apollo experience report power generation system

8/8/2019 Apollo Experience Report Power Generation System

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N A S A T E C H N I C A L N O T E NASA TN D-7142

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APOLLO EXPERIENCE REPORT -..".,. , , . .,i . , . - . i . .POWER GENERATION SYSTEM

Munned Spucecrufi CenterHozlstotz, Texus 77058

N A T I O N A L A E R O N A U T I C S A N D S P ACE M I N I S T R A T I O N W A S H I N G T O N , ' 6 . C. M A R C H 1973


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1. Report No. 2. Government Accession No.NASA TN D-7142APOLLO EXPERIENCE REPORTPOWER GENERATION SYSTEM

4. Title and Subtitle

7. Authork)David Bell 111and Fulton M. Plauchb, MSC 1 10. Work Unit No.

3. Recipient's Catalog No.

5. Report DateMarch 19736. Performing Organization Code

8. Performing Organization Repo rt No.MSC S-347

9. Performing Organization Name and Address

12. Sponsoring Agency Name and Address

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

Technical Note

11. Contract or Grant No .

17. Key Words (Suggested by Author(s)IElectr ica l Power Generation System'Fuel Cell

18. Distribution Statement

National Aeronautics and Space AdministrationWashington, D. C. 20546

21 . No. of Pages19. Security Classif. (of this repor t) 20 . Secu rity Classif. (o f this page)None None 20

14. Sponsoring Agency Code

22. Price$3.00

15. Supplementary NotesThe MSC Director waived the use of the International System of Units (SI) fo r this ApolloExperience Report because, in his judgment, the use of SI Units would impair the usefulnessof the r eport or r esu lt in excessive cost.

16. AbstractA comprehensive review of the design philosophy and experience of the Apollo elec tri ca l powergeneration system is presented. The review of the system covers a period of 8 years, fromconception through the Apollo 12 lunar-landing mission. The program prog ressed from thedefinition phase to hardware design, system development and qualification, and, ul timately, tothe flight phase. Several problems were encountered; however, a technology evolved that enabledresolution of the problems and resulted in a fully manrated power generation system. Theseproblems are defined and examined, and the corrective action taken is discussed. Several rec -ommendations are made to preclude simila r occurrences and to provide a more reliable fuel-cellpower system.


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APOLLO EXPERIENCE REPORTPOWER GENERATION SYSTEM

B y Dav id Be l l I I a n d F u l t o n M. P l a u c h 6M a n n e d S p ac e cr af t C e n t e rS U M M A R Y

The Apollo power generation system consisted of t hr ee fuel-cell modules that weredesigned and qualified to generate electrical energy at a ra te of approximately 2 kilo-watts for a 14-day mission. The fuel cells used cryogenically stored hydrogen and oxy-gen as the reactants, which were supplied to the fuel cel ls in the form of high-pressuregases. Therma l control of the fuel cel ls was provided by the use of independent therma lcontrol sy st em s that rejected heat fro m radiato rs mounted on the skin of the spacec raft.

The Apollo fuel cell was the fir st pract ical application of a high-temperaturehydrogen/oxygen fuel cell. The design, because i t w a s selected a t a n early develop-mental stage, resulted in many inherent system-design deficiencies that made the fuelcells sensitive to operator e r ro r. Thus, many operational problems were encounteredthat had to be identified and resolved. The eventual solutions to the problems creatednew functional modes and bet ter defined nonoperating modes of the fuel cell s.Hardware development was difficult in fuel-cell design because of the functional

environment of the components. For example, electrolyte seals had to be developedthat would maintain their shape and sealing capability at 500" F in a solution of potas-sium hydroxide. Also, a lightweight, highly reliable, long-life pump had to be devel-oped that would operate a t 200" F in a 60-psi wet-hydrogen environment. Additionally,the development of a highly rel iable low-power glycol pump that could function at tem-pe ra tu re ex tremes of -40" to 180" F w a s required.After the fuel-cell components were designed, built, and tested, quantity produc-tion was necessary fo r support of the flight program. A new set of problems (such asprocess controls , contamination, sp ar es production, and traceability) evolved. Al-though these were not considered major problems, careful control wa s required so thatreliable f light-qualified hardware w a s produced.The Apollo space flights were relatively free of fuel-cell failure s. One problemthat was noted early in the program was related to the secondary coolant loop. A i r andparticulate contamination held immobile i n a one-g environment was relea sed duringthe low-gravity phase of the flight. This phenomenon resulted in the loss of thermal-control capabilities for the secondary coolant loop. New checkout techniques and pro-cedures minimized these conditions on future flights. An oscillating condenser-exittemperatu re was another flight problem. After many tests and analyses, the cause ofthe problem was determined to be another low-gravity effect that was not detrimenta l


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to the fuel cell. Corrective inflight operations that could be used to control such oscil-lations were determined but never were required.

I N TR O D U C TI O NIn 8 years, the Apollo P rogr am elect rical power syste m prog resse d fr om thedefinition phase to the hardware design, syst em development and qualification, andultimately, to the flight phase. Although a number of problems were encountered, res-olution of the prob lems result ed in a satisfactory electrical power syste m.The primary power generation system (PGS)hat was developed had to generateelec trica l energy at a rate of approximately 2 kilowatts for a 14-day mission. The sys-tem had to pos sess adequate mission flexibility; fo r example, no constr aints could beimposed on launch dates, photographic reconnaissance time, o r navigational observa-tion time. In addition, it had to be operationally adaptable to changing requirementsfor successive missions without a subsequent requiremen t for design changes.High reliab ility and safety that were consistent with sys tem weight were prime

considerations. The concept of redundancy grea tly increased the reliab ility of mos tcandidate syst ems. Factors affecting reliability, such as multiple sta rt s, were to beavoided, and simplicity of design w a s desired.In addition to normal operational requ ireme nts, other requirements we re imposed.The requiremen ts involved operational considerations fo r the crewmen (minimum noiseand vibration in addition to operational simplicity), integration with other vehicle sys-te ms , high heat-rejection temperatur e for heat-engine concepts, syst em volume, mini-mum development time , and minimum cost . After a number of electrical powersyst ems were considered, the fuel cell syste m ultimately w a s selected. The advantagesof the fuel cell syst em included the fuel -cel l developmental sta tus and availability, therela tive ly light weight of the system, and the grea te r degree of flexibility than existed

in the other sy stems that were considered. The us e of fuel cells in place of so la rarrays resulted in simplified launch preparations and rendezvous maneuvers andthe use of less fuel for attitude-control maneuvers. Also, extended periods of non-so la r orientation were possible for guidance fixes and star tracking. Batt erie s wouldbe required to supply peaking power during high-load-demand periods . In addition,power t o meet load demands during the entry phase ( afte r the fuel cel ls were jettisoned)would be supplied by entry bat ter ies .After fuel cells were selected as the PGS, he mission requirements for electri-cal energy were refined , and a procuremen t specification stipulated that the PGS wouldhave the capability of generat ing 575 kW-hr of elec trica l energy from thre e fuel cells ata minimum rat e of 563 watts and at a maximum rate of 1420 wat ts per fuel-cell module.

DES IGN C O N S I ERATlONSBefore the design for the Apollo PGS could be selected, the operational require-ments had to be known. After the requ iremen ts were determined, a number of

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electrical power system designs were developed. The design that contained thebest features was selected for further refinements.

R eq u emen sThe Apollo PGS had to generate, supply, regulate, and condition all electricalpower required by the command and service module (CSM) for the duration of a full

mission, including recovery but excluding the life support syst em requirem ents. Thespecif ic requir emen t was the generation of elec tri cal energy at a rate of approximately2 kilowatts fo r a 14-day mission.

System Select onIn the process of selecting a power system to meet these requir ements, all knownsolar, nuclear, and chemical conversion techniques were investigated individually; ap-pro pri ate combinations of individual sys tem s al so we re consider ed. Many designs thatwer e suggested by vendors and NASA we re studied and rate d with resp ect to weight,reliab ility , safety, power capability, tolerance to the mission-environment profil e, andmis sion flexibility. With resp ect to developmental stat us and availability, a require-ment was that all designs were t o be f ree of any known development- improvement re-quiremen ts that might ser iously impact program cost s and schedules.

Sys tem Desc r i p t ionThe syst em that was selected wascomposed pf thre e alkaline fuel cells, oneof which is shown in figur e 1. The sche-matic diagram of the PGS (fig. 2) containsviews of t he thr ee fuel-cel l modules thatare integrated with two hydrogen-storage

tanks and two oxygen-storage tanks. Hy-drogen and oxygen are used as the fuel andoxidizer. The resultan t electrochemicalreaction in the fuel cell produces electricityand water. The water is a byproduct andis used for the environmental control sys -tem for cooling and is consumed by thecrewmembers.Each Apollo fuel-cell module pro-duces direct- current electri cal power overa norm al rang e of 563 to 1420 watts at avoltage of 27 to 31 volts. The module is44 inches high and 22.5 inches in diame-t er and weighs approximately 245 pounds.Each module is composed of four dis-tinct sections or systems: an energy-conversion section, a reactant-control Figure 1. - Apollo fuel-cell module.

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IIIIIIIIIIIIIIIIIIIIIIIIIIL-

Figure 2 . - Simplified power generation system schematic.section, a thermal-control and water-removal section, and the neces sary instrumenta-tion. Except for the energy-conversion section, all of these sections ar e included inthe accessory portion of the fuel cell.

The energy-conversion sect ion, shown in figure 3 , cons ists of a cell stack com-posed of 31 alkaline series-connected cel ls and the associated gas manifolds and con-necting leads. The energy-conversion section is housed in a press urize d jacket that isin an insulated support assembly. The single-cell assembly (fig. 4) converts the chem-ic al energy of the reac tants to electr ica l energy and byproduct water. It consists of twohalf cells; hydrogen is the anode and oxygen is the cathode. The elec troly te between thehalf cells is an 80-percent aqueous solution of potassium hydroxide (KOH). A Teflonseal around the cell contains the electrolyte in the c ell and provides electr ica l insulationbetween the diaphragm port ions of the elec trodes . The cell operates at temperaturesbetween 380" and 460" F and at reactant pre ssures of approximately 60 psia.

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Figure 3 . - Energy-conversion section.

The primary coolant loop contains hydrogen and water in the fo rm of superheatedsteam at 60 psi and 430" F. The excess steam is condensed in the 160" F condenser.Then, the liquid/vapor/gas mixture ent er s the hydrogen pump/separator unit (fig. 6)that se par ates the liquid water out of the st re am by centrifugal action and pumps t hegaseous hydrogen back to the fuel-cell stack . The wat er- separa tor portion of the pumpis shown in cr os s sect ion in figure 6. Gaseous hydrogen and water e nter the rotating

The components that for m the ac ces -sory section are mounted on a Y-frame(fig. 5). The accessory section consistsof a nitrogen pressurization system, three

fittingId

(electrical insulator)MoIyWenumlmanganese slurry, fired

Tubing insulatorlconnectorFigure 4 . - Single-cell assemb ly.

regul ators , a-p rim ary loop (hydrogen andwater vapor), a secondary loop (glycol and water ), heat exchangers, motor-drivenpumps, and plumbing. A condenser connects the two fluid loops.

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Figure 5. - Fuel-cell accessory section.separation scr ee n because of s urf ace tension and is forced down into the pitot-tubepickup are a by the centrifugal force that is generated by the rotating s cr ee n and housing.The pickup end of the pitot tube is stationary and is submerged in the liquid water. Therotation of the water along with the rotating housing produces a 2-psi r am pressur e inthe pitot tube; th is pr es su re is trans mitted by the water to the water-removal-valvediaphragm and spring. This proce ss causes the valve to open and water to be expelledfr om the se parat or into the pitot tube. If a gas bubble en te rs the pitot tube, the ra mpressure is lowered sufficiently to cau se the water- remov al va lve to close. Then, thegas bubble is forced through the pitot tube and out the unsubmerged end. The actua loperation consists of a continuous flow of wate r and gas bubbles through the pitot tube,during which time the periodic opening of the water-r emova l valve oc cu rs , dependingon the deg ree of submersion of t he r am end of the pitot tube.


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Figure 6 . - Block I1 hydrogen pump/separator .DEVELOPMENT AND FLIGHT-TEST DIF FIC UL TIE S

During the initial syst em development of the PGS, a number of problems occurredthat necessit ated furth er component development. During the earl y flight te st s of thePGS, additional problems were noted that re sulted in the need for improvements inserv ice procedures and for further system refinements.

Co m p o n e n t De ve lo p me nt a n d P ro d u c t i o nDevelopment problems w ere encountered during the tr ansition f rom prototypehardware to production-type components that would meet the mission requirements.Many of thes e problems w ere unique to the fuel-cell operations and required that newtechnology be developed.Electrolyte s eal. - A s a result of cell peripheral electrolyte-seal leakage, a con-siderabl e delay occurred in the ear ly stage of development. After an exhaustive surveyof ma ter ial s, Teflon was selected as the sea l material. The seal had to contain thehighly concentrated potass ium hydroxide, which is very corrosiv e (especially at 400" to500" F). Also, the se al had to contain the pressure and act as an electric al insulator.At elevated temperat ures , form ed Teflon has a tendency to return to its original shape.Also, the surf ace of the seal is slick, which enables it to slip or extrude (cold flow)through two parallel sealing surfaces. The critical temperature at which these

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phenomena occur is close to the operating temp erat ure of the ce ll (500" F). Be-cause no other material met the seal requirements, a decision w a s made to acceptthe relatively sma ll electrical-pe rformance penalty by operating at lower temperatu res(nominally 400" to 425" F). Also, the flat-cell seal design w a s changed to a nL-configuration, as shown in figure 4, and both mating sur fa ce s we re roughened toallow fo r better containment. This action virtually eliminated all seal leakage problems .Cell flooding. - The two half cel ls (e lectrodes) that for m the single-cell assembly(fig. 4) are composed of dual-porosity sinte red nickel fo rmed fr om nickel powder thatis pre ssed into sheets. The liquid-electrolyte-to-gas eactant interface is maintainedwithin the sin tered nickel by means of a controlled 10.5-psi pr ess ure differential be-tween the electrolyte and the reactant compartments. If either the hydrogen or oxygengas pressure is more than 2.5 p si below or 15 ps i above the electrolyte pre ssu re, abreakdown of the liquid/gas int erf ace possib ly would occur.During the preproto type design stag e of the single cells, many electro lyte leaksdeveloped acr os s the electrode interface as a result of p re ss ur e differentials thatcaused flooding, allowing KOH to enter the reactant cavities. The resu lt w a s the fail-ur e of the individual ce ll to maintain an elec trical load. The manufacturing procedure

was changed so that the porosity of the nickel electrodes w as more uniform, thus in-creasing its bubble pressure and decreasing its susceptibility to flooding. Also, a coat-ing of lithium-impregnated nickel oxide was added to the electro lyte s ide to inhibitoxidation; by this method, the configuration w as controlled during operation. Thesechanges constituted modest impro vements , but the fundamental problem of ground-testcel l flooding caused by gas -pr ess ure imbalance remained throughout the program . Thisground-test operational effect w as minimized by the improvement of ground-support-equipment (WE) gas distribution sys tems and operational test procedures and by care-fu l handling.Dendrite formation. - During qualification tes ting , it w as discovered that, aftertwo 400-hour mission duty cycles, the fuel cell shorted out internally during shutdown.

Nickel ions dissolved fro m the oxygen electrode into the electrolyte and formed nickeldendri tes when reduced at the hydrogen electrode. Eventually, the dendrit es bridgedthe sp ace between the hydrogen and oxygen electrodes (fig. 4) and resu lted i n elec tric alshort circuits that were interna l to the affected cell.It w a s determined empirically that the reaction r at e w a s temperat ure and timedependent. Because thi s w a s a well-defined fa ilur e mechanism, the failu re mode wascircumvented by means of operational procedures. A crit erio n was established requir-ing the removal of the fuel cell s fro m the spacecr aft if an equivalent life of 840 hours(based on an operating tem pera tur e of 400" F) accumulated before launch. Fue l-ce lloperation during the buildup and checkout of spacecraft was minimized; as a result, nofue l cell s have been removed fro m spa cec raft because of the equivalent-life operationalrule.Pressure vessels. - Titanium proved to be an excellent pre ssur e-ve ssel mat eria lfo r fuel cells; however, a higher d egre e of quality control was req uire d during the man-ufacturing proces s than was imposed originally. Two titanium pr es su re vess els areused on each fuel cell: one on the GN2-storage tank and one on the GN2-blanket-

pressure-tank assembly. Contamination and inadequate process cont rol caused the

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tanks to have defective welds that resulted in failures. The two components that wereaffected and the welds that were involved are shown in figure 7.

Flange to form ed sheet m e t a ldome (weld I1-Cover

fl ange- lo -cy l inder we ld+wacketLongi tud ina l cy l inder we ldDom e- lo -cy l inder we ld Power -sect ionpressu re vesse llapproximately 20 i n . l o n g ,12- in . d iam eter )

Figure 7. - The GN2-blanket-pressuretank and power-section pr ess ure-vess el assembly components andwelds.

The first indication of a titanium-weldproblem was a GN2-storage-tank girth-weldfailure that occurred during a component-acceptance test. The cause of fa ilure wasoxygen contamination of the weld caused bythe us e of improper procedures for a girth-weld repai r. Because of the sensitivity ofpressure-vessel failure during flight andthe resultant damage that could have beencaused by such a failure, a thorough reviewof al l procedures and pro cesses that hadbeen implemented was made, and all tita-nium welds were inspected for f l aws . Be-cause of the hydriding that occurred in manyof the titanium welds, all of the high-pressure storage vessels were rejected.As a res ult, improved procedures wereimplemented that included the requirementfor the new GN2-storage tanks to be weldedinside a hard chamber (instead of a bag)that had a capability of vacuum purge andinert-gas backfill. The dewpoint was con-trolled to -30" F maximum. A single weldpass was allowed (previously, two passes

were allowed), and no weld re pai rs were permitted. Each tank-was X-rayed, the in-side and outside surfaces were inspected visually, and the welds were dye-penetrant-inspected before and afte r a proof te st of 3000 ps i and a tank helium-leak check. Theweld specificatdon was revis ed to prohibit high-density inclusions. The quality-controlsamples were required to be with the tanks during the weld and stress-relief operations.All the procedures were reflected in the traceability records that were made more spe-cific. When applicable, simil ar procedures were incorporated into the production ofthe GN2-blanket-pressure tanks. The precautions that were taken have resulted in norecurre nce of fa ilur es in fuel-cell pres su re vessels after acceptance tests wereconducted.

Ceramic-insulator leakage. - Each fuel cell has a hydrogen and oxygen reac tant in-let and a hydrogen and oxygen reactan t outlet (fig. 4). The inle ts of the cel ls a r e joinedto form a manifold (that is; the hydrogen inl et s of the 31 cells a r e joined with a commonsupply), as are the outlets, as shown in figure 3. The piglet tubes between each cel l andits manifold are isolated elect rically by means of cer amic i nsulators (figs. 3 and 4) , tomaintain the desired series electr ica l connection.During the acceptance test of one fuel cell, excessive hydrogen w as detected inthe GN2 blanket, indicative of leaking ceramic insulators in hydrogen piglets . This

indication was confirmed by means of additional te st s. It was shown that the manufac-turer had produced one lot of these in sula tors for which a high-temperature bake cycle

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and assembly leak-check procedures were omitted. The fuel cells that had insulatorsfrom this lo t were recalled; the insulators were replaced with properly processedones. All insulator reco rds were reviewed for quality control to ensure that necessarytests and procedures had not been omitted. More stringent quality-control procedureswere instituted, and the ceramic -insula tor leakage problem did not re cu r.

GroundTestsGround tests indicated that the PGS could operate while the fuel cell s were inte-grated with the spacecraft sys tems during extrerne environmental conditions. Duringthese tests, incompatibilities were noted in the reliabi lity of the components in theintegrated systems; definition and correction were r equired .Accumulator. - An accumulator (fig. 8) is provided as pa rt of the water-glycolcoolant system to maintain a constant coolant pre ssu re, rega rdl es s of the volumetricchanges associated with coolant tempera ture variations. This pr es su re control is ac-complished by imposing a regulated nitrogen-blanket pr es su re on the coolant sys tem bythe use of a flexible bladder.

Approximatelyr 4 i n . 1i

Bladder-Ni t rogen-

Water -ql ycol

I

Figure 8. - Coolant accumulatorassembly.

nately

During early system tests, it w asdetermined that the accumulator siz e w a snot sufficient to function as a pressure-control device for the total temperatur erange of the fue l cell . The problem w asnoted during boilerpla te 14 tes ts, at whichtime the coolant pr ess ure increas ed be-cause thermal expansion of the water-glycolextended the accumulator bladder to itslimit. A larg er accumulator w a s added toproduction fuel cell s, and the problem didnot recur.

Cell separation. - The electrolyte,80 percent KOH, is a porous solid at am-bient tempera ture. Therefore, sm all quan-titie s of reactan t gases can permea te theelectrolyte as it drie s and hardens duringshutdown of the fuel cell. The ea rly shut-down depressurizatio n procedure w as ac-complished by opening the reactant-gaspurge valves and rapidly reducing the cellpr ess ure . When the cells are rapidly de-pressurized, the forces exerted by the ex-pansion of the trapped gases can break thebond between the electrode and the solidifiedDotassium hvdroxide. Thi s process is calledcold popping. On re st ar t of a cold-popped ckll, the trapped reactant ga s fo rm s a bubblebetween the electro lyte and the electrode. It manifests itself a s a reduction in the ac -tive electrode ar ea , with a resultant los s of performance.

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The occu rre nce of cold popping was virtually eliminated by careful adherence toa controlled, slow depress uriza tion of the cell reac tant gas es, which allowed the re -actant trapped in the solidifying electrolyte t o diffuse out.Water-glycol pump. - Several problems were associate d with the coolant pump(fig. 9), including leakage and failure to st ar t. The leakage wa s noted fir st during test s

Carbon en d platesQ u i l l A sealshafl

Figure 9. - Block I1 water-glycol pump.

on spa cec raf t 012. The leakage was causedby a damaged seal surf ace under the motor-to-housing O-ring se al . Although initiallythe damaged area under the se al was notlarge enough to cause leakage, cor ros ion ofthe damaged area outside the seal wasaugmented by the air environment. Thisaugmentation caused an enlargement of thedamaged area and the growth of a largecrys tal under the O-ring seal, allowingleakage. To prevent recurrence of theproblem, all water-glycol pump housings(3561'6 aluminum) wer e ele ctr ole ss nickelplated to re si st damage or corrosion. Norec urr enc e of thi s problem was noted.

During te st s on spacecraft 2TV- 1, hewater-glycol pumps failed to start afterbeing dormant for se ve ra l weeks. By in-vestigation, it was determined that thepump was sticking. The sticking wascaused by deposi tion of nickel phosphate inthe pump gea rs , which was caused by thepresence of chlo rides in the coolant. Thewater-glycol mixture that was used for servicing the 2TV-1 radia tors did not requ ire achloride analysis. A sam ple of the water-glycol m ixtu re that was used was analyzed;the chlorine content w as 37 pa rt s pe r million. Concentrations of grea te r than 25 to30 par ts per million are considered to be intolerable fro m a corro sion standpoint; con-sequently, the chlorine content was limited during servicing to 5 par ts p er millionmaximum.

During the normal acceptance tes ts of new hardware, seve ra l of the water-glycolpumps tended to stick during the firs t st ar t. By investigation, it was shown that, duringa final flush and dryout procedure before sto rage, a residue was left on the shaft andthat the shaf t could not rot ate because the pump ha s a low (4 inch-ounce) starting torque.After the water-glycol pumps were started, the resi due was dissolved and no furth erproblems o ccurr ed. A new rin se and dryout procedure that eliminated furt her problemsof t his type was incorporated.

Hydrogen-vent port. - Two reactan t purge port s, one fo r hydrogen and one for oxy-gen, a r e provided on each fuel cell to allow the purging of i mpur itie s (nonreactant ga ses )that may accumulate in internal cell reactant cavities. Under ext reme therma l condi-tions, during the spacecraft 008 tests on t he original spacecraft design, the water vaporcondensed and froz e at the purge-port opening, preventing furth er hydrogen purging.Two heate rs w ere added to subsequent flight vehicles; thes e hea ters are connected11


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electrical ly in para llel for redundancy. Each hea ter has two elements that oper ate at2 watts pe r element. The hea ter s are activated 20 minutes before a fuel-cell hydrogenpurge and a re turned off 10 minutes a fte r purge termination.Water relief valve. - The fuel cell does not have a water relief valve as such;however, the environmental control sy stem (ECS) has such a valve that in terf aces withthe fuel cel l system. Through this valve, water is removed that is generated by thefuel cell when both ECS water tanks are full and the water pres sure exceeds 45 psia.

Fail ure of t h i s valve to relieve and vent water causes the fuel cell to abs orb an exces-sive amount of water, thus requiring in crea sed volume and resulting in cell flooding andsubsequent failure .Certain ground conditions of the ECS could cause back pres su re s in exces s of45 psia that would be applied to the fuel-cell water supply system as a result of thewater head pr es su re caused by the one-g for ce of the earth. This problem was avoidedby veri fication of the position of the ECS tank-valve switches (thr ee) and by the place-ment of a redline maximum value (12 to 15 pounds) on the amount of wate r in each tank.This problem has not occurre d.Hydrogen pump/separator . The fuel -cel l hydrogen pump/separator (fig. 6)

se rv es two functions : it circulat es moist hydrogen from the fuel-cell stack through thecondenser, regenerator, and inline heater and back to the stack; it removes the con-densed water fro m the hydrogen loop and discharges this water to the spacecr aft water-collect ion tanks of the ECS. The hydrogen gas that is pumped through the motoreliminate s the need for shaft seals in the hydrogen pump and cools the motor. Becausethe hydrogen is saturated with water vapor, many electric al problems were caused un-ti l a sati sfac tory waterproofing epoxy insulation w a s found and a sati sfac tory method ofapplication of the epoxy w a s developed.FI ght Exper iences

Anomalies occ urred on the e arl y Apollo short-duration flights in both Block I andBlock 11 fuel cells. During la te r flights of longer duration, the effects of ze ro g, cou-pled with the thermal cycling, were a problem in the operation of the fuel cell s. Theanomalies were corrected, and operational procedu res were incorporated that resultedin satisfactory fuel-cell performance.craft 011 ( he fi rs t use of fuel ce ll s on an Apollo flight), which had two operable fuelcells (numbers 1 and 3). Fuel cell 2 had been rendered inoperative before launch be-cau se of a checkout-equipment-wiring e r ro r . During the lat ter portion of the flight,the condenser-exit temp eratures of fuel cell s 1 and 3 were out of regulation (high), al-though t he overall performance w a s unaffected. It was concluded that the cooling capac-ity of the secondary coolant loop w a s reduced, and the most likely cause w a s attributedto inadequate radiator servicing that re sult ed ir, coolant-pump cavitation. To preventrecu rren ce of t h e problem, a coolant-system compressibility test w a s designed to iden-tify entrapped air in the PGS radi ator s to verify adequate syst em servicing. A maxi-mum of 8 cubic centimeters of air at a tem perature of 120" F and at a normal operatingpre ssu re of 54 t 2 psia w a s allowed. The compress ibility checks were performedtwice on all vehicles after the occu rren ce of the problem on spac ecra ft 011: once just

Block I flight anomalies. - The only Block I fuel-cell anomaly w a s on space-

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afte r coolant-system servicing and again as late as possible before a launch. The ap-plication of these procedures virtually eliminated the possibility that preflight leaks o rga s bubbles of significant volume would occur in the fuel-cell secondary coolant sys tem.As a result, this problem was eliminated.Block I1 flight anomalies. - When the CSM was maneuvered into lunar orb it, ex-tr em e operational conditions were encountered as a resu lt of exposure of the fuel cel lsto transient thermal conditions.Condenser-exit tempera ture: Thecondenser-exit temperature (TCE) sensor

is a functional pa rt of the secondary water-glycol coolant loop (fig. lo), the output ofwhich is used to maintain a temperatureequilibrium in the wat er condenser of thehydrogen system. This allows a constantwater-vapor pressure in the inlet-hydrogenmanifold throughout the operating powerrange of the fuel ce ll, which, in turn, con-tr ol s the water content of the electrolyte.

Condenser-exit temperature anomalieswere noted on all flights except the Apollo 8mission. The anomalies were in two cate-gori es. The fi rs t anomaly was evident onthe Apollo 7 and 9 missions and was charac-terized by TCE excursions to higher-than-normal temperatures, which were indicativeof res tr ic ted secondary-coolant bypass-valve travel. By analysi s of the data , itwas shown that the res tric tion w as causedby coolant-loop contamination on bothflights. The coolant sys tem had a historyof chronic contamination. The combinationof the coolant, coolant -corros ion inhibitors,and aluminum Dlumbing caused the forma-

c

From To in te r fa c eradiator radiatorsystem system

Figu re 10.- Schematic diagram of fuelcell system.

tion of a gelatinous product after dormant-stand periods of m ore than 3 months duringwhich no loop circulat ion occurred. The gelatinous product, re leased in a zero-g en-vironment during flight, res tric ted the fluid travel in c ertain crit ical control valvesand caused mission-constraint problems. The servic ing procedures were revi sed fo rall subsequent spacecraft that had not already been serviced. Thereafter, the coolantsyste ms we re serviced with water-glycol at the NASA John F. Kennedy Space Cente rrather than at the spacecraft manufacturing facility to reduce the dormant-stand time.Also, the coolant-sampling schedules were revised to include the requirement for morefrequent sampling of the coolant loops. If any samples were questionable, the coolantloops were flushed with fr es h water-glycol while the radi ator panels were vibrated man-ually in an effort to shake adherent contamination loose. The flush and vibration opera-tions apparently minimized the contamination because this type of anomaly was notedin only one of the fuel cell s on the remaining Apollo flights . Spacecraft 107 (Apollo 11)and the subsequent spacecraft were retrofitted with fuel cell s that had a modifiedBlock I secondary-coolant bypass valve in place of the Block 11valve. By means oftesting, it had been shown that the Block I retrofit valve was far less susceptible to


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contamination than w a s the Block II valve. The TCE excurs ions of this type did notoccur in fuel cell s that were installed in spa cecraf t subsequent to that used on theApollo 9 mission.On the Apollo 10 mission , 18" F amplitude (peak-to-peak) TCE oscillations at afrequency of approximately 2 cycles per minute were noted in fuel ce ll 2 during lunarorbital f l igh t (fig. 11). Fuel cell 2 wa soperating at a load that w a s higher thannormal at the time because of the pumpcircuit problem i n fuel cell 1 that causedthe temporary electr ica l isolation of thatfuel cell. The overall fuel-cell perform-ance w a s unaffected, but the caution-and-warning al ar m for low TCE was trippedrepeatedly and had to be re se t manuallyevery 5 minutes, creat ing a nuisance fo rthe crewmembers. The oscillations tendedto damp out at radiator exit tempera turesof 110" to 120" F and recommence at tem-

pera ture s of 60" to 80" F. Five separateinstances of the oscillations occurr ed.After an extensive examination of the flightdata, the presence of a periodic disturbancew a s noted in the TCE and w as characterizedby a 1.5" to 2.0" F drop over a 2-secondperiod with a n approximate 10-second re -covery t ime to the original tempera ture(fig. 12) . Periodic disturbances wer epresent on this fuel cel l throughout theApollo 10 flight. Typically, the dis turb-ances were 8 to 10 minutes apar t and in-crease d to a frequency of 4 to 5 minutes atloads above 30 to 35 ampe res on the affectedfuel cell and served as the trigger mecha-ni smfo r the oscilla tions. The other two fuelce ll s were unaffected. By the use of com-puterized analytical stud ies and ground testsat the NASA Manned Spacecraft Center(MSC) Thermochemical Tes t Area (TTA),it w a s shown that TCE oscillation wa s afunction of load and radiator temperatureand that induced oscillations could alwaysbe damped out by reducing the fuel-cellload. Both the analysis and test programconfirmed that the oscillations were notdivergent and caused no damage to the sys-tem. The only concern w as the effect ofthe repeated caution-and-warning al ar ms onthe crewmembers.

14

1701 I I I I 1 116616 2

?< 15 8L

0,LE 154t

1%14613 3 MOO I 3 3 30 20 13 3 3040 13 3 31 00 13 3 31 20 13 3 31 40 133 3 2 0 0

T i m e hr m in sec

Figure 11.- Typical fuel cel l 2condenser-exit tempera tureoscillations during lunarorbit.

162

iij 160CLsI-!

158156030 20 30 40 50

Time, secFigure 12 . - Typical fue l cell 2 TCE

disturbance.


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An extensive review of all previous flight data resu lted in proof that this type ofdisturbance occur red in at least one fuel cell per spacecraft on all Block II flights. Ex-tens ive investigation of the cause of the TCE disturbance was indicative that water slug-ging out of the pr imary (hydrogen) loop condenser in a zero-g environment w a s themost probable cause . Because of the absence of a gravity field, the water probably didnot leave the condenser in a uniform manner, but tended to accumulate at the condenser-exit plenum until some crit ical level w a s reached, at which time the subcooled waterwas released in the form of a fair ly large globule. Presumably, the subcooled watercontacted the TCE probe and the secondary-loop bypass-valve-control sensor , causingthe cont rol disturbance and result ing valve oscillations under the high loads and lowradiator tempera tures that occurred in lunar orbit on the Apollo 10 mission.

Provisions were made in flight procedures to unload a fuel ce ll that exhibited anoscilla ting TCE. Through the Apollo 16 mission, however, no other fuel cell has dem-onstrated this phenomenon, probably because all th ree fuel cell s have s hared loadsthroughout all other missions.Hydrogen pump: During the Apollo 10 miss ion, a sho rt circuit in the alternating-cu rren t pump package of f ue l cell 1 caused the associated circuit breaker to tr ip open.

The breaker would not reset; thus, a permanent sho rt was indicated. Fuel cell 1 wasremoved immediately from the bus because both the hydrogen and coolant pumps wereinoperat ive. This fuel cel l was maintained in an operative standby mode by placing iton the bus when the fuel-cell-stack temperature reached 370" F and open circui ting itwhen the stack temper ature reached 420" F. After the hydrogen-pump failure in fuelcell 1, the maintenance of this fuel cell in the operative standby mode was accomplishedby periodic loading, which caused the concentration of water in the elect roly te to in-cr ea se eventually to approximately 33 percent. At this concentration, the fuel ce ll soonwould have flooded i f water could not have been removed. A continuous hydrogen purgewas initiated to reduce the water concentration af ter approximately 167 hours of theflight. After 3 hours, at which time the fuel cell w a s sufficiently dry, the purge valvewas closed and the hydrogen vent-line hea ter was turned off.Circuit anal ysis , inverter testing, and review were indicative that the failureprobably w a s a phase-to-phase sho rt circuit in the hydrogen-pump sta tor windings,probably caused by insulation breakdown. Because the motor windings are exposed tohot, wet hydrogen, the limited life of the motors is attributed to the basic design. Inendurance test ing, the minimum length of t ime before a motor s tato r failed was1000 hours; the maximum time before a failure was 3960 hours. Except fo r a majorredes ign of the hydrogen pump, no procedural o r design changes could be identifiedthat would enhance the reliabil ity of the pump. Consequently, the use of the pump wasretained throughout the flight program without hardware changes. This type of f ail uredid not occur in any flight fuel cell other than i n the one on the Apollo 10 mission.Hydrogen purge flow and pr es su re excursion: After the extended purge of theflooded fuel c el l on the Apollo 10 spacecraft, the hydrogen-supply flow failed to shutoff completely. Normally, hydrogen flow should have decreased to ze ro in less than1 minute; however, the flow ra ted ecreased very slowly. The purge valve was reopened,and the flow rate increased to the upper limit, which was indicative that the purge valvewas functioning. The valve was closed again, but the flow decr ease was s til l very slow.After approximately 30 minutes, as the flow rate approached zero, the regulated

15


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hydrogen pre ssu re for fuel cel l 1 began to incr ease and reached a maximum pres sur eof 72 psia before it decreased slowly to the normal pre ssu re of 62 psia (fig. 13).A test program was conducted at the

:so vendor facility to te st the flight hardwareunder simulated environmental conditionsf m,n at the time of failure . During the te st s,g , the regulator temperatur e reached -23" FL 24 in 5 minutes and reached -100" F in 15 min-e 16 U t e s during hydrogen purges . At a tempera-I m tu re below -10" F, both the regulator supply- and relief (vent) valves, al so a part of theregulator assembly, leaked because these al s stiffened and did not conform to the

a

,n

fw

" 01 ',; 4; 'i; sb 'i7dOQ im 18 170x: in 42 i m K 171 06Time. hr min

sea ts. Prope r sealing was rest ored when-10" F.Figure 13*-Hydrogen rate and the regulator temperature increased topres sure af ter purge.

The test resu lt s were indicative that the extended hydrogen purge in flight withouta hydrogen-preheat capability created low temperatures in the regulator; the consequentleakage of both regulator s ea ts was the reason for the continued flow. Both the purgeline and the vent line a r e vented overboard through a common header, and the header isprotected from freezing by the heat ers that were discussed previously. In this instance,the hea ter s wer e turned off when the purge valve was closed ; thus , the continued ventthrough the regulator caused cooling and subsequent freezing in the overboard vent line.Normally, the regulator controls overpr essur e by venting on the downstream side.Consequently, blockage of the vent line caused an inc rea se i n regulated hydrogen pres-sure. Subsequently, tests at the TTA resulted in proof that subcooled hydrogen can bepurged safely and continuously fo r a t l eas t 9.5 hours if the reactant-preheat capabilityis maintained.

A change was incorporated in the Apollo Operations Handbook requ iring that thevent-line heater be kept on for 10 minutes after the termination of a hydrogen purge .Extended hydrogen purging was not req uired a fte r the Apollo 10 mission.Hydrogen gas in potable wate r : Gaseous hydrogen (GH2)n the spacecraft potable

water was a problem on all manned Apollo flights. The operation of the fue l-cell pump/separ ato r caus es expulsion of water that is free of gas except for the na tural absorptionof GH2 n water at 60 ps i and 160" F. When the environment of the water is changed tothat of the potable-water tank (25psi, 80" F), 3 Xof gaseous hydrogen is expelled from each pound of water that is collected fro m thefuel cell. The theoreti cal GH liberation value was shown and veri fied in controlledtests at the MSC.

pound (10.3 cubic centimeters)

2The hydrogen that was entrained in the potable water posed no hazards t o thecrewmen or missions but did represent a nuisance when the water was used by causinguneven water-flow slugging and some physical discomfort to the crewmen from the gasaccumulation in their digestive systems. A hydrogen-gas sep arato r was added to theECS potable water syst em to reduce the occurrence of this problem.

16


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C O N C L U D I N G REMARKS

1. System selection/design cr it er ia should include susceptibility to damage asa result of operational e r ro r s . To supplement these cri te ri a, ground-support equip-ment should be designed so that, when a hardware fai lure occ urs, the mode of failurewill preclude damage to the spacecra ft hardware. Also, thorough training prog ramson fuel-ce ll operation should be conducted for engineers and technicians who per-fo rm field operations in this ar ea to ensur e adequate protection in ca se of off-designconditions.

During the cou rse of the Apollo Prog ram, the fuel cell w a s proven to be a rugged,reliable, and vers ati le electrical-power-generation device. The fuel cel l operatedsatisfactorily during spacecraft launch/boost vibration, in zero g, and in a space/vacuum environment. The fue l cel l met al l elect rical demands that were imposed on itfrom low (t rans lunar coast) to high (Apollo 10 two fuel-cell return) power levels.

The experience gained from the ground- and flight-test phases of the Apollo fuel-ce ll program was indicative of cer tain hardware design and operational sensitiviti esthat, although not considered major problems , required a great deal of t ime and effortto correct or work around. Some of the problems were unique to the fuel cell ; otherproblems w ere caused by integration with other spacecraft syst ems,Operational er ro r s caused the costly failure of seve ral fuel cells during earl yservicing and checkout operat ions in the Apollo Pr ogram. The high frequency of thisoccur rence was caused by the complexity of the fuel cell s and their sensitivity to off-

design conditions. After the fuel ce ll w a s operating and was on internal spacecra ftreactants, the system operated reliably.Contamination, par ticu lar ly in circulating-f luid sys tems (for which hardwaretolerances a r e critical), wa s a serious problem for spacecraft subsystems, as evi-denced by the chronic contamination of the fuel-cell coolant loops caused by the forma-tion of a gelatinous product after dormant-stand periods during which no loop circulationoccur red. These problems were minimized before subsequent flights by means ofground flush and vibration procedures that (presumably) removed most of the gelatinousproduct from the coolant loops.The fuel-cell condenser-exit-temperature disturbance that occurred during the

Block II flights is believed to have been caused by a condensation phenomenon in thezero-g environment. Enough evidence was produced during ground tests to verify thisas the most likely cause of the disturbance. The apparent instability of the fuel cellunder these conditions was shown by a mathematical model that re pres ented the fuel-ce ll module.The redundancy philosophy that was instituted by the fuel c el l sys tem des ignersresulted in syst em and mission flexibility. The value of redundant powerplants wasevident during the Block 11flights , during which a fuel c el l was shut down and had noeffect on the accomplishment of the mission objectives.To preclude s imi lar problems on future spacecraft and to provide a more reliable

fuel-cell power syst em, the following recommendations ar e made.


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2. System and spacecr aft interface s should be defined carefully, and the defini-tion should be adhered to during subsys tem design and integration phases .3 . The compatibility of circula ting fluids with syste m hardware must be verif iedthoroughly by ground testing.4. A l l fluid loops should have fi lte rs installed upst ream of all critical

components.5 . Critica l automatic control devices , such as the transduce r used to control thewater-vapor pre ss iir e in the hydrogen loop, should be used in a manner that will pre-clude control-sensor operation in a two-phase-fluid medium.

Manned Spacecraft CenterNa t i ona l Aeronautics and Space AdministrationHouston, Texas, October 14, 1972914- 11-10-00-72

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apollo experience report power generation system

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