d .  We see f rom f i g u r e  3 t h a t  they  
experienced a peak power d i s s i p a t i o n  of Pmax/4 when shunting about one-half 
t h e i r  maximum cu r ren t .  This i s  about 7-1/2 watts, an acceptable power l eve l ,  
bu t  not  i d e a l  f o r  long l i f e .  Indeed on SAS-A, we d i d  experience f a i l u r e s  o f  
the  shunt d r i v e  t r a n s i s t o r s  w i t h i n  a year. This was found t o  be due t o  a 
defect  i n  t h e  t r a n s i s t o r  which was corrected on SAS-B by  rep lac ing  the  
t r a n s i s t o r  w i t h  one o f  a d i f f e r e n t  manufacture. 
SAS -C 
Due t o  a near doubl ing of t he  power on SAS-C, t h e  charge c o n t r o l  system 
was expanded t o  inc lude d i g i t a l  shunts (see f i g u r e  4). The so la r  a r ray  was 
292 
d iv ided i n t o  24 equal, p a r a l l e l  segments; each w i t h  a  sho r t i ng  t r a n s i s t o r  
con t ro l l ed  by CRAM. The l i n e a r  shunt could then be s ized t o  have s l i g h t l y  
more capac i ty  than one o f  t h e  d i g i t a l  shunts. Although the  shunt d r i v e  t r a n -  
s i s t o r  s t i l l  had t o  be designed f o r  9 watts power d i ss ipa t i on ,  i t was placed 
i n  a  benign environment i n s i d e  t h e  spacecraft  s ince the re  was o n l y  one o f  
them. 
Any abrupt change i n  t h e  power balance was accommodated by the  l i n e a r  
shunt. I f  it became saturated (o r  empty), then the  d i g i t a l  shunts were 
shorted (o r  opened) sequent i a1 l y  unt i 1 t h e  power imbal ance could be accepted 
by the  l i n e a r  shunt. SAS-C was a lso designed t o  operate wi thout  a  b a t t e r y  
(Solar Only). Since t h e  b a t t e r y  was one o f  t he  few non-redundant items on the  
spacecraft, i t was thought t h a t  we should a t  l e a s t  be able t o  operate dur ing  
the  s u n l i t  p o r t i o n  o f  t he  o r b i t  i n  the  event t h a t  t he  b a t t e r y  f a i l e d .  An 
Ac t i ve  R ipp le  F i l t e r  (not  shown) was employed t o  reduce bus t rans ien ts  when 
swi tching d i g i t a l  shunts wi thout  a  b a t t e r y  on t h e  i i n e .  
The charge c o n t r o l  concept o f  a  recent  APL s a t e l l i t e ,  AMPTEICCE (The 
Act ive  Magnetospheric pa r t i c l i e  Tracer ExplorersICharge Composition Explorer),  
i s  shown i n  f i g u r e  5. This i s  the  f i r s t  charge con t ro l  system a t  APL t o  use a  
microprocessor, an RCA 1802. The f i g u r e  shows one s ide  o f  a  two b a t t e r y  
system f o r  simp1 i c i t y .  The actual  a r ray  power, shunt power and load are 
double those shown. The coulometer has no con t ro l  f u n c t i o n  i n  t h i s  system, 
but  i s  s imply a  monitor.  Also, there  i s  no l i n e a r  shunt, bu t  o n l y  a  shor t ing  
t r a n s i s t o r  f o r  each so la r  a r ray  segment. The t r a n s i s t o r s  used i n  t h i s  case 
are MOSFETS IRF130 w i t h  a  very low sa tu ra t i on  res is tance,  which f u r t h e r  reduce 
the  power d i ss ipa t i on .  
This system i s  s impler  and o f  lower power d i s s i p a t i o n  than i t s  predeces- 
sors, bu t  i t  does r e s u l t  i n  some add i t i ona l  bus noise as shown i n  f i g u r e  6. 
Also, t h i s  system does not have any 'Solar  Only '  c a p a b i l i t y .  
OPEN PROGRAM 
Figure 7  shows s t i l l  another charge con t ro l  concept. One t h a t  i s  being 
proposed f o r  t h ree  s a t e l l i t e s  fo r  t h e  OPEN (Or ig in  of Plasmas i n  t h e  Ear th 's  
Neighborhood) program and which w i  11 make extensive use o f  a  microprocessor. 
At f i r s t  glance, the  much higher power l e v e l s  o f  470 t o  800 wat ts  from the  
so la r  a r ray  would seem t o  d i c t a t e  another d i g i t a l  (non-d iss ipa t ive)  system. 
However, t he re  i s  a  s t rong des i re  t o  have very low conducted emissions on 
these s a t e l l i t e s ,  a  requirement t h a t  has forced us t o  re - th ink  the  l i n e a r  
shunt. 
Sixteen l i n e a r  shunts w i t h  a  capac i ty  o f  25 wat ts  each w i l l  accept the  
e n t i r e  800 wat ts  f rom t h e  array.  The d i s t r i b u t e d  r e s i s t o r s  can once again be 
placed on the  s o l a r  panels under the  so la r  c e l l s .  To reduce the  power i n  the  
shunt d r i v e r  package, each o f  t he  shunts w i l l  be d r i ven  sequen t ia l l y  i n  
tandem. That i s ,  o n l y  one d r i v e  t r a n s i s t o r  w i l l  be allowed t o  operate a t  the 
maximum power d i s s i p a t i o n  po in t .  A l l  o ther  energized shunts w i l l  have been 
d r i v e n  t o  sa tura t ion .  With t h i s  type o f  design, aided by the  use o f  MOSFETs, 
the  maximum power i n  t he  d r i v e r  package can be contained t o  approximately 10 
watts.  This  design should prov ide  us w i t h  a  q u i e t  bus and a l l ow  us t o  once 
again consider  t he  use o f  t h e  'So lar  Only'  mode. 
V-T CURVES 
Although the  e l e c t r o n i c  coulometer has proven t o  be a  very good charge 
c o n t r o l  method, the  simple vol tage l i m i t e r  has a lso  proven t o  be very e f f e c -  
t i v e .  Using an e m p i r i c a l l y  developed t rans fe r  func t ion ,  t he  vo l tage versus 
temperature (V-T) curve, the  vol tage l i m i t e r  causes ac tua t ion  o f  the  shunts t o  
keep the  b a t t e r y  vo l tage  below the  l e v e l  def ined by t h i s  curve. Figures 8  
through 10 show t h e  V-T curves t h a t  were used fo r  t h e  SAS and AMPTE s a t e l -  
l i t e s .  The more recent  sate1 l i t e s  have used more than one V-T l e v e l  which can 
be se lec ted  by  ground command. The reason f o r  t h i s  i s  t o  p rov ide  f l e x i b i l i t y .  
I f  the  power d i s s i p a t i o n  i s  t oo  h igh  f o r  a  g iven thermal cond i t ion ,  i t  i s  
poss ib le  t o  swi tch t o  a  lower curve which w i l l  r e s u l t  i n  a  lower b a t t e r y  over-  
charge ra te .  Also, when the  b a t t e r y  charge vo l tage increases w i t h  age, i t i s  
poss ib le  t o  swi tch t o  a  h igher  V-T curve i n  order  t o  charge the ba t te ry .  
ACKNOWLEDGEMENTS 
The Author wishes t o  acknowledge t h a t  a  l a rge  number o f  people were invo lved 
i n  the  design and t e s t  o f  these power systems. However, the  c o n t r i b u t i o n s  o f  
t he  f o l l o w i n g  th ree  people were major: 
M r .  A r thur  F. Hogrefe--Design o f  the  CRAM 11 and I 1 1  charge c o n t r o l  
systems f o r  t h e  SAS sate1 1  i t e s .  
Mr .  Stanley Mantel --Design o f  t h e  b a t t e r y  charge regu l  a to r  (BCR) f o r  the  
AMPTE s a t e l l i t e .  
Mr .  Walter E. A l l e n  --Power system design f o r  t he  AMPTE s a t e l l i t e .  
REFERENCE 
Sul l ivan,  Ralph M.: Design and Test o f  t he  SAS-A Power System. Appl ied 
Physics Laboratory TG 1106, 1970. 








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BATTERY TEMPERATURE ( OC ) 
FIGURE 10 AMPTE/CCE BATTERY VOLTAGE LIMITS 
A. Rogers, Hughes Aircraft: The newer s i  1 icon transi s tors  woul d 
probably be a quarter of tha t  in the same current device. 
A. Sullivan, APL: Oh you are  saying that  there ' s  another t rans is tor  I 
should be looking a t .  Okay thank you. 


Nickel Cadmium Charge Control Concepts

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