apollo experience report battery subsystem

8/8/2019 Apollo Experience Report Battery Subsystem

1/23

N A S A T E C H N I C A L N O T E NASA TN D-6976

'>ho*40dz

4v)c

;Ff CASE F I L Ecop,!f &

APOLLO EXPERIENCE REPORT -BATTERY SUBSYSTEMby J. Barry TrontManned Spacecrat Center

ZGSitO 058

N A T I O N A L A E R O NA U T IC S A N D S PA C E A D M I N I S T R A T I O N W A S H I N G T O N , D. C. SEPTEMBER 197


2/23

3. Recipient's Catalog NoIGovernment Accession NoIReport NoNASA TN D-69767. Author(s1

J. Ba rry Trout, MSC

4. Title and SubtitleAPOLLOEXPERIENCEREPORTBATTERY SUBSYSTEM

8. Performing Organization Report No.MSC S - 3 3 3

5. Report DateSeptember 1972+-. Performing Organization Code9. Performing Organization Name and Address 10 . Work Unit No.914-11 10-00- 22. Sponsoring Agency Name and Address

Manned Spacecraft CenterHouston, Texas 7705813. Type of Report and Period CoveredTechnical Note

11 . Contract or Grant No

this page) 21 . No. of Pages 22. Price22 $3.00

National Aeronautics and Space AdministrationWashington, D. C. 20546 14. Sponsoring Agency Code5. Supplementary Notes

The MSC Director waived the us e of the International System of Units (SI) for thi s ApolloExperience Report, because, in his judgment, the us e of SI Units would impair the usefulnessof the repo rt or re sul t in excessive cost.6. Abstract

Exper ience with the Apollo command-service module and lunar module bat teri es is discussed.Significant hardwar e development concepts and hardware test r esu lts a r e summarized, andthe operational performance of batteries on the Apollo 7 to 1 3 missions i s discussed in termsof performanc e data, mission constraints , and basic hardware design and capability. Also,the flight perf orma nce of the Apollo battery char ger is discussed. Inflight data are presented.

7. Key Words (Suggested by Author(s1.Battery Subsystem* Apollo Program.Electrical Power System

9. Security Claqif. (o f this report) 20. Security Classif. INone I None

18 . Distribution Statement

' For sale by th e N a t i o n a l Technical Information Service, Spr ing f ie ld , Virginia 22151


3/23

APOLLO EXPERIENCE REPORTBATTERY SUBS YSTE M

B y J. B a r r y T r o u tM a n n e d S pa c e cr a ft C e n t e rS U M M A R Y

The req uire ment s established for the Apollo command-service module and lunarmodule batteries were well within the existing state of the art for batterie s; hence, nounique p robl ems w ere identified o r experienced during battery development and quali-fication o r during short-time unmanned flights.

The only significant problems resulted from the use of a relative ly new type ofnonabsorbent s epa rat or (Permion 307) in the command module entry and postlandingbatteries and from failure to verify t h e effectiveness of the battery-charging systemfor those batteries. Thes e two fact ors jointly resulted in se ve re undervoltage a t thecommand module main buses at command module/service module separati on duringthe Apollo 7 mission. The final solution of these probl ems for the flight of Apollo 11w a s achieved by reverting to the original absorbent cellophane separator materialand by raising the output voltage of the command module battery charger.With the possible exception of auxiliary battery 2 in the unmanned Apollo 6 flight

(insufficient data to prove a battery failure), no battery failure occurred in any flightthrough the Apollo 16 mission . This is proof that the design- and verifica tion-te stprinciples wer e valid and that proc edur es and crit er ia for acceptance testing at thebattery vendor's facility and preparation for flight at the Kennedy Space Center wereeffective in culling batteries that might fail in flight.INTRODU CTI ON

The batteries discussed herein a r e the operational flight batterie s developed forus e in the Apollo spacecraf t. The re po rt includes a disc ussi on of the command module(CM) entry and postlanding batter ies and battery cha rg er , the lunar module (LM) mainpower batt erie s, and the CM and LM pyrotechnic batteri es. This re po rt does not in-clude a discussi on of the miscellaneous off-the-shelf bat ter ies use d in the vari ous de-velopmental boilerplate flights. A l l Apollo bat ter ies we re of the silver-oxide/zincalkaline type. The LM main power batteries wer e pr im ar ie s (single discharge), where-as the CM batt erie s and all pyrotechnic b atteries were low-cycle life secondari es(rechargeable) . The Apollo 7 to 1 3 missions a r e covered.


4/23

COMMAND-SERVICE MODULE BATTERIESThe unmanned command-service module (CSM) flights used from two to eightentry batteries in the command and servi ce modules, two pyrotechnic ba tteri es in theCM, and two pyrotechnic ba tt er ie s in the SM to supply power to mission-sequencing

Electr icd lc o n t r o l a s w n b l y

,-Aft equiprnent ba y

control devices and se rv ic e module (SM)jettison control devices. The manned ve-hicles used three entry batte ries and twopyrotechnic batteries, all in the CM. Themounting locations of the variou s b att eri esused in manned missions are shown infigures 1 to 3 .

Pyrotechnic bat tery APy ro tec hn i c ba t te ry B

I n v e r t e r 2I n v e r t e r 3I n v e r t e r d c \ Battery charger

7- Ba t te rv A/.-- To battery vent valve onwaste-management pan el

Bat tery vent l i neBat tery C -/ Battery B

Figure 1. - Command module batteryand charger locations in the lowerequipment bay (block 11).

l n v e r

For+w rd

(a) Ascent components.

(b) Descent components.Figure 2. - Locat ions of LM mainelectrical power systemcomponents.

2


5/23

Entry a n d P o s t l a n d i n g Ba t te r i e sThe entr y and postlanding bat teri eswer e used t o supplement the main CSM

syst em gimbaling); to supply cert ain para-power source (that is, the fuel cells) duringpeak loads (for example, se rv ic e propulsionsitic loads that had to be kept separa te fromthe CM main dire ct-c urre nt (dc) buses; and,after CM/SM separation, to supply all CMelectrica l loads.

&I. E ~ ~ I ~ ~ , ~ ~ - ~ ~ ~ ~ ~ ~b a ~2 1

Figure 3. - Locations of L'M explosive- The origina l vers ion of the entr y anddevice batterie s. postlanding battery had a capac ity of 25 A-h.The battery had a welded titanium case and amachined, anodized- magnesium c over. Acoat of beryllium-aluminum paint was applied fo r increased emissivity and for c orro-sion resistan ce. The plate-sepa rator material in the cell s was Permion 307, a n irra-diated polyethylene fil m sele cted becaus e of its thermal resistance. A thermalenvironment of 250" F in th e CM lower equipment bay was expected fo r approximately30 minutes because of e ntry heating. The available energy was 725 W-h p e r battery at29 volts or 2175 W-h p er shipset of thr ee batte ries . The batt ery had to be capable ofsix cha rge s and discharges, the dischar ges consisting of 30 minutes at 35 amp ere sand then 225 minutes at 2. 0 amperes (all at more than 27 volts).

The 25-A-h batt ery wa s qualified but was neve r flown. By mea ns of better defi-nition of the block I1 vehicle battery-energy requirements, it was shown in mid-1964that the requ ired entr y and postlanding energy wa s 2234 W-h, which exceeded the en-erg y availab le even without a battery failure. It was determined that the batteriesshould be capable of 40 A-h (1160 W-h each ) to satis fy vehicle req ui rem ent s with oneallowable batt ery failure. In October 1964, a contractor was authorized to proceedwith battery redesign.

The 40-A-h batt ery als o contained Perm ion 307 plat e-se para tor mat eri al andrequired additional active material to deliver the additional energy.rial was changed to cemented Plexigla s, which was enveloped i n glass-r einf orce d epoxyand covered with gray plastic paint. Also, this batt ery had to yield six completecharge/discharge cycles. The discharge profile was changed to 25 am pe re s for 1 hourand 2 amper es for 7.5 hours; a 25-ampere load was mo re representa tive of spacecraftbatt ery loads. The weight and dimension s of the 40-A-h batt ery are shown in table I.The voltage-current chara cter isti cs of this battery are shown in figu re 4. Thi s batt erywa s qualified and flown on all Apollo missions through Apollo 13.

Th e case mate-

3


6/23


7/23

thermal requirements, a pro gra m wa s initiated to requalify the battery, using conven-tional cellulosic (cellophane) separator material that had an intermediate layer ofplastic film.

A pol l o 7 B a t te ry A 0

A further indication that a change in separator material was required was ob-se rv ed in reduced battery-voltage ch arac teri sti cs afte r prolonged exposure of the bat-te ry to ze ro gravity (fig. 5). The very low voltage shown for the Apollo 7 batteries(relative to those of Apollo 8) wa s caused by the unexpected inability to r ech arg e thebatt erie s fully. Although the reduced voltage-current char acte rist ic could not be repro-duced in ground-based tests, it w a s concluded that hydrogen bubbles, fo rme d by natu ralgassing at the negative plate, displaced elec trol yte between the plates . This phenom-enon is possible because the Permion separator is not absorptive and the hydrogenremai ns in place on the negative plates in zer o gravity. Loss of electro lyte de cr ea se sconductivity, lowering the voltage chara cter ist ic. After qualifying the battery contain-ing cellophane separators, one redesigned battery w a s flown on the Apollo 10 space-craft; on subsequent flights, all th re e batt erie s we re of this design. The improvementobtained from the redesign is shown in figure 6. The absorptivity of the cellophaneelimina ted the zero -gravity effect by retention of the electrolyt e between the plates.

30r

\25 1 I I I 1 I0 5 10 15 20 25

B a t te ry A or B current. A

Figure 5. - Entry-battery flight-performance data at the timeof CM/SM separation.

25 I ' I I 1 I I I I I I2 4 6 8 10 12 14 16 18 20 22Current. AFigure 6. - Comparison of entry-batterycharacterist ics at the time of CM/SMseparation.

Mounting. - The ent ry and postlanding battery w as hung from its top surface inthe lower equipment bay (fig. 7). Active thermal control was not r equired because ofthe mild the rm al environment of the CM cabin.


8/23

Figure 7. -The 100-ampere cir cuitbreaker and entry-battery mount-ing adapter.

P y r o t e c h n i c B at te r yN o significant proble ms we re encoun-te re d in the development of the pyrotechnicbattery. Qualification was completed inJuly 1965. The CSM pyrotechnic batt eryhas been flown successfully on Apollo space-cra ft fro m boilerplate 22 to Apollo 13.Performance requirements. - Thesalient electrical requi rements were as fol-lows. The capacity was 75 amperes for36 seconds at mo re than 20 volts at 60" to143' F o r for 15 secondsat more than20vol tsat 50" ? 3" F. Regarding cycle life, thebattery had to deliver 6 cycl es of 75 am pe re sfor 36 seconds at any time within 36 daysafter activation. The battery als o had tosatisfy the preceding requ irem ents whenactivated after stora ge for 1 gear f rom the

date of manufacture. The batt erie s wer e furnish ed in the dry , charged condition, andthe electrolyte was supplied in sep ar at e contain ers. The weight and dimensions of thisbattery a r e given in table I. The sepa rat ors were thr ee la ye rs of cellophane (only).The battery c as e was a hand-fabricated monobloc of G-10 circuit board, cemented to-gether and lined with an epoxy cement. The cell relief valves, which had crackingpre ssu res of 2 to 1 0 psig, wer e vented into a manifold spa ce that w a s insulated fro mthe intercell connectors; this sp ace had a relief valve that had a cracking pr es su re of30 f 5 psig.Mounting. - The pyrotechnic battery wa s hung fr om its top surface i n the right-hand equipment bay by four bolts (two at each end) that threaded into the st ruct ure.Active thermal control w a s not requi red becaus e of th e mild the rm al environment of

the CM cabin.Spec ia l P ro tec t ive Dev ices

Before the spacec raft 012 fi re, silicone -rubber potting wa s applied to all exposedterminal and cable-lug metallic surfaces after the installation of the CM flight batter-ie s. (The potting was not used fo r te st batter ies. ) A s a result of the review after thefire, it was concluded that not only was th e silicone rubber flam mabl e but that the ex-posed terminals of the test batteries were a spark source i f shorted. It was also con-cluded that additional overload protection was r equi red f or the long cable that connectedeach entry battery to its circuit b rea ker in the battery circ uit-breake r panel (panel 250).Entry battery. - To provide additional c irc uit protection, a 1 0-ampere circuitbrea ker was added at the battery. The batt ery mounting-adapter plate wa s extendeddownward t o provide a bracket so that the 100-ampere ci rcuit br eak er could be con-nected to the batt ery by a ve ry sho rt cable. The use of silicone-rubber termin al pot-ting was discontinued; instead, a polyimide shield w a s designed. The shield wa s


9/23

attached to the mounting bracket and contained integrally molded baffles that blockedac ce ss of st ra y materia ls that might short circuit either the battery o r the circuit-breaker terminals.Pyrotechnic battery. - The use of silicone-rubber ter mina l potting was discontin -ued. Pod-shaped polyimide shields were designed. Each terminal had two shiel d

halves: one shield half was mounted on the termina l behind the connecting cab le lug;the oth er shi eld half, which enclosed the termi nal and lug, snapped onto the first shieldhalf. Becau se the press uriza tion-t est plug on the termin al face of the ba tte ry physi-cally in ter fe red with the installation of the pods, the plug had to be removed.

Entry-Batte ry ChargerThe salient characteristics of theentry-battery charg er were as follows.The alternating-current (ac) input was1 1 5 to 200 volts, 400 hertz, three phase;

the dc input w as 28 volts . The output isshown in figur e 8. The entry-batterycha rge r weighed 4 . 3 pounds (maximum),and the maximum input power was 5 5 wattsac and 84 watts dc. A simplified sc hemat-ic of the charger electronics is shown infigure 9. N o problems occu rred duringthe development of thi s unit. Qualificationwas completed without complications i nFebru ary 1966.

40I. 5 1.0 1.5 2.0 2 . 5

Output current. A

Figure 8. - Charac teristic s ofentry-battery charger.

Battery chargerOFF

...

Currentamplifier Schmit l trigger Voltage amplifier Comparator circu itI tSwitch diode Battery C -1 -

e I ,0 M V dcII). Transformer ~~~~~~r Auxiliary

powerectifierIUPPlY~ circuit

0 c dI0 B \I powerIF supply

115 v II" B A dl - - ti"K I

Figure 9. - Battery-charger functional diagram.


10/23

Operation. - The ch ar ge r was operated manually by the crewmen. When a batterywas to be charged, the charg er w a s turned on, the battery was removed fro m the bat-tery relay bus, and charging power wa s applied to the battery through a selectingswitch and the applicable battery bus (fig. 10).

- I ZDEC-AII I *Voltmeter lpyro bat A )I 20 A Bat bus A to !i vr o Ibus A- , -MESC Abat A IQ O2 pyro bus tie IN.0.1 fue l dump sys Ab u s 6 I r - - ~I SeqB 20 A1 CB17 - ;LDEC BI- 1 , {I-* ZlESC B

*voltmeter ipyro bat B lBat bu s B to \I L---II cB19 I , u f l dump sys B+pyrobustielN. 0.1 -- 1 I Batj Bat A pw rentrv and tostCBl5 Bat C tobat bus A lh.O.1

I landing

II Bal C pwr

I bu s A

Bat CC B43 I C Voltmeter (Bat ci

partlanding 0 EDS 7Bat chqrO A I Bat Bat B pw r entry and p o s t -

I landing 1 bu s B

r i a n i ~F5postlanding I 80 AI Panel 250 IRHEBlL ~ ~ ~ _ _ ~ ~- J

N 0 normall) open-~ _ _ _ _ ~ ~

C B circ uit breaker P w r powerC hg r iha ger Fyro pyrotechnicED5 emergency detection system RHEB rig ht hand eqiiipment baytDEC lu na r docking events contr oller Seq seauencehlESC master event sequence control ler, ~ I D C main display console Sys systemL-~~-~~~ .______-

bus tie

Figure 10. - Battery circuits.Flight use . - Extensive testing and evaluation of t he ch arg er and ent ry batter ywere performed at a contractor facility and at the MSC to define the interdependentchar acter isti cs and rul es for inflight charging timing and termination. The se labora-tory te st s showed that the ch arg er could recha rge the batt eries fully in a time approxi-

mately equal to T = (A-ho/lA) + 0.5, where T is the time required for charging, inhours, and A-ho is the amper e-hou rs removed fr om the battery. This relationshipwas based on terminating charg es when the char ger cur ren t decrea sed to 0 .4 ampere.Typical charging c urv es fro m these tests a r e shown in figure 11.

8


11/23

z.5r

1 I I I I I I I I J0 1 2 3 4 5 6 1 8 9Time, hr

Figure 11.- Entry-battery chargingcurrent as a function of priorampere-hours output.

5c ' \

7 xpected

I A ~ o l l of Aoollo 9I l l 1

O l 0 - a Y 1 I ; ; d ; Q ' 9Hrin

Time

Figure 12. - Charging experience fo rapproximately 10 A-h output.

The first formal test of the C M battery-charging syste m o ccur red during the flightof Apollo 7 (spacecraft 101). The batt erie scould not be recharged. A comparison ofthe expected and the actual pe rfor manc e ofthe Apollo 7 charging system is shown infigure 12. Postflight anal ysis showed thatthe line impedances between the charger andthe battery (previously unevaluated) weresufficient to re duce significantly the c harg evoltage applied at the battery. The char geroutput voltage, although within specifica tionlimits, was on the low side of th e allowabletoleran ces, aggravating the low-voltage con-dition. It was fur ther postulated that thebattery had a higher-than-expected internalresistance because of the gassing effectsmentioned previously.Reduction of l ine imp edances was notfeasible for subsequent flights because itinvolved breaking into al read y fabricat ed,complex wiring ha rn es se s. Reduction ofbattery impedance by using absorben t cello-phane s epa rat ors w as not feasible untilthe Apollo 10 mission . Although it wasproven possible to dril l a hole in the charg erdown to the adjus tmen t sc re w of an outputpotentiometer (raising the charger outputvoltage) and then reseal the charger, thereliability of this adjustment could not be

verified until after the Apollo 1 0 mission.Hence, fo r the Apollo 8, 9, and 1 0 mis-sions, chargers were sele cted that had out-puts on the high side of the specified voltagetolerances. Wherever necessary, ch arge rswere removed from other command modulesand install ed i n the next flight vehicle. This procedure perm itted full recharging but re-quired ex cessivel y long tim es (fig. 1 2 ) . In the Apollo 11 and subsequent vehicles,cha rge rs w ere used that had adjusted outputs permitting charging at the rates expectedbefore the Apollo 7 mission. The small incr ease in voltage required to achieve thisimprovement is shown i n figure 13.

9


12/23

40 I LUNAR MODULE BATTERIES

u>

*3-cmm 3 7 -

36 -

The LM battery complement consist edof two 29-volt, 296-A-h batteries in theasce nt stag e and four 29-volt, 400-A-h bat-teries in the descent stage. The LM origi-nally was configured fo r fuel-cell powergeneration. Development of the LM fuelcel l was begun with the c ont ract or who de-veloped the CSM fuel-cell syst em. The LMfuel cell w as to have the following genera lcharacter ist ics .1. Each of t hre e fuel cells was toproduce 32 kWh at 900 watts within the bandof 27 to 36 vo lt s.

I I I I 2. The minimum load wa s to be\ * .' 100 watts.5 5 1.0 1 . 5 2. 0Current, A3. The peak load wa s to be 1125 wattsf o r 1 hour.igure 13. - Char acte rist ics of entry-battery chargers on the Apollo 7nd 11 missions. 4. The fuel cell was to accept tra n-sient load changes of 450 wat ts within volt-age regulation.

5. Each fuel cell wa s to weigh 76 pounds (maximum).6. The fuel cell s wer e to be the open-cycle-operation type.

Regarding open-cycle operation, hydrogen was to be u sed both as a fuel and as a coolant,without recycling fuel. By thi s means, each fuel cell was not to dissip ate mor e than75 Btu/hr at 160" F to the str uctu re. By mid-1964, significant technical difficultieswere being experienced in both the CSM and LM fuel- cell developments. Additionally,the specified tim e fro m lunar lift-off to docking with the CSM wa s redu ced m or e than20 hours. The resulting decrease in energy requirements made a battery system fea-sible.the effect of converting to a silver-oxide/zinc alkaline primary- battery power source.Thi s type of bat ter y was consi dered to be reliable, and was avai lable off the shelf rela-tive to fuel cells, and had been flight proven during Proje ct Mercury and the GeminiPro gr am. In the study, it was shown that a battery system could be developed th atcould provide elect rical power fo r a 35-hour lunar stay but that would r esul t i n approxi-mately 1 00 pounds mor e of powe r-syst em weight than would the fuel-cell syst em. Adecision was made to accept this penalty as being favorable over the developmental andreliability risk s assessed against the fuel cells. Therefore, the LM fuel-cell programw a s reduced in April 1965 and terminat ed on June 30, 1965 (concurrent with the award-ing of the LM battery subcontract). The battery -system configuration that wa s select edwa s composed of two 8.7-kWh bat ter ie s in the LM asce nt st age and four 11.6-kWhbatteries in the descent st age (a total of 63.8 kWh of installed energy).

A s a consequence, a feasibility study was initiated by a contractor to evaluate

1 0


13/23

Bat tery development began in Jun e 1965 and culminated i n battery qualification byMar ch 1968. In addition to 20 monthly pro gr ess re po rt s, special report s were issuedregarding ascent- and descent-battery efforts on the following subjects.1. High-temperature performance te sts2. Electr olyte volume and cell position3 . Thermal-runaway test4. Short-circuit test (descent battery)5. Batter y-cont ainer weight-reduction study6. Thermal - and ac-impedance te st s7 . Specific-heat determination8. Regression analysis of proce ss and usage fact ors

Capacity-prediction re por ts w ere issued on the following subjects.1. Transient-voltage performance2. Separator- materia l selection and evaluation te st s3 . Systems simulation tests4. Ascent- and descent-battery load sharing5. Pa ram et ric testing of electri cal char acte rist ics6. Qualification test7 . Postqualification te st8. Supplemental qualification test

An integrated thermal /electr ical computer model of the ascen t and descent batterie s(as i f mounted on the cold rails and surrounded by the LM s truc ture in space vacuum)wa s genera ted and was verifi ed by the use of in strumen tation on the LM-1 and LM-2flights.The only significant problem encountered in the development and qualification ofthe ascent batt eri es was undervoltage. This problem was encountered during larg e loadtransi ents r elated to single-battery abor ts and abort- sta ge maneuvers. The problemwas resolved by better definition of permissible maneuvers and by requiring "condition-ing" dis cha rges of 2 . 5 to 20 A-h at low loads immediately before mi ssion stag es in-volving possible lar ge load inc rea ses on the ascent batteries. This procedure preventedundervoltage transients by a mechanism that is not fully understood at this time.

11


14/23

The only significant problem encountered in the development and qualification ofthe descent batteries was random-vibration-test failures of the magnesium-alloy bat-tery container. Cracks occurred in the area where the ends and sides of the containerwe re welded (just above the mounting rail). The fail ure mode wa s elimi nated by addinga second gusset a t the end of the mounting rail (fig. 14).

h e l d crack a r e a

A d d e d gussel-../

(a) Fai lur e configuration. (b) Corre cted configuration.Figure 14. - Fail ure and correc ted configurations of the battery container.The explosive-devices (ED) battery was identical to the CSM pyrotechnic b attery,except that the pressurization port was retained in the ED battery. The ED battery wassubjected to qualification te st s (in addition to those re quir ed f or CSM us e) to verify the

ability of the battery to withstand LM-vibration spectra and thermal vacuum in the cold-plated, thermally wrapped configuration. N o difficulties relat ive to design o r operationwe re encountered.

A s c e n t-S t a g e Ba t te r i e sThe requirements for the ascent-stage batteries were as follows. The specifiednominal capacity wa s 296 A-h within 28.0- to 32. 5-volt l imit s and at rates equivalentto 350 to 1600 watts. The specified single-batter y abo rt capacity was 268 A-h within27. 5- to 32.5-volt limits and at ra te s equivalent to 2200 to 3000 watts. The acceptance-te st capacity (sample cells only) was 296 A-h a t 50 am pe re s to a minimum of 1. 41 volts

per cell. It was required that the specified nominal and abor t capacities be deliverableafter idle stand for 30 days at 80" F. The required capacity on the acceptance test wasdetermined after a 10-day charged stand at 95" F. The specified weight was a maxi-mum of 131.7 pounds (activated), and the quali fied weight was 123. 7 pounds (activated).Dimensions and other pertinent data a re summarized in table I . The battery wasmounted in the vehicle in su ch a way that the plane of the pla tes wa s perpendicu lar tothe, natural gravity vect or (earth or moon). Originally, it was required that the battery1 2


15/23

yield one 1 0-percent-capacity disc harge and (with re ch ar ge s) two 85-percent-capacitydischarge s. The requirement for the las t two cycles w a s deleted because no chargingis required in flight.

De s c e n t -S ta g e B a t te r i e sThe requi remen ts for the descent-stage battery were as follows. The specifi ednominal capacity was 400 A-h within 28. 0- to 32. 5-volt limits at a 25-ampere dischargerate. A tapped connection was provided fr om the 17th cell of the batte ry to preventapplication of overvoltage to the LM sy st em s at low loads when the battery w a s at near-full capacity and could deliver the high silver dioxide voltage on discharge at low cur-rents. In a typical mission, the 17th-cell tap was used to supply small loads forhe at er s in the LM only fr om prelaunch until after CSM transpositi on and docking.After this , these lo ads wer e supplied fr om the CSM through the CSM/LM umbilicaluntil LM powerup, at which time the LM loads were high enough to d ep re ss the voltageto acceptable levels. Use of the full battery or the 17th-cell tap was controlled manuallyby the LM crewmen. The specified contingency (one of four bat ter ies failed) capacitywas 389 A-h within 28. 0 to 32. 5 volts at ra te s equivalent to 1 5 to 1330 watts. The con-

tingency mission requirements resulted in higher individual battery loadings and ashortened luna r staytime. The acceptance-test (sample cel ls only) capacity was400 A-h at 25 amp ere s to a minimum of 1.4 1 volts pe r cell. The specified capaciti eswere required to be delivered after the sa me wet-stand time periods and tempera ture sas wer e used for the ascent batteries . The specified weight was 139.6 pounds and thequalified weight w a s 132.7 pounds. Dimensions and othe r pertinent data are summa-rized i n table I. The mounting orientation and the cycle-life information wer e the sa meas for the ascent-stage batteries.

Explos ive-Dev ices Bat teryThe ED battery was identical to the CM pyrotechnic batter y except that the pre s-surization por t was retained in the ED battery. The us e environment in the LM differedfr om that in the CSM; thus, the LM ED battery w a s mounted on cold rails outside thepres suri zed cabin and w a s wrapped in a thermal blanket to limit temperature extremes.The ED batt ery supplied power to the explosive devices that enabled such functions asstaging, landing-gear extension, helium- and fuel-tank pressurization, and s o forth.

F l i g h t P e r f o r m a n c eThe LM ascent, descent, and ED batt erie s per for med in accordanc e with specifi-cations and satisfied the electrical-energy requirements for all LM flights. The

voltage-current cha rac ter ist ics of these batteri es a r e shown in figures 1 5 and 16. Theperformance of these batteries w a s noteworthy during the r es cu e phase of the aborte dApollo 13 mission. The bat ter ies supplied total vehicle power for 83 hou rs under con-ditions for which they had not been qualified o r tested; that is , low electrical loads(350 watts comp ared with a nominal 1000 watts), continuous ze ro grav ity while loaded,and continuous low temperature (approximately 37" F). Telemetry measurementsindicated that two of the descent bat ter ie s were use d 5 to 6 percent above the specifiedcapacity. A CM entry battery w as charged fro m L M battery power in preparation forCM/SM sepa rati on and CM entr y.13


16/23

10090 -8 0 -70-60 -50 -40-30 -

-

2 0-acalL2 10 -5 7 -

9 -c 8 -

6 -5 -4 -3 -

2 -

o L M - ~ l ight data0 LM-3 fl ight data3 LM-4 fl ight dataD Qual i f icat ion-testA LM-5 f l ight data

D

Temperature range : 5' t o 8 5 ' FCurrents between 1 an d 1 0 A were measured after dischar gingapproximately 20 percent of the available capacityper battery

1 L . U I I I22 24 26 28 30 32 34 36Battery voltage, V dc

100-90 -80 -70 -6 0 -50 -40-30-

2 c -accE l 0 -" 9 -c 8-x 1 -m 6 -

5 -4 -3 -

2 -

00

aD0

I ILM- I f l ight dataLM- 3 f l i gh t da taLM-5 f l i g h t da taQual i f icat ion-test dataLM-4 f l i gh t data

I '1 1emperature range = 50- to 75 FCurr en ts be tween 1 an d 10 Awere measured after dischargingapproximately 20 percent of the I ,available capacity per battery. I

0

I22L4 26 A8 30 32 34 36Battery voltage, V dc

(a) Ascent battery. (b) Descent battery.Figure 15. - Lunar module battery characteristics.

I I I I0 25 M 75C u r r e n t . A

Figure 16. - Characterist ics of the pyrotechnic and ED battery.14


17/23

COMMAND MODULE BATTERY-FAILURE EXPERE n t r y a n d P o s t la n d i n g B a t t e ry

ENCE

With the possible exception of the Apollo 6 auxiliary battery 2, fo r which no telem-etry data are available for verification of an inflight malfunction ("Operational per for m-ance, ? * p. 4), no inflight failure of an entry and postlanding battery has occurred.Ground-based failures include the Permion-separator failures (previously discussed)and the following problems .

Low voltage occu rre d after vibration tes ts during qualification testing. The volt-age dec re as e was caused by groundpaths that resulted from electrolyte leakage fro mcracked cell cases. This problem was solved by reducing the cell-relief-valve pr es su refrom 50 to 30 psig and increasing the internal corner radii i n the molded cell cas es(that is , making the corners thicker and stronger).An incremental qualification test was performed on the en try and postlandingbattery to determine i f it could withstand the incr eased vibration lev els of th e SM for

us e without shock mounting in spacecra ft 009. The battery failed because of cell-casecracks; thus, it was shock mounted for flight.The two terminal-end cells of a battery in a laboratory became grounded by elec-trolyte leakage through cell-ca se cr ac ks located direct ly behind the battery vent elbow.Thi s fa ilu re was the re su lt of mishandling the elbow, which caused it to be jammedbackward into the cell ca ses .

P y r o t e c h n i c Bat te ryN o inflight fa ilur e of a C M pyrotechnic battery has occur red. None of the ground

fai lur es was significant. to battery design o r performance capabilities.B a t t e r y C h a r g e r

N o recorded fa ilur es of the C M battery charger have occurred. Two chargerswer e damaged during laboratory te sting because of inco rre ct connection of ac inputpower.

LUNAR MODULE BATTERY-FAILURE EXPERIENCEMany L M battery failures have occurred; therefore, the failure s are summarizedin broad groupings.

1 5


18/23

Q u a I f c a t o n T e st F a iIu e sQualification-test failures are summarized as follows. Descent -batte ry weldcr ack s occurred during the random vibration testing; the solutions to this problemincluded the following.1. On t h e basi s of data obtained fr om vibratio n testing of an L M test vehicle, thevibration levels w ere reduced.2. The inspection procedur es w er e revised.3. A gusset w a s added on the mounting rail.Descent-battery cell-cas e cra cks occu rred after a -20" o 160" F hermalshock (unactivated) because of differential t her mal expansion of the ce ll c as e and ther -mal fins i n the battery ca se to which the cell s wer e potted. The following changes weremade.1. The lower tempera ture w a s changed from -20" to 20" F, based on the abilityto control shipment and sto rage te mper atur es easily within this limit.2. The cell-case w a l l s we re lubricated to prevent adhesion to the thermal fins.An ascent-battery voltage transient occurred below 28.0 volts (27.6 volt s) whenthe thermal-vacuum discharg e load w a s changed fr om 50 to 80 amperes. The specifiedvoltage minimum w a s reduced to 27.5 volts, based on a review of the capabil ities ofthe voltage-limited equipment operated from the battery. The ED battery w a s intern-ally grounded because of el ectr oly te leakage (caused by impro per addition of e lect rol yteduring activation). The activation procedur e and equipment were improved.

Accep tance-Tes t Fa i u e sN o ED-battery fail ures have occurred. However, two ascent- batter y fai lur esoccu rre d: one because of low insulation resi st anc e between the battery ter mina l andthe cas e, caused by inadequate drying of potting mate rial, and the second because ofbroken case s on sample cells , caused by vibration overtest. A s of December 1969,26 descent-battery fai lure s had occurred. A ll fail ures except one we re caused by im-proper processing, improper handling during fabrication, o r improper testing. Theone exceptional failure (low capacity) is still unexplained.

F a i l u r e s o f F l i g h t B a t t e r ie s U se d in V e h i c l e T e s t i n gA l l but one of the 15 eported ED-battery failures were caused by improperusage (that is, use beyond specified life or im pr op er handling). In the one exception,the battery had a groundpath fr om terminal to c as e that could not be analyzed becausethe battery was discharged completely through an external load after discovery of theground. No significant ascent-battery problems occurred . Six descent-battery failuresoccurred because of cells t h a t leaked electrolyte when the batteries were stood on endduring activation o r while the batteries were in a vehicle at the Kennedy Space Center.

16


19/23

The leakage was attributed to e xcess free electrolyte cau sed by a reduction of plateporosity in long-term stor age at temp eratu res between 50" and 9 0 " F. Requirementsfo r refr iger ated storage were imposed. Also, a battery used in L M - 4 tests had lowvoltage in one cell after discharging 320 A-h in the laboratory before being return ed tothe vendor for u se as scrap. A lump of mate ri al found on a plate lug had punctured asep ara tor , causing loss of ce ll capacity by means of a shor t between the positive andnegative plates. Clos er inspection by the vendor and hand cleaning of plate tabs wereinstituted to elimina te this type of f ailure.

F a i l u r e s of Ba t te r i e s P re p a re d for FlightN o ED-battery failures occurred. The one ascent-bat tery fail ure was caused byan apparent cra ck in one plastic cell cover at the flow mar k nea r the fill screw. Theflow mark was made to leak air slightly after extreme pr ess ure cycling during thefai lur e investigation. Th er e had been no leakage before cycling. An inspect ion pointwas added to the test and checkout procedu re to ens ure that possible "leakers" wouldnot get pas t the battery shop. One descent-battery fai lure occur red because a cell wasinstalled in the reversed-polarity position in a battery, lowering the battery voltage by

the equivalent of two cells.open-circuit voltages lower than required ( 3 7 . 0 volts) by 0.005 volt or less. The open-circuit-voltage cri teri on was reduced to 3 6 . 9 8 volts. N o battery fail ures occurred inflight. The incident involving batt ery 2 on the Apollo 1 3 flight (following section) didnot diminish the abil ity of tha t battery to continue to deli ver power in a normal manner.

Three fail ures occurred because activated batteri es had

APOLLO 13 BATTERY EXPERIENCEDuring the Apollo 1 3 cryogenic-oxygen-system fai lur e and the subsequent inabil-ity t o oper ate the CSM fuel cells , the CSM first obtained power (approximately

400 watts) fr om one entr y and postlanding battery. Subsequently, the CSM/LM assem-bly was switched onto LM-battery power. The spa cec raf t was powered down to3 5 0 watts. Inte rmit tent spikes of an additional 7 0 watts occurred as heaters cycled onand off. Approximately 9 7 hours into the mission, the L M pilot reported "a littlethump" in the LM descent stage . Within a few minutes, he repo rted snowflakes in thearea of descent-stage quad IV , i n which descent batteries 1 and 2 we re mounted. Fr oma review of th e telemet ry data., it was concluded that, at the time of t he re port ed thump,a n apparent sho rt circu it of approximately 1 0 0 to 150 amp ere s occurr ed between thebattery 1 and 2 electrical control assembly and the battery 2 ground. The battery 2malfunction light subsequently cycled on and off fo r approx imately 1 day, even thoughall batteries per for med nominally fo r the remainde r of the mission. This malfunction-light cycling was att ributed to a malfunction of the battery 2 overtemperatur e- sensingcircuitry.se ns or s do not activate at temperatures less than 140" F.The battery tempe ratur e was approximately 37 " F, and the overtemperatur e

From the foregoing information, it was inferred (but never proved) that electro-lyte escaped from some cells in battery 2 , was displaced to the termin al end of thebattery by the tran seart h descent-engine burn, and shorted a positive terminal to thebattery case (mounted to a grounding structure ). This s hor t could have ignited the hy-drogen and oxygen mixture under th e battery cover , causing the thump and an efflux


20/23

of electroly te fr om the battery, which produced snowflakes. The capability of such ashor t to carr y the estimated fault curre nts was verified by a battery test performe d bya contractor.On the pre mi se that the postulated failure sequence might be valid and becausethe quantity of el ectro lyte in the descent-ba ttery ce ll s was considered to be higher thanwas n eces sary to obtain full battery capacity on discharge, a prog ram of corre cti veredesign and investigation was initiated. The significant findings and design improve-

ments a r e summa rize d in the following sections.F i n d i n g s

Twenty cubic centimeters (5.6 perc ent of 360 cub ic' cen timeter s) of the electro-l y t e in a descent-battery cell could be removed fr om a cell without affecting its abilityto deliver capacity (420 to 450 A-h using 360 cubic cent ime ters comp ared with the speci-fied 400 A-h). This reduction of elec tro lyt e volume bec ame a design change.Cells that had plates mor e than 280 days old (from completion of the plate-makingproc ess es until the day of activation), whether st or ed ref rigera ted or unrefrigerated,

absorbed significantly le s s electrolyte than did younger cel ls and tended to leak. Cellsthat had plates younger than 280 days caused no problems. No work was done o r isplanned to identify and define the mechanism of th is plat e behavior.While on discharge, silver-oxide/zinc pr ima ry ce lls can evolve significantamounts of oxygen. Volumes grea te r than 50 perce nt of the effluent cel l gas we remeasured. Work is planned to identify and define the ba si s for th is behavior, which isbelieved to be related to the plate-formation process, formation-bath temperature andcomposition, drying conditions, and plate st orage ti me and temp erat ure beforeactivation.

D e s i g n Im p r o v e m e n t s

18

The inside su rfa ces of the battery container and lid wer e coated with two la ye rsof insulating, alkali- res ist ant paint. Individual cel l vents we re manifolded togetherand were vented external to the battery through a relief valve. Thi s design provided apath overboard of the batte ry for any elect rolyte that leaked fro m the cell vents andsubstantially reduced the volume of any detonatable ga s mix tu res that ca me fro m in-side the cells. An analog tem pera ture s ens or was removed fr om the batter y becauseit was no longer used in flight and because it complicated potting of the inside of thebattery. To have both ascent-battery t er min als at the sa me end of t he batt ery (whichwas a single row of 20 cells), the terminal lead fro m the cell at the far end of the bat -te ry ra n the full length of the inside of the batt ery and emerge d at the terminal end.In lieu of this arra ngem ent, the design was revi sed to have a terminal at each end of thebattery. This redesign eliminated both an internal potting complexity and a possiblesou rc e of internal short s. The internal battery potting, which had previously coveredall current-conducting pa rt s but had not adhered reliably to some parts, was revi sedS O that it adhered proper ly and extended ov er the e ntir e top s urf ac es of all cells. Thispotting reduced the possibility of internal grounds cau sed by escaped electro lyte.


21/23

A s an improvement to the overall Apollo spacecraf t, an LM descent battery w asadded to subsecpent service modules (fig. 17) . This batt ery could have provided12 kWh of emer genc y energy and was designated the SM auxilia ry battery. It couldbe connected to the C M main buses through the distribution syst em for fuel cellnumber 2.

Figure 17. - Location of auxi lia ry battery in sec to r IV of the SM.C O N C L U D I N G REMARKS

The design features and methods of development, qualification, acceptance test-ing, activa tion, and prei nsta llat ion checkout of all the Apollo command-service moduleand luna r module bat ter ies have prevented inflight fail ures , with the possible unprovedexception of auxi lia ry bat te ry 2 on the Apollo 6 spacecraft.

1 9


22/23

Two major conclusions, both based on the Apollo 7 problems, a r e made. First,untried and novel components (such a s new sep ar at or m ate ri al s) should not be used inbat ter ies unless no other feasible, standard altern ative exists. Second, each space-craft system design (such as charging cir cui ts) must be verified quantitatively, eithe rin a flight vehicle o r in an accu ra te simulation, to avoid unexpected developments inflight.

Manned Spacecraft CenterNational Aeronautics and Space AdministrationHouston, Texas, April 11, 197291 -11 10- 0- 2

20 NASA-L angley , 1972- 1 s- 3 3 3


23/23

n v n I I V I X M L a C t T W I X M U 1 IC.2 MIIU JTMbC f iYM13 I RH IUNWASHINGTON. D.C. 20546 [SOSTAGE AND FEES P A I DNATIONAL AERONAUTICS AN DFIRST CLASS MAIL SPACE ADMINISTRATIONFFICIAL BUSINESSPENALTY FOR PRIVATE USE $300

NASA 451

If Undeliverable (SectionPostal Manual ) Do NOI

The aeronautical and space activities of the United Stares shall beconducted so as to contribute . . . to the expansion of human R n w l -edge of phenomena in the atmosphere and space. Thk Administrationshall provide for the widest psacticable and afpropriate disseminationof inforiiiation concerning its activities and the results thereof!-NATIONAL AERONAUTICSND SPACE ACTOF 1958

NASA SCIENTIFIC AND TECHNICAL PUBLICATIONSTECHN ICAL REPO RTS: Scientific andtechnical information considered important,complete, and a lasting contribution to existingknowledge.TECHN ICAL NOTES: Information4ess broadin scope but nevertheless of importance as acontribution to existing knowledge.TECHNICAL MEMORANDUMS:Inform ation receiving limited distributionbecause of preliminary data, security classifica-tion, or other reasons.

TECHNICAL TRANSLA TIONS: Informationpublished in a foreign language consideredto merit NA SA distribution in English.SPECIAL PUBLICATIONS: Informationderived from or of value to N ASA activities.Publications include conference proceedings,monographs, data compilations, handbooks,sourcebooks, and special bibliographies.TECHNOLOGY UTILIZATIONPUBLICA TIONS: Information on technologyused by NASA that may be of particularinterest in commercial and other non-aerospaceapplications. Publications include Tech Briefs,~~~~~l~~ utilization R~~~~~andTechnology Surveys.

CONT RACTO R REPORTS: Scientific andtechnical information generated under a NASAcontract or grant and considered an importantcontribution to existing knowledge,

Details on the availability of these publications may be obtained from:SCIENTIFIC AND TECHNICAL INFORMATION OFFICE

NATIONAL AERONAUTICS AND SPACE ADMINISTRATIONWashington, D.C. PO546

apollo experience report battery subsystem

Documents