Apollo Experience Report Command and Service Module Environmental Control System

8/8/2019 Apollo Experience Report Command and Service Module Environmental Control System

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APOLLOEXPERIENCE REPORT -C OMMA N D A N D SER V I C EMODULEENVIRONMENTALCONTROL SYSTEMby Frunk H . Sumonski, Jr., und Elton M . TzcckerM a n n e d Spucecrd t CenterHouston, Texas 77058

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. ,.. r . I I , ./ .N A T I O N A L A E R O N A U T I C S A N D S P A 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. M A R C H 1972


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.TECH LIBRARY KAFB, NM

APOLLO EXPERIENCEEPORT March 1972COMMAND AND SERVICE MODULENVIRONMENTAL 6. Performing Organization CodeCONTROL SYSTEMFrank H. Samonski, J r . , andlton M. Tucker, MSC 1 MSC S-2797. Authcds)

-~ ~~~~ ~ ~~~ ~~~ .~8. Performing Organization Report No.

"_ ." . _ ~ " ~ 10 . Work Unit No.9. Performing Organization Name and Address 914-11-10-00-72Manned Spacecraft CenterHouston, Texas 77058 I 11. Contract or Grant No.I 13 . Type of Report and Period Covered

AgencyNameandAddress Technical NoteNational Aeronautics and Space AdministrationWashington, D. C. 20546 I 14. SponsoringAgency CodeI~-~5. SupplementaryNotes ~~The M SC Director waived the use of the International System of Units (SI) for

this Apollo Experience Report, because, in his judgment, use of SI Units would impair the usefulnessof the report or result in excessive cost.~~ - . . -" ~~ ~~~16 . Abstract .~ - ~~~

This pape r pre sen ts a comprehensi ve review of the design philosophy of the Apollo environ-mental control system, and the development history of the total system and of sele cted com-ponents within the system. In partic ular, discussions are presented relative to thedevelopment history and to the problems associated with the equip.ment cooling coldplates,the evaporator and its electronic control system, and the space radiator system used forrejec tion of the spacecraft thermal loads. Apollo flight experience and operational diffi-culties associatedwith the spacecraft water system andhe waste management system aredisc uss ed in detail to provide definition of the proble m an d the corrective action taken henapplicable.

I~ ~ ~~.._ _ _ _ _ _ . _" ~~ ~ ~~ ~ .17. Key Words Suggested b y t h o r k l i 18: Distribution Statement~ ~~- Space Radiator *Waste Management' Evaporator - Environmental Control* Liquid/Gas Separator * Atmosphere Selectionk n e ." . - ." - -~ ~. . . . . ."" ~ . . ~ ~ _ _ ~19 . Security Classif. (of this report) 20 . Security Classif. (of this page)T None 21. NO. of pages 22. Rice*29 $3 00~~~ ~~- "

Fo r sale by the N ationa l Te chnica l Inf ormatio n ervice, Springfield, Virginia 2 2 1 5 1


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A POL L0 EXPER IENCE REPORTC O M M A N D A N D S E R V I C E M O D U L E E N V I R O N M E N T A L

CONTROL SYSTEMBy F r a n k H. S a m o n s k i , J r . , a n d E l t o n M. T u c k e rM a n n e d S p a c e c r a f t C e n t e r

S U M M A R YThe Apollo environmental control system w a s designed and qualified to supportthree crewmen for 1 4 days and to maintain electronic equipment within operating ther-mal boundaries. The system maintains the pressure atmosphere of 100 percent oxygen

and removes trace contaminants and metabolic carbon dioxidey absorption in charcoaland lithium hydroxide beds. Temperature control is provided by heat rejec tion fromradiators and a water evaporator. Oxygen is supplied by the cryogenic storage system,and water is supplied as a byproduct of the fuel cell s. The knowledge gained from ex-tensive ground testing and inflight experimentson the behavior of water in zero gravi tyled to the incorporation of a wick-type porous-plate condensate separator.The two hardware items requiring the most extensive development were thewaterevaporator and the radia tor. During the Apollo Program, continuous refinements havebeen required in the construction, material selection, and quality control of the evap-orator and its control system. The wide range of the maximum and minimum heat loadsled to the use of a selective stagnation radiator designed to employ the viscosity char-

ac te ri st ic s of the coolant fluid (ethylene glycol and water). The other major problemsexper ience d w ere in mate rials selec tion o reduce corrosion (particularly in the coolantsystem), materials selection for fabrication of porous plates and heat exchangers, andmater ial and proces s refin ement s to eliminate weld crazing.

I NTRODUCTI ONDuring an Apollo mission, the environmental control system (ECS) of the com-mand and service module (CSM) provides life support for the flight crew and thermalcontrol for the vehicle electronics systems. During the major portion of the mission,the CSM serves as the living quarters forall three crewmen. The components of theECS are located in both the command module (CM) and the service module (SM).This report presents a discussion of the functional and physical ECS design re -quirements based on the integration of the ECS with the CSM and the mission param-eters. A descr iption of systems opera tions and of some significant problemsencountered during development and flight testingf the hardware is al so included.


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DES I GN C O N S DERAT l ONSR e q u i r e m e n t s

The initial ECS desig n requirem ents w ere in the fo rm of general guidelines,which later were refined for contractual purposes. A 5.0-psia pure oxygen atmospherwas requ ired in t he pres suriz ed CM, and a shirtsleeve environment for the crewmember s w a s established as the normal mode of oper ation . Pressur e suits were to be wornonly during critical mission phases, such as launch and entry. In the event of anemergency cabin depressurization, the system should maintain the cabin-pressurelevel within acceptable limits for a time period sufficient for he crewmem bers to dontheir pressu re suits . To provide this capability, statistical data were used to estab-lish a design criterion for maintenance of cabin pressure above 3.5 p sia for 5 min uteaf ter a 1/2-inch-dia meter puncture in the pressure vessel. Should lo ss of cabin pres-surization occur, life support in the pressure-suit mode would be provided for a suffi-cient duration to permit the safe ret urn of the crewm emb ers.

The design requirements limited the carbon dioxide partial pressure to 7.6 t o r rand specified that the carbon dioxide removal technique w a s to be chemical absorptionby lithium hydroxide. In the earl y design stages, a regenerable molecular sieve hadbeen considered but was rejected for the more reliable lithium hydroxide absorptionprocess. The CM gas temp eratu re was to be maintained at 75" * 5" F, except duringent ry when 100" F maximum was permissible. The cabin relative humidity was lim-ited to 40 to 70 percent. In addition to the atmospheric requiremen ts, the ECS wa s toprovide thermal control for electrical equipment, and critical equipmen t was not to bedependent upon the cabin atmosphere for cooling.

The conversion of these d esign requiremen ts into operational hardware nvolvedseveral additional factors associatedwith the interface definition of the syst em withthe spacecraft. The more significant considerations were w eight and volume limita-tions, power requirements, and reliability. In addition, the design approach was fre-quently dicta ted by the environments in which the equipment would operate.

W e i g h t a n d V o l u m eThe free gas volu me of the CM is approximately 320 cubic feet as comparedwith the 80-cubic-foot volume of the Gemini spacecra ft. Judicious management ofsystem volumes in the design phaseof the CM was necessary to provide the free vol-ume required for the crew functions unique to theApollo missions, such as crewtransfer to the lunar module (LM) in a pressurized environment. The volume occupiedby the ECS, including stored expendab les such as the lithium hydroxide canisters, isapproximately 20 cubic feet. Negligible system volume growth has occurred duringthe program.Equipment weights have proved to be more difficult to predict than the equipmentvolume. Ekperience with the earlier manned spacecraft has demonstrated an increasein weight with time, as estimated nu mbers we re replaced by weights calculated fromdetail drawings which, in turn, were revised as actual hardware weights became available . Three poin ts in the weight history of the CSM ECS are presented in table I.

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TABLE I. - ENVIRONMENTAL CONTROL SYSTEM WEIGHT S U M M A R Y

. " -.I DateIi..anuary 1965July 1966July 1969 . .-~ "

Weight of ECSCM, lb Total, IbM, lb

44 252 8553

A significant portion of the weight- increase trend is associat ed with the designchanges, which usually resulted from difficulties encountered during the developmentof the ECS. Another source of weight increase is the modification of ECS requirementsas a resu lt of experience with the system, the spacecraft, and the mission itself. Also,system requirements may shift as more refined definitions of the int erf ace s with othersystems become available. When configuration control to the component level is estab-lished, then the weight increases are minimized and generally occur, if at all, as theresu lt of failures during the qualifica tion test program.Pow e r

The requirement for electrical power is one basic element which is optimized inthe trade-off studies conducted during the preliminary design of the spacecra ft. TheApollo spacecraft uses three fuel cellsas the primar y sources of power, with batteriesproviding a supplemental source during peak-load periods. A sum mar y of the elec tri-ca l power requiremen ts of the ECS is given in table 11. This load is constant and in-dependent of the mission phase.

TABLE 11. - ENVIRONMENTAL CONTROL SYSTEM POWER SUMMARY

ComponentsSuit compresso rPumpCabin fan (2)ControlsInstrumentation

Total 1972255-

aThe values for ac power are as supplied to the rotating machinery. All ac power(115/200 volts, 3 phases, 400-cycle) for the spacecraft is supplied from three centralstatic inverters which energize two independent buses.

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R e l i a b i l i t yReliability was another prime consideration in the design of the ECS. Therefothe keynote of the ECS design is redundancy. The system must operate continuouslythroughout the mission, and although inflight maintenance was considered, it did notappear practical. Instead, redundant features were used whenever possible, and inother situations, completely independent manual-override capability was provided.

Test experience inall pha ses of the program , from design-fea sibility testing througflight performance evaluation, has been used to increase the confidence in the reliabiity of the system. In addition, reliability data collected from comparable hardwareand design techniques used in aircraft and other spacecraft systems contribute to heconfidence level. These data ar e us ed in conjunction with reliability logic diagramsand failure modes and effects analyses o gain further confidence that the system de-sign meets its reliability goals.E n v i r o n m e n t

The environment in which the system must operate frequently dictates the de-sign approach. The 5-psia pure oxygen atmosphere eliminates consideration of manystandard parts and materials because of the. haza rd of combustion. Other versatilematerials are rejected because they outgas at reduced pressures. This outgassingresu lts in toxic-co ntaminant buildup in the sealed environmen t of the C M cabin. (Ex-cept fo r the 0.2-lb/hr maximum external leakage, the CM is completely sealed. ) Thus e of the ethylene glycol and water coolant presented a toxicological problem. A l-though the effects of acute ingest ion were known, the toxicity resulting from continuouinhalation of gaseous o r aerosolized glycol resulting from a leak was unknown. Toestablish acceptable levels anda mean s for detection of glycol in the atmosphere, aseries of tests w a s perf ormed on eight different mammalian species, ranging from thmouse to the chimpanzee. Based upon the result s of the animal testi ng, human volunteers were exposed to glycol aerosols and vapors for periods up to4 days. It w a sfound that levels whichdid not produce subjective responses of a sweet odor orpharyngeal irrit ation would not produce physiological changes in humans in 14 days.Once the sense threshold was passed, however, prolonged exposure was irritating ansubjectively intolerable. In the event the threshold were to be exceeded, it would benecessary to isolate the crewmembers from the atmospherey use of the emergencyoxygen masks o r the intravehicular space suits.

The vibration and acceleration loads encountered during launch and entry conute to the complexity of the design. By thems elves, these dynamic force s usually arenot difficult to overcome in the design of the individual components, but frequently, thefor ce s are amp lif ied y the manner of equipment packaging o r by secondary supportinstructures in the spacecraft.E N V l R O N M E N T A LC ON TR O L S Y S T E M D E S C R I P T I O N

The following description of the ECS is presented to provide orientation for thmore specific developmental problems to be discussed later. The schematic diagramof the ECS (fig. 1) may be convenient ly divided into the oxygen, water, coolant,pressure-s uit, and cabin circuits. For orientation within the spacecraft, refer to fig-u r e s 2, 3, and 4 which show the spacecra ft insta llat ion of ECS equipment.4


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The pri mary oxygen source cons ists of two supercritical cryogenic tanks locatedin the SM. These tanks also supply the oxygen r e q i r em en t s of the fuel cells and aregenerally considered as.pa rt of the electrical power system. The two tanks contain atotal of 640 pounds of oxygen, and the design speci fica tion allocates 172.6 pounds of thisamount to the ECS (table In). In comparison, the actual oxygen allocation to the ECSfor the Apollo 11 mission was 72.4 pounds for planned use, 10.4 pounds for LM support,and 15.6 pounds fo r contingency use. The reduced consumption during the Apollo 11mission resulted because the mission duration (196 hours) w a s less than that of thespecification mission (336hours), and the cabin leakage and crew metabolic require-ment values were lower than design specification requirements (table V).

TABLE III. ENVIRONMENTAL CONTROL SYSTEMOXYGEN- PECIFICATION REQUIREMENTS

AllocationC r e w consumption (1.8 b/man-day for 14 days)Cabin leakage (0.2 b/hr for 336 hours)Cabin repressurizations

Oxygen, lb75.667.211.7

One CM puncture

172.614. 5M support3.6

Total

TABLE IV. - ACTUAL ENVIRONMENTAL CONTROL SYSTEM

Apollo mission7891011

OXYGEN CONSUMPTION

I IDuration, hr Quantity of oxygenconsumed, b

259.7 102146.5240.5190.0196.0

51997182

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Oxygen delivery to the ECS at 900 f 35 psia is limited to 4 . 5 lb/hr from eachtank by the combination flow restrictors and heat exchangers, which ensure a minimutemperature of 0" F at the maximum flow rate. Short-duration demand s in excess ofthis capability are satisf ied by a surge tank (3.7 pounds at 900 psia) and a repressuri-zation pack (3 pounds at 900 psia) . Eithe r or both supplies may be isolated to preservthe oxygen for en try or other use. T he ma in oxygen pressu re reg ulato r, which supplithe components and functionsas indicated in table V, reduces the circuit pressure to100 5 10 psig.

T A B L E V. - ENVIR ONMENTAL C ONTR OL SYSTEM OXYGEN-C OMPONENT SUMMAR Y

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The primary source of w ate r forthe ECS is the fuel cells, which produceapproximately 0.77 lb/kWh as a byproductof fuel-cell operation. The water storageprovisions consi st of a 36-pound-capacitypotable water tank (fig. 5), and a 56-pound-capacity waste-water ank. Excess mois-ture in the cabinor suit circuit gas isremoved by the water separator in thesuit heat exchanger and s t ransferred bythe cyclic accumulator to the waste-watertank for subsequent useas an expendablecoolant. The effluent fro m the fue l ce llsis directed to the potable-water tank andis used for drink ing and ood reconstitu-tion. Periodic injection of chlo rine by thecrewmembers maintains bacteria controlin the potable-water system. When thepotable-water tank is full, the water cir-cuit automatically diverts the fuel-celloutput to the waste ank by elevating thewater-system pressure from 25 to 30 psia.When both tanks are full, the water-systempressure is increased to 40 psia, and the

Figure 5. - Potable water tank disassem-bled with the bladder, support frame,and quantity transducer.fuel-cell effluent is dumped directly overboard. E xcess wate r may also be dumpedmanually, and this capability has been used in all missions. This manual operationwas chosen to prec lude inter ference with photography, sightings with the guidance andnavigation equipment, and trajectory determination.

The coolant system consists of a pr imar y loop, which is operated continuously,and a secondary loop, which ser ve s as a backup system. The primary loop uses acentrifugal pump to circulate 200 lb/hr of coolant (ethylene glycol and wate r) throughthe heat-absorption and heat-rejection equipment in the CSM. If the coolant returningfrom the space radiator is less than 45" F, it is mixed with fluid from the CM the rma lload, which has bypassed the radiato r, to obtain a mixed-coolan t temperature of 45" F.Under mission conditions when the space radiator cannot reject the tota l load, no by-pass occurs; instead, the glycol evaporator cools the 200-lb/hr flow to 41.5' F byevaporating water at a contr olled pre ssure of approximately 0.1 psia. The coolantflow leaving the evaporator is divided into a 35-lb/hr flow directed to the inertialmeasurement unit (IMU) of the guidance and navigation equipment, and a 165-lb/hrflow is routed to the suit heat exchanger through the drinking-water chiller shown infigure 6. The suit heat exchanger, shown in figure 7, provides the humidity controlfor th e CM. The coolant leaving the suit heat exchanger enters the cabin heat ex-changer and abso rbs heat from the CM lighting, the electronic equipment not mountedon coldplates, the environmenta l loads, and the crewmembers in the shirtsleeve mode.The effluent coolant from the guidance and navigation equipment mixes withhat fromthe cabin heat exchanger, and the 200-lb/hr flow is directed through a series-parallelarrangem ent of 22 coldplates, which absorb the major portion of the thermal load. Theheat from the coldplate network may be diverted to the cabin heat exchanger throughthe cabin-temperature control valve for heating the cabin, when required. The fluidleaving the cabin-temperature control valve enters thepump, and the flow is directed

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Figure 6. - Drinking-waterhiller. Figure 7.- Suit heatxchanger.to the space radiator. A secondary coolant loop is provided as a backup for the pri-ma ry loop and may be operated at the disc retio n of the cre wme mbe rs. Both loopsprovide cooling for the suit and cabin atmospheres and fo,r the electronic equipment.The secondarv loor, does not have cabin-heat ing capabili ty, nor does it provide coolingto the guidance and navigation equipment.

The pressure-suit circuit controlsthe lev els of carbon dioxide, odor, andhumidity and can provide a habitable en-vironment for the crewmembers if cabinpressurization is los t. When the crew -members are in the pressure-suitmode,they are isolated from the cabin. Theventilating gas flow leaving the pressuresuits passes through a deb ris trap , shownin figure 8, which remove s parti cleslarger than 0.04 inch. Suit circuit flow isaccomplished by one of two centrifugal-flow comp res sor s which deliver 55 lb/hrof suit-circuit gas (35 cu ft/min) at apressure rise of 10.0 inc hes of wate r withan inl et densi ty of 0.0266 lb/cu ft.

A s the ventilation gas passes throughtwo par all el ele men ts of lithium hydroxideand activated carbon, the carbon dioxideand odor control for the C M is accom-plished. Each element is sized for Figure 8. - Apollo solids trap.12


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1.5 man-days of operation at the design metabolic loads, and the elements are changedby the crew alternately every 1 2 hours. Twenty elements are carried for 8- to 10-daymissions.Theelemehtholder, o r canis-ter, incorporates the necessary checkvalves, diverter valve, and interlockmechanisms which permit the changing ofelements in a depressurized cabin. Thecanister is also designed to preclude in-adverten t depressurization of the suitcircuit.

The gas leaving the carbon dioxidecanister enters the suit-circuit heat ex-changer, where suit-circuit heat loadsare absorbed by the water and glycol. A tthe heat exchanger, the moisture is con-densed, removed by the wicking, andtransferred to the w aste-wat er circuit bypneumatically actuated accumulators(fig. 9) which are cycled every 10 minutesby a timing device. The normal gas exittemperature from the heat exchangeris50" F.

The cool gas is distributed to thethree suit-hose-connector units, which in-corporate a flow-control adjustment leverand a flow-limiting Venturi tube. Whenthe crewmembers are in the shirtsleevemode, the ir por tion of the suit-circuitflow is delivered to the cabin through anorifice in the connector unit which approxi-mates the pr es su re drop of the suit. Thisflow is returned to the suit circuit for car-bon dioxide and humidity removal by thecabin-air-r eturn valve located upstream ofthe suit compressor s. During mannedground testing and during launch, thecabin atmosphere is a mixture of 60 per-cent oxygen and 40 percent nitrogen. Thisis the minimum oxygen concentrationwhich will provide a viable atmospherewith a reduction to 5.0 psia in the cabinpressure. Subsequent to orbit insertion,a bleed flow overboard establishes ademand on the cabin-pressure regulator(fig. 10) and enriches the mixture to sea-level equivalent (an oxygen par tia l pre s-sure of 3.1 psia). The nitrogen content

Figure 9. - Apollo cyclic accumulator .

Figure 10. - Cabin pressure regulator.1 3


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is reduced further by leakage o r LM pressurizations. Technical considerations as-sociated with the select ion of the launch environment are presented in the followingsection.The crewmembers undergo a period of oxygen prebre athing prior to inser tioninto the CM suit ci rcuit , which has been purged to an oxygen level greater than 95 per-cent. This oxygen prebreathing minimizes the possibility of aeroembolism during the

boost phase when cabin pre ssu re is reduced from 14.7 to 6 psia. To prevent nitrogenleaking into the suit circuit, a positive pressure relative to the cabin pressure ismaintained by a 0.5-lb /hr excess flow.Finally, the cabin circuit consists of two axial-flow fans. Each fan has a capac-ity of 86 cu ft/min at 5 psia, which circulates the CM gas through the cabin heat ex-changer. A cabin-pressure relief valve relieves the cabin pressure at a differentialof 6.0 psi during ascent of the sp acec raft and re press urize s the cabi n during des cent,when the am bient pre ssure exceeds cabin pre ssure by approximately 1 psid. Aftersplashdown, a postlanding ventilation system, consisting of an inlet valve and fan andan outlet valve, is activated by the crewmembers to ventilate the cabin until recove ry.In the event of smoke, a toxic gas, o r another harmful atmosphere in the cabinduring the shirtsleeve environment, three oxygen masks are provided. The mask isa modified commercial full-face-type assembly with headstraps to hold it on. Theoxygen is supplied at 100 psi through a flexible hose from the emergencyoxygen andrepressurization unit. The mask has an integral regulator that supplies oxygen ondemand when the crewman inhales.

DEVELOPMENT AND FLIGHT T E S T I N G DIFFICULTIESGround Test and Launch Environment

A s a re su lt of the crit icali ty of the gas environment in the spacecraft duringlaunch, detailed technical considerations involved in the selection of a two-gas atmos-phere fo r the Apollo CM during ground checkout and launch a r e presented. The designanalysis included a comprehensive engineering trade-off study which resulted in severalgeneral conclusions concerning atmospheric selection. The analysis provided thenecessary data required for the final s election f a 60-percent oxygen, 40-percent ni-trogen launch atmosphere.The purpose of the evaluation was to determine the feasibil ity of using a two-gasatm osphere in the CM during ground checkout and launch. Selection of this atmosphererequired the consideration of several factors direc tly affecting crew safe ty: (1) he

physiological acceptability of the cabin atmosphere, (2) the capabili ty of reduc ing thefire hazard, (3 ) the opera tiona l characte ristic s of the spacecra ft hardw are when sub-jected to the atmosphere, and (4) he additional crew procedures required.The crewmembers undergo a lengthy oxygen prebreathing period prior to launch,and subjecting them to even a smal l amount of di luent for a sho rt tim e could destroy thebenefits of this prebreathing period. Therefore, a 100-percent pure oxygen environmentduring prelaunch checkout and launch was physiologically desirable because it elimina-ted any potential dysbarism. The pure oxygen environment was least desirable from a

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fire-hazard standpoint, since material flammability tends to be proportional to atmos-pheric oxygen parti al press ure. Thus , from a fire-hazard viewpoint alone, pure diluentin the cabin atmosphere seemed most desirable.Of course, a pure diluent cabin atmospherewould surround the suited c rewmanwith a nonviable atmosphere during the entire ground checkout and launch period. Eventhough the crew is not dependent on the cabin atmosphere during this time, it wasdeemed desirable to maintaina livable atmosphere to accommodate any unforeseen

emergenci es. These considerations dictated the requirement for an analysis of theoptimum mix ture of oxygen and diluent to be used in the cabin during prelaunch check-out and launch.During the trade-off study, the following requirements basic to all considera-tions were made:1. The fire protection capabili ty of the launch atmosphere must be enhanced bya two-gas mixture , through the use of a n in er t diluent.2. The two-gas atmosphere considered must provide a livable cabin atmosphere.3. If, after orbital insertion, the diluent must be removed to meet physiologicalrequirements, crew operational procedures must be minimized.In accordance with these requirements, the atmosphere-selection study w a sconducted, using ECS hardware compatibility testing, systems performance analysis,weight trade-off studies, and materials compatibility studies. An extensive programw a s organized for analyzing and testing flammability characteristicsf componentsand entire systems withina C M test vehicle.The trade-off study demonstrated that an air cabin environment would provide aviable atmosphere surrounding the crew during the relatively long prelaunch period,but in order to maintain a viable atmosphere during the entire launch phase and to

avoid high purge rates during launch, an oxygen partial pressure higher than thatpresent in air must exist in the cabin a tmosphere prior to launch. The oxygen partialpressure in the cabin could be continuously maintained through launch in a physiologi-cally acceptable range w i t h a 50- to 60-percent oxygen concentra tion. This concentra-tion provides an atmosphere with flammability characteristics near those f a 6-psiapure oxygen environment.Use of the two-gas cabin atmosphere required certa in precautions because thesuit-circuit pressure upstreamof the suit-compre ssor inlet is normally less than thecabin press ure. Suit-cir ,cuit leaks i n that vicinity pr ior to cabin diluent removal wouldtherefore cause the diluent -rich cabin atmosphere to enter the circuit. Since thediluent is not consumed in metabolism, it would continue to build up in the suit circuit.The oxygen partia l pressur e in the suit circuit would correspondingly diminish, re-sulting in potential hypoxia of the suit ed crew or in aero embo lism upon launch. Ableed rat e of approximately 0.80 pound of oxygen per hou r into the sui t cir cui t through-t he manual oxygen-metering valve and vented through the demand pressure regulatorwas used to provide positive suit pressure to prevent diluentuildup.The trade-off study showed that, in view of the definit ion of the cabin oxygenpartial pressure required to support the crew during launch and abort modes and in

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view of the requirements associated with the ire haza rd of va rio us oxygen diluent mix-tures, a two-gas mixture for Apollo ground checkout and launch is both acceptable anddesira ble. Furth ermo re, the analy sis estab lished the feasib ility of using a two-gasmixture with current spacecraft hardware for prelaunch and launch. The most de-sirable composition was ound to be a mix ture of 50 to 60 perc en t oxygen in n itrogen.A i r and m ixtures of helium and oxygen were also considered, but these combinationsprove d to be unde sirabl e. Several methods of removing the diluent, such as cabinleakage, cabin purging, and depressurization techniques, were investigated. The useof cabin purge through existing overboard dump nozzleswas found to be most desirableand has been adopted for eachmanned mission.

ColdplatesThe original design of the Apollo coldplates used parallel passages milled intoa 3/8-inch shee t of 6061 aluminum on approximately 1-inch cente rs. Excess meta lwas milled chemically from between the strips o lighten the coldplate.In the original design, the passage through which the fluid travels was small,

approximately 0.044 inch wide and 0.100 inch deep. Distribution and collection pas-sages were provided at the inlet and outlet, giving the core the appearance of a ladder-type struc ture. Base and face sheet s of 0.030-inch 5052 aluminum were thensilver-eutectic bonded to the core or ladder. One side of the face she et was electro-plated with silver and then onded to th e coreby the application of heat and pressure.This proce ss resul ted in an assem bly hich had silver and aluminum eutecticas a bondbetween the core and face sheet, and this bond was exposed directly to the water andglycol coolant.The ethylene glycol and water solution is a moderately good electrolyte, and thesilver and aluminum couple in the coldplates hada high electromotive potential. There-fore, a galvanic corrosion problem developed in the coldplate, causing concern about

potential leaks and flow blockage. Also, the ratio of su rface area to coola nt volum ewas large becau se of the sm all pa ssage flow area in the coldplates. This conditiontended to cause local depletion of the in hib itors at points where the coolant fluidwasnot circulating, permitting corrosion to occur at these points. Also, the small pas-sage cro ss section made the coldpla te sensitive to obstruction from the corrosionproducts. Parallel effor ts w ere initia ted both to cope with the problem for the firstfew spacecraft and to eliminate the problem for the later vehicles.Special ground support equipmentwas designed to circulate the ethylene glycoland water coolant during periods of spacecraft activi ty when the coolan t c ircuitw a snot normally in use. This technique prevented local inhibitor depletion. Also, Sam-ples were withdrawn periodically to be analyzed for inhibitor concentration. When

this check indicated a decrease in the inhibitor level, the ethylene glycoland wate rcoolant circuitwas exchanged and filled with f re sh solution.A redesign of the coldplate w a s initiated not only to resolve the corrosion prob-le ms but also to improve the heat-transfer and fluid-distribution characteristics. Thecore is eloxed fr om a shee t of 6061 aluminum to obtaina staggered patte rn of0.05-inch-diameter pins on 0.125-inch centers, with a 60" angle between the centerlines. . The pin height va ri es fr om 0.062 to 0.172 inch with the coldplate application.

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The face sheet has a 0.003-inch coating of a fluxless braze alloy consisting of si liconand aluminum, thus eliminating the silve r from the bonding process. The new cold-plate, referred to as the pin fin configuration (fig. ll), represents an improvementover the ladder-type structure because of a lower surface-to-volume ratio .and be-cause of lar ger passages. The heat-transfer capability w as increased from l. 0 o3.0 W/sq in. for the average diss ipati on ate and from 2.0 to 6.0 W/sq in. fo r themaximum local rate.

(a) Pinnd inlet manifold details. (b) Coldplatessembly.Figure 11. - Pin fin coldplate.

Space RadiatorThe ECS space radiator for the Apollo spacecra ft w a s designed as an integralpa rt of the SM structure and originally consistedof two panels located on oppositeside s of the SM. Each panel had 30 square feet of radiator area. A four-tube circuitand a six-tube circuit were provided in each panel, with parallel flow through the tubes.The two panels were linked in parallelso that the flow path within the radiator w a sthrough 20 parallel tubes on four separate circuits, each circuit having an isolationcapability. This configuration resulted in a radiator having a minimum pressure dropand was consistent with optimizing the pumping power requirements of the coolantcircuit.Three factors made theselected configuration inadequate to reject the spacecraft

thermal loads. First, gradual growth of power requirements by other spacecra ft sys-tems resulted in thermal loads that exceeded the heat-rejection capabilityf the60 square feet of radiator area (3700-Btu/hr capability as compared to the 4850-Btu/hrrequirement for an average earth-orbital environment). The difference between thecapability and the requirement had to be rejected by water evaporation and consequentlycurtailed the mission-duration capability. Because the radiator was an integral partof the SM structure, it w a s not practical to increase thearea without a major impact

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on the program. Second, a problem associated with unequal fluid flow rates throughthe tubes was causedby the outer tubes on each panel being longer thanhe other tubes.The resulting effect was magnified by the disproportionately larger finareas associ-ated with the longer tubes. Third, no prov ision was made to con trol the flow betweenthe two panels, which a r e usually exposed to different radiant environments.When the radiator was exposed to a cold environment, the combina tion of thelonger outer tubes and the lack of t herm al isola tion from theSM, coupled with thevisco sity char acte ristic s of the ethylene glycol and water coolantat low temperatures,acted to induce flow stagnation in the outer tubes. A slight plumbing change to theouter tubes reduced the "sensitivity to flow imbalance, '' but it was still necessary toadd a supplemen tary water supply of 1 1 2 pounds in the SM to provide the requiredheat-rejection capability for a 14-day earth-orbital mission.The redesign of the CSM to a config-uration which implemented the lunar-mission capability provided an appropriatechange point to cor rec t the ECS radiatordeficiencies. The experience gained wasused to eliminateall th re e of the undesir-able features in the earlier radiator. Theresulting design consists of a primary anda secondaryadiatorystem.he in- l ine heatermary radiator circuit consists basicallyof

two radiator panels, each with an ar ea of50 square feet, located on opposite sidesof the SM (fig. 1 2 ) . With th is arra nge-ment, one panel may be exposed to deepspace at the time the opposite panel is ex-posed to a heat source such as the sun,earth, o r moon. These extre mes n en-vironments will produce large differencesin the effec tiveness of each panel and, con-sequently, in the fluid outlet temperatures.The panel exposed to deep space can re-ject more heat than the panel receivingexternal radiation; herefore, he overallheat rejectio n of the system can be in-cre ase d by increasing the flow to the coldpanel.

The flow through the radiators is S e r v i c e m o d u l econtrolled by a dual-flow-proportioning Figure 12 . - Selective tagnation/flowvalvessembly.uringperation, i f a proportioningadiatorystem.difference in radiator-panel outlet tem-perature occurs, the flow-proportioningvalve will be positioned to increase the coolan t flow to the colder radiator panel. A t atemperature diffe rent ial of 10" F, the flow-proportioning valve will divert approxi-mately 95 perc ent of the flow to the cold radiator. A redundant flow-proportioningsystem is provided; the system contains a logic network to initiate and indicate anautomatic switchover if improper operation occurs.18


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In situations when the radia tor inlet temperature is low and the panels haveafavorable environment for heat rejection, the radiator outlet temperature starts todecrease, and the hot-glycol bypass flow within the CM is initiated. A s mor e flow isbypassed, the radiator outlet temperature decrease s. An in-line heater upstream ofthe radiator is automatically turned on when the outle t temperat ure of the rad iat ormixed coolant drops to-15" F.In addition to flow proportioning and heater control, a pas siv e syst em of effectiveradiator area control called selective stagnation is incorporated. The two radia torpanels are identical, with five tubes in parallel and one tube inseries downstream ofthe other five. The five tubes have manifolds sized to provide graduated flow throughthe individual tubes. Thus, for equal fin areas, the tube with the lower flow rate willhave the lowest outlet temperature. As outlet temperature decrease s, the flow re-sistance in the minimum flow tube increases, thus further reducing thelow rate andoutlet temperature. A s the fin area around the tube gets colder, it draws heat fromthe a djacent tube, and the same process occurs with each successive tube. In a fullystagnated condition, there w i l l be essentially no flow in three tubes . The pri mary(minimum resi stance) flow tube w i l l ca rr y mo st of the flow with a sm al l flow throughthe adjacent tube.A s the heat load increases, the radiator inlet temperature increases; and moreheat is transferred into the stagnated tubes, resulting in success ive increases in flowra te and ultimately full panel operation. Therefore, at high heat loads the panelsautomatically provide a high effectiveness (completely thawed panels operating at ahigh average surface te mperature); and at low heat loads, the pane ls provide a loweffectiveness (stagnated panels operating at a low average surface temperature).The secondary radiator is provided as a backup in the eventof fa il ur e of theprima ry system . This radiato r consist s of four tubes placed close to the hottest pri-mary circuit tubes so that the ethylene glycol and water coolant in the secondary tubes

will not fr eeze when the secondary circuit is inoperative. The selective stagnationprinciple is not used in the s econdary r adia tor becaus e of t he sma lle r r ang e of heat-load requirements. This lack of a passive-control mechanism causes the secondarycoolant circuit to be dependent on the heater control system at low heat loads and onthe evaporator at high heat loads for contro l of the ethylene glycol and water coolanttemperature.

Evapo ratorThe glycol evaporator (fig. 13 ) us esa plate-fin sandwich construction in a cross-counterflow arrangement. The core is comprised of brazed modules of finned glycolpassages manifolded together and of s team pass age fins brazed to each sidef the ex-terior surface. Nickel felt-metal wicks of 15-percen t density are sandwiched between

the glycol modules, and the assembly is brazed to form the evaporator core. Thewater inlet and distribution plate s composed of a solid plate of sta inl ess ste el in whichchannels are milled for waterflow passages. To this plate is brazed a sinteredstainless-ste el porous plate which has a pore ra ting of 5 microns. This assembly dis-tributes the water uniformly over the entire surface f the plate and is bolted to a flangearound the base of the core. A cellulose sponge pad is compressed between the distri-bution plate and the baseof the wicks to ensure contact and uniform ater distribution

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Distribution wick

Waterinlet

(a) Ex te rior view. (b) Core detail.Figure 13. - Original evaporator design.

to all the wicks. To complete the component, the steam exhaust pan, with an integralflange to accept the backpressure valve, is brazed to the top of the core.The evaporator, which is automatically actuated when the fluid temperatureleaving the space radiator exceeds 49" F, presented a difficult developmental problembecause of three cons ide rat ions involving systems integ ration. The first factor in-volved the need for water management in the ove rall spacecraft mas s and energybalance. The glycol evaporator is required to operate frequently and at high evapora-tion rates because of the marginal heat-re jection capability of the radia tor when oper-ating in the relatively warm env ironments of earth orbit and lunar orbit. This fac tordictated an evaporator design requirement of high efficiency, that is, exit steam witha quality approaching 100 percent. The second fac tor involved the requirement foraccu rate and cons tant temperature control of the coolant supplied to the spacecraftguidance and navigation equipment, particularly the IMU. The design t h a t was selectedcontrolled the coolant temperature by controlling the pressure at which evaporationoccurred. A backpressure valve, close -coupled to and linking the evaporator with thesteam exi t duct, allowed the evaporator to be maintained in a wet condition, ready tofunction immediately upon demand should he radiator exit temperature begin to rise.The third factor involved the steam duct leading from he evaporator to space vacuum.The final design configuration wasa 2-inch-diameter duct approximately 8 feet inlength and having thr ee 90" bends because of trade-offs on equipment locations withinthe crew compartment and the spacecraft-a ttitude requirements during entry. Thelength and shape of the steam du ct emphas ized the requirement for free water control,because any free water leaving the evaporatorwould tend to freeze, accumulate, andeventually obstruct the passage.Early developmental testing of a control system, which metered water to theevaporator as a function of the coolant outlet tempe rature error, proved unsuccessfu lbecause of its slow response. The backpressure control method was then developed,and its performance is satisfactory for rapid startup and accur ate tem perature con-trol. However, problems were encountered in the control function which supplied themakeup water to the evaporator, and these problems merit further discussion.

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The backpressure-controlled evaporator configuration useda limit switch withinthe backpressure valve to serveas a permissive function. The switch enabled the wateror wetness control to be activated hen the backpressure valve moved off the fullyclosed position. The wetness control function w a s isolated from the backpressure (ortemperature) control function, except for the activation switch in the backpressurevalve. The major components that provided wetness control included a solenoid-actuated water-inflow valve, an electronics package or controller, an evaporator inletglycol-temperature sensor which served as a variable-reference temperature for thecontroller, and a wetness sensor. The wetness sensor consisted of a glass-sealedthermistor enclosed by a coil-type heater. The assembly w a s encapsulated in a metalcylinder. The sensor was located on the steam exhaus t pan of the evaporator abovethe co re and upstream of the backpressure valve so that the sensorw a s directly in thepath of the s te am flow. The concept involved the cooling effect of wet steam on theheated sensor and subsequent turning offof the water inflow to the evaporator . Devel-opmental testing demonstrated that the wetness sensorw a s not sensi tive to the qua lityof the steam and tha t drople ts of water impinging on the . sen sor wer e required to pro -duce the cooling effect necess ary to deenergize the water inflow valve. Furthe rmore ,it w a s discovered that the evaporatorw a s not actually wick-fed but w a s a pool boiler.When operated with the water inlet port orientedat the bottom (the attitude in which itis installed in the spacecraft), the level of free water in the core would vary as afunction of the heat.load being rejected. A t low loads, 500 Btu/hr , the core would benearly filled with water and boiling would occur at the top of the core. In this mode,droplets of water needed to t ravel only a short distance before they reached the wetnesssensor. Conversely, at the high loads of 7600 Btu/hr, no free water could be observed.

A control system which permitted free water to accumulate in the corew a s notacceptable for two significant reasons. The first reason involved the characteristicwhich the evaporator exhibited when exposed to an increasing heat load. Under con-ditions of increasing inlet glycol temperature, tests indicated that free water w a sejected from the evaporator. The mechanism for the liquid carryover appeare d o bebubbles of vapor which passed rapidly from the bottom to the topf the liquid mass,carrying liquid over as the bubbles broke the surface. Also, it was demonstrated thatthis quantity of water w a s sufficient to obstruct the steam duct. Tes ting in this attitudewould not be r epr esenta tive of flight conditions, because gravity forces were acting toretain the liquid in the core and, in the absence of gravity, a ia rg er quantity would beexpelled under similar conditions. This attention to predicted performance in theweightless environment w a s the other factorwhich caused concern over the operatingcha rac ter ist ics of the wetness controlfunction.A redesign w a s initiated (fig. 1 4 ) which requir ed t hat two basic changes to theevaporator be incorporated. First, the heater w a s removed from the wetness sensor,and the thermistor was relocatedso that it was embedded in a wick. This improvedthe response characteristicof the wetn ess senso r becauseconduction, rather thanconvection, became the mode by which the s ens or me asu red t he de greeof wetness in

the evaporator. Second, th e core was redes igned so that a sponge pad connected thetop edges of all wicks, thus permitting redistribution of water between the wicks tocompensate for nonuniformity in wicking rates. This redistribution capability resultedin a uniform wick wetness throughout the core and permitted the wetness sensor loca-tion in one wick o be repre senta tive of all wicks.

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S e n s o r e l e c t r i c a l c o n n e c t o rContact s p n g e l 7 W a t e r t r a n s f e r w i c k

(a) Exte rior view.

h

\

Water transfer wick

74Water redistribution pad A(c) View of sensor deta il .Figure 14. Modified evaporator design.

S i n t e r e d - m e t a ld i s t r i b u t i o n p l a t e sW a t e r d i s t r i b u t i o n sponges

(b ) Cutawayview.

Because of the problems encounteredwith the attitude sens itivi ty of the evapora-tor during testing and becausef the diffi-culty in extrapolating these test results toa prediction of performance ina weightlessenvironment, a requirement was placed onthe evaporator that it perform in all atti-tudes without wate r ca rry ove r. Althoughthis requirement was stringent, it wasdeemed necessary to demonstrate thatcapillary forces in the wicks dominate theeffects of gravity. A s pa rt of the redesigneffort to meet the all-attitude test criteria,the one-piece water distribution plate wasreplaced by nine small plates, each sep-arately fed by a water distribution mani-fold, for mproveddistribution. naddition, the material was changed fromsintered stainless-steel porous plate(5-micron rating) to a sintered laminateof nickel screens, each having a 12-microneffective pore size. This produced an in-creased pressure drop, permitting a moreuniform distribution of the w ater ov er thethe plate to the wicks.

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Qualification testing revealed that the small passages within the sintered nickelplates became obstructed becausef the formation of nickel oxidation products, whichbuilt up with exposure to water. Also, this small pore size was very sensitive to par-ticle contamination from the upstream plumbing. The nickel plates were subsequentlyreplaced by a st ack of five stainless-steel pla tes in which 0.005-inch-diameter holeswere drilled, thus producing a series-orifice effect which adequately met the uniformdistribution requirement and w a s much less sensitive to contamination. The finalconfiguration of the evaporator, w ith these changes incorporated, w a s flown on theApollo 11 mission, and the performance was entirely satisfac tory. However, to avoidperturbations of the spacecraft attitude, the evaporator was used only during launchand entry.The interaction of the var iou s con-trol funct ions of the coolant cir cuit is ex-plained in figure 5 which illustrates thesystemperformance of the Apollo 8 space- ,,,craft during lunar orbit. During theeighth revolution, the radiator outlet tem-perature decreased as the CSM entered the

darksi de of the moon. A t approximately $ 504 9 " F, theevaporatoroperation w a s ter- $minated,and he CM thermal oadbypass +w a s initiated to maintain the evaporatoroutlet temperature at a nominal tempera-tu re of 4 5 " F. Radiatoroutlet emperature " R e v o l u t i o n 8 - ~ R e v o l u t i o n 9 +Revolution 10"decreased at 18" F minimum as the CMS 10 I 1 - Ileft the lunar night and then increased to a Lun::nght -5 86 81 88 89maximum of 6 5 " F under the temperatureincrease as a re su lt of sola r and lunar ra - Figure 1 5 . - Coolant system performancediation.hevaporatorperation w a s duringhe Apollo 8 mission.initiated automatically as the radiator out-let temperature exceeded a 4 9 " F nominaltemperature.Thecoolant-loopheat oadswereessentiallyconstantduring hisperiod, as indicated by a stable radiator inlet temperature of 7 3 " to 7 5 " F.

80-'

-

Mission elasped t i m e , hr

Water SystemThe Apollo CSM wate r proces sing and distr ibution system s the product of in itialdesign concepts modified by extensive developmental and flight testing experience. Thedesign of the system for Apollo 11 and subsequent missions is adequate for the intendeduse but could not be considered optimized from systems-engineering standpoint.All potable water 'available in theCM, except for that initially loaded preflight,is supplied by the spacecra ft fuel cells located in the SM. Medical requirements dic-tate the repeated injec tion of chlorine into the potable water, for bacteria control. Thesystem distribution plumbing is such that all fuel-cell water is transferred to the CMand is either stored in thewaste- or potable-water tanks, taken directly to the useports, o r vented directly overboard. System design is such that available water willbe preferentially directed to the potable-water tank at any time the tank is not full.When the potable-water tank is filled, product water will be automatically divertedinto the waste-wa ter tank.

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During the system development and qualification testing and during spacecraftflight testing, several problems were identified and required either operational changor hardware modifications. The problem areas included plumbing corrosion , undesir -able chlorine taste, and excessive quantities of fr ee gas.After numerous development tests and evaluation of the water system and com-

ponents, the amount of plumbing corros ion w a s determined to be undesirable but represented no hazard to mission success for 14-day missions. In an effort to minimizesystem corrosion, operationa l constraints were made to preclude exposure of the sys-tem to chlorine until countdown operations began for launch. These procedures haveproved adequate to preclude inflight failures.During the early Apollo flights, the crewmembers reported on severa l occasionsthat the water had a strong chlorine taste. In most instances, the difficulty was tracedto a procedure error occurring during the injectionof the chlor ine and buffer ampules.When cle ar and concise procedures were developed and used, the crews had no objec-tion to the taste of the water.The potable water generated by the fuel cells is saturated wi t h hydrogen at thefuel-cell operating pressure and temperature . When the pressu re on the water is re-duced from the fuel-cell pressure (60 psia) to the cabin-use pressure (5.0 psia), largequantities of fr ee ga s evolve in the water. The fre e ga s consumed by the crew duringeating and drinking caused discomfort to the crewmembers. Water/gas separatorswere developed by NASA and ins talled to provide gas-free water for crew consumption.The water/gas separators (fig. 1 6 ) were designed to use the unique surface-tens ion characteris tics avai lable from hydrophobic (gas-permeable Teflon) and hydro-philic (stainless-steel filter mesh) membranes. In use, the two membranes are placedin close proximity to each other, and the water/gas mixture is forced to flow betweenthem (fig. 17). The hydrophobic membrane selectively passes the gas from the mix-

tu re while rejecting the liquid. The gas flowing through the membrane is dumped tothe spacecra ft cabin. The hydrophilic membrane, when wetted with water, allowswater flow, while acting as a barrier to the gas in the mixture. The water is collecteddownstream of the hydrophilic membrane and delivered as gas-free water for crewconsumption. The water/gas separator was designed to permit inflight installation onthe use ports (water gun o r food preparation unit) as shown in figure 1 8 .Operationally, the CSM water system has proved to be very satisfactory. Par-ticularly, the flight crews have noted that the availabilityof hot water for food anddrink prepa ration was desirable and should be pa rt of future spacecr aft system s.The chilled-water port w a s not used often because the water gun was more convenientand also provided chilled water.

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Figure 17. - Cutaway view ofwater/gas separator.Figure 16. - Water/gas separator.Waste Management System

The waste management system isrequi red to dispo se of c rew waste solidsand liquids. The equipment is designed tomeet requirements for collection and directtion and stowage of feces .

Gas separator ca rtridge overboardventing of urine and for coll ec-SEB-39104914The urine collection and transfersyste m cons ists of a receiver and collec-tion bag, a transfer hose, miscellaneousvalve s and plumbing, and an overboard

dump nozzle with an electric heater toprevent nozzle freezing. The urine collec-tion assembly consists of a pliable reser-voir bag interfaced to the spacecraftplumbing with a quick disconnect fittingand to the crewman with a roll-on cuff(rubber tube) that servesas an externalto r inlet valve. During urination, the

Food preparation panel

Water pistol catheteretweenheenisndheollec-Figure 18. - Inflight nsta llat ion of liquid is venteddirectlyoverboard. Whenwater/gaseparator.heicturition rate exceedshelow-ratecapability of the ove rboard dump nozzle,

the reservoir bag servesas a collector tobe vented overboard after terminati on of ur in e flow. Also, the volume of the reservoirbag is adequate to store one complete urination volumehen it may be undesirable tovent directly overboard. The roll-on cuffs tend to deteriorate during usage, and re-placement cuffs are provided for inflight replacement.Equipment forfecal collection and stowage includes the fecal collection assembly,cleansing tissue, and waste stowage compartment. The fecal collection assembly

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con sis ts of a fecal collector and container bag open at one end. The open end of thebag has a flange with a tape surface for adhesing the flange to the buttocks. Each bagis packaged with a w et cleansing cloth and a germicide. After use, the germicide isplaced in the bag and the bag is sealed. After the feces is perm eated with the germi-cide, the bag is stowed inside the waste stowage compartment. The capability is pro-vided to purge the waste stowage compartment of objectionable odors if necessary.

Both the urine colle ction and trans fer system and the fec al collection and stowsystem have been the subject of exten sive c ritic ism by mo st of the c rewm embers.After the Apollo 11 mission, the urine collection and transfer assembly w a s replacedwith a urine rece iver asse mbly (figs. 19 and 20) which, unlike the prior system, doesnot require crew contact with the unit to prevent urine spillage during us e.

Figure 1 9 . - Urine eceiverassemblyFigure 20. - Urine eceiverassemblyw i th cap removed.Cabin N o i s e

During the early Apollo missions, the crews registered numerous complaintsabout the excessive noise level created when the cabin fans were operated. Methodsof acoustically isolating the fans were identified but never implemented on flight ve-hicles. It was determined from additional flight experience that the fans were notmandatory for cabin thermal control for the typesf mis sio ns which were then planned,and the expense of the modification w a s not justifi able.

Both in Apollo missions and in long-term ground-based tests, it has been con-cluded that added emphasis mustbe placed on reducing or controlling the noiseoutputfrom spacecraft components and fluid flow sys tem s.

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CONCLUDING R E M A R K SDuring the course of the development-qualification and flight-test programs, theenvironmental control system has performed in an outstanding manner. Within theflight sys tem, there have been o environmental control system failures ha t haveplaced the crew or the mission in jeopardy. There have, howeve r, been malfunctions

which caused changes to planned operational procedures.During the period of qualification and flight testing, considerable difficulty wasexperienced with the coldplates used for electronic equipment cooling, the evaporatorsfor supplemental heat rejection, the radiators for primary heat rejection, the potable-water system, and the waste management system. The difficulties associated w i ththese items are of a basic design nature and should receive prime consideration onfuture programs.

Manned Spacecraft CenterNational Aeronautics and Space AdministrationHouston,Texas, May 24, 1971914-11-10-00-72

Documents

Apollo Experience Report Command and Service Module Environmental Control System