Apollo Experience Report Service Propulsion Subsystem

8/8/2019 Apollo Experience Report Service Propulsion Subsystem

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T E C H N IC A L N O T E

EXPERIENCE REPORT -PROPULSION SUBSYSTEM

R. Gibson a n d James A. WoodB. Johnson Space Center

I O N A L A E R O N A U T I C S A N D S P AC 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. A U G U S T 1973


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1 Report No.1- NASA TN D-7375

Lyndon B. Johnson Space CenterHouston, Texas 77058_ _ _ - _ _ _ ~12. Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationWashington, D. C . 20546 -__I-____-____II

I2. Government Accession No. I 3. Recipient's Catalog No.

1 1 . Contract or Grant No.

13. Type of Report and Period Covered__echnical Note-14. Sponsoring Agency CodeI --

_I_--I ---I.-5. Report Datet ugust 1973_____--- --"L_..-.---_._.--.-I__l_l____4. Tit le and Subtit let

17. Key Words (Suggested by Author(s)'Apollo Project*CommandSer vic e Modules.Spacecraft Propulsio n* Reliability Control* Propellants

APOLLO EXPERIENCE REPORT1 SERVICE PROPULSION SUBSYSTEM

la. Distribution Statement

6. Performing Organization Code

20. Security Classif. (o f this page) 21. No. of Pages9. Security Classif. ( of this repo r t )None None 25

-l____l____ - . _ - ~7. Au tho t (s ) -8. Performing Organization Report N o.

22. Price$3.00

JSC S-37810.Work Un i t No ..___I_ -eci l R. Gibson and James A. Wood, JSC- 111.1 ---I-._.9. Performing Organization Name and Address I 914-13-00-00-72


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CONTENTS

SectionSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .DESIGN PHASE . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Block I Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Block Il Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

DEVELOPMENT AND QUALIFICATION PHASE . . . . . . . . . . . . . . . . .Component Development and Qualification . . . . . . . . . . . . . . . . . .Subsystem Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Integrated Systems Te st s . . . . . . . . . . . . . . . . . . . . . . . . . . .Flight-Simulation Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . .

SUBSYSTEM FLIGHT RESULTS . . . . . . . . . . . . . . . . . . . . . . . . .BlockIFl ights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .Block I1 Flights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Page112366714161717171920

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TABLES

Table PageI RESULTS OF BLOCK I1 SPS ENGINE ALTITUDE TEST PROGRAMS

AT AEDC . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11I1 BLOCK 11WSTF TEST SUMMARY . . . . . . . . . . . . . . . . . . . 15FIGURES

Figure Page1 Service propulsion subsystem . . . . . . . . . . . . . . . . . . . . . . 32 Service propulsion subsystem propellant feed assembly . . . . . . . . 43 Service propulsion subsystem engine assembly . . . . . . . . . . . . 64 Service propulsion subsystem engine baffled injector . . . . . . . . 8

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APOLLO EXPERIENCE REPORTS E R V I C E P R O P U L S I O N S U B S Y S T E M

B y C e c i l R . G i b s o n a n d J a m es A. W oo dL y n d o n B . J o h n s o n Space C e n t e rS U M M A R Y

A review of the 9-year development program of the Apollo service propulsion sub-system fr om the initial concept to the Apollo 11 lunar landing is given in this report.During late 1960, separate contractors, under the supervision of the National Aeronau-ti cs and Space Administrat ion, prepared feasibility studies of an advanced manned space-craft program. In 1961, the lunar-orbit-rendezvous mode w a s selected to replace thedirect-lunar-landing mode. By 1967, development and qualification of the serv ice pro-pulsion subsystem were completed. The continued refinements and the evolution of theserv ice propulsion subsys tem and it s related assemblies we re prime facto rs in the suc-cessfu l manned flights of the Apollo Program.

INTRODUCTIONThis review of the Apollo ser vi ce propulsion subsys tem (SPS) covers the 9-yearper iod from the conception of the subsystem to the Apollo 11 lunar landing. During thi speriod, the program progressed fr om the definition phase to the hardware design phase,to the subsys tem development and qualification phases, and, ultimately, to the flightphase. Because of the siz e of the subsystem, it w a s not feasible to provide totalredundancy. For thi s reason, the SPS had to be extremely reliable. Simplicity of de-sign and extensive ground testing were required in o rd er to achieve the desired confi-dence level.qualification phases. These problems, and the actions taken to eliminate them, a r ediscussed in the descrip tion of th is phase of the program.

Several problems were encountered during the development and

In October 1960, the National Aeronautics and Space Administration (NASA) se-lected thr ee contractors to prepare feasibility studies of an advanced manned spacecraftas a pa rt of the Apollo Pr og ram. Several types of propulsion systems, cryogenic pro-pellants, storable propellants, and solid-propellant rocket motors were studied to sup-por t the various proposed Apollo configurations. By early 1962, the basic Apolloconfiguration and the engine subcontractor had been selec ted.The evolution of the subsystem began with the definition of design cr it er ia thatwould be consistent with mission requirements, reliability goals, and spacecraft-designconcepts. The development and qualification of the subsystem were accomplished by


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developing and testing components, by testing major assemblies, and by full-scale s y s -tem testing on propulsion test f ixtu res. Detailed test objectives to verify inflight sys-tem performance also were incorporated in the ear ly Apollo missions.Analyses that encompassed conceptual studies , weight trade-offs, configurationchoice, size, and establishment of performance c ri te ri a resulted in a final design defi-

nition for the subsystem. Materials and proc es ses were investigated to support subsys-tem development. In connection with the analytical studies, laboratory r es ea rc h wasconducted to verify analytical techniques, to improve component design, and to resolveproblems. Material-properties r esea rch was conducted to determine the emissivit iesof nozzle and nozzle-coating materials. In addition, nozzle-material welding techniqueswere investigated. Tube brazing and weld techniques were improved by means ofpropellant-metal compatibility s tudie s and brazing-welding metallurgical investigations .Thrust-chamber ablative material s were selected after the performance of laboratorytests that limited the mater ials list before any thrust-chamber testing. Laboratorystudies were conducted on 42 potential thrust-chamber-material samples; the studiesincluded high-temperature vacuum te st s and thermal- and structura l-proper tiesinvestigations.Seal materials for propellant equipment were selected after investigation of elas-tomer and pseudoelastomer compounds to s cr ee n fo r propellant compatibility, swell,creep, resilience, and other required seal pro per ties. Zero-gravity propellant-motionproblems were investigated by means of theoreti cal and experimental r es ea rch in fluidmechanics. The goals of th is re sea rc h were new modeling and scaling techniques fo rea rt h simulation of zero-gravity effects and an improvement in the understanding offundamental phenomena. The simulation techniques and facil ities that were used we rethe prime cont ractor drop tower (low gravity), scaled transparent tanks (one g) withslow-motion picture simulation of inflight real time, and the U . S . Air Fo rc e KC-135airplane flying laboratory.The command and ser vi ce module (CSM) flight progra m was the final phase of theSPS development. Inflight testing of the SPS was accomplished in sequence with thevehic les to produce a high-confidence, proven subsystem for the lunar-orbital-rendezvous mission. Because the SPS played a major ro le in the achievement of crewsafety and mission success , an attempt to qualify the subsystem under all space-operational conditions w a s made during the ground-tes t program. However, limita tionsin the ground-test facilities; limited zero-gravity periods i n earthbound test vehicles;the impracticability of simulating the combined pr es su re , temperatur e, and gravityenvironment; and environmental unknowns prevented complete demonst ration of subsys-tem performance conditions before and during the earl y unmanned flights. Thus, theSPS was used conservatively ear ly in the flight pro gram, but the complexity of operating

modes and of subsystem demands was in crease d with each flight as experience and con-fidence were obtained.

D E S I G N PHASEDesign reliability was a product of simplicity. Propellants wer e selected on thebas is of experience with other pr og ra ms and because of propellant earth-storability ,

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hypergolicity, performance, and high-density properties. A pressure -fed engine,which eliminated the need for the par ts and controls that ar e required fo r pump-fed en-gines, was selected to decrease the operating complexity. High-pressure (4400 psia)helium, stored at ambient temperature and regulated to 180 psia, was selected as thepropellant- tank pressurizing agent. An ablative thrust chamber was selected ratherthan a regeneratively cooled chamber to decrease the possibility of propel lant freezing.Also, an incr ease in the operational limit s fo r the propellant r atio, the propellant tem-pe ra tu re s, and the chamber pr es su re was more compatible with ablative chamb ers. Inaddition, studies were indicative that most propulsion-system failu res were caused byfailu res of control s, valves, and solenoids rather than by failure s of in jecto rs o r thrustchambers. There fore, redundancy was used, where practical, to incr ease reliabilityin these ar eas .

Approximately halfway through the SPS development program, data generatedfr om an in-house-managed re se ar ch and development contract were indicative that theSPS could supply a specific impulse of 3 to 5 seconds higher i f the oxidizer-to-fuelweight ratio was 1.6: 1 rather than 2 : 1. The higher performance level was desirableto provide fo r the lunar module (LM) weight growth and for additional flexibility in mis-sion planning. Therefore, the SPS was redesigned fo r the new ratio. The 2: 1 ratiowas used in the Block I vehicles. The vehicles using the 1 . 6 : l ratio were designatedBlock II.Block I Configuration

The SPS Block I mission and operational requirements dictated that the subsys-tem consist of a helium-pressurization assembly, a propellant-supply and propellant-distribution assembly, a propellant-utilization and propellant-gaging assembly , arocket-engine assembly, instrumentation,ment is shown in figure 1.

storage tank

Figure 1.- Service propulsionsubsystem.

and displays and controls. The SPS arrange-Helium-pressurization assembly. -The SPS helium-pressurization assembly(fig. 2) consisted of a high-pressure heliumsGpply; two spherical helium tanks; and theassociated pres su re regulators, isolationvalves, check valves, and pressure-rel iefvalves. The tanks wer e connected to acommon helium-distr ibution line. Twoparallel solenoid-operated poppet valves,

installed downstream from the helium tanks,isolated the helium supply when the propul-sion system w a s inoperative. Pr es su re wasregulated by two pressure -regulato r assem-blies. Each regulator assembly incorpo-rated a prima ry and a secondary regulatorin ser ies . The primary regulator reducedthe high upstream pr es su re to the down-str eam operating pr es su re . The secondary

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Oxid izerstoraget a n k

Oxid izerf i t t ingquick-di

L

Figure 2. - Service propulsion subsystem propellant feed assembly.regulator w a s calibrated to regulate a t a higher pressure; thus, the secondary regulatorremained open when the pr im ar y regulator functioned properly . The installation ofparal lel-s eries check-valve assembl ies downstream fr om the helium-pressure regula-to rs prevented backflow of propellants into the pressuri zat ion system. An oxidizer-and a fuel-pressure-relief-valve assembly (each consisting of a relief valve, a burstdiaphragm, and a filter s cre en) were provided fo r the propellant-tank sys tem s. Thebur st diaphragms were used to isolate the valve se at s from the propellant vapors.

Propellant supply and distribution assembly. - The propellant supply and distribu-tion assembly (fig. 2) contained and distributed oxidizer and fuel . The oxidizer supplyof the assembly was contained within two cylindrical tanks that had hemisphericaldomes. The two tanks were connected in ser ies by means of a propellant-transfer line.The upstream tank was used as the sto rage tank; the downstream tank w a s used as thesump tank. The oxidizer tanks contained 30 000 pounds of usable propellant, and thefue l tanks contained 15 000 pounds of usable propellant after allowances were made fo rthe loading tole rances, the res iduals , and the required ullage. A propellant/heliumheat exchanger was incorporated in each propellant line fo r the rma l conditioning of thehelium supply.

Titanium was determined to be the best pressure-tank material because of itshigh strength-to-density rat io and i t s compatibility with the propellants . Initially, thegross weight of the tanks w a s 49 850 pounds; the gros s weight was reduced to39 500 pounds in July 1962 because of the change to the lunar-orbit-rendezvous mode,

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which requir ed le ss propellant and tankage. The maximum tank pr es su re was 240 psig,and the operating-temperature limits were 104" F maximum and 44" F minimum.Zero-gravity expulsion techniques that were studied included the use of mechanicalbellows, an umbrel la spring-loaded bladder, full bladders, and reaction-control-engineullage maneuvers. The reaction-control-engine ullage maneuver w a s selected to settlethe propellants f o r expulsion. However, th is technique necessi tated the design and de-velopment of the propellant-retention re se rv oi r in the sump tanks as a backup to the re-action cont rol system (RCS).Propel lant utilization and gaging syst em. - The propellant utilization and gagingsystem (PUGS) consisted of primary and auxiliary propellant-quantity-sensing devices,an electr ical control unit, an oxidizer-flow-control-valve assembly, and a crew dis-play panel. The syst em w a s used to monitor the quantities of usable propellan t thatremained i n the propellant tanks so that the desired oxidizer-to-fuel rat io could be ad-jus ted manually during propellant expulsion fo r simultaneous depletion of the oxidizerand fuel. The primary quantity se nso rs were cylindrical capacitance probes that weremounted axially i n each tank. The auxiliary gaging syste m had impedance-type point

se ns or s coupled with a nominal-flow integrator between sen sor levels . Oxidizer flowwas controlled by means of a motor -operated, redundant, double-blade valve assemblythat was used to provide increased, decreas ed, or normal oxidizer flow ra te s. Thecontrol unit wa s used t o compute tota l propellant quantities fr om individual tank quan-tit ies, the propellant imbalance, and the oxidizer-to-fuel r atio; also, the contro l unitcontinuously compared the total propellant quantities that were indicated by the p rimaryand auxiliary s ys te ms . The crew display panel provided the onboard output indicationsthat wer e requir ed and provided the switches for use i n control functions and onboardtesting.Engine assembly. - The SPS engine (fig. 3) w a s a nonthrottleable, gimbaled,pressure-fed rocket engine that consisted of an ablative-cooled thrust chamber, a

radiation-cooled nozzle-(extending from an a r e a ra tio of approximately 6: 1 to 62.5: ),a bolt-on aluminum injector, a thrust and gimbal mount, a bipropellant valve, gimbalactuators, an electri cal harn ess , and propellant feedlines. The engine operated at achamber p re ssu re of 102 psia and produced a vacuum thrust of 21 500 pounds. Theaverage specific impulse was 309 seconds. The engine w a s capable of at least 36 s ta r t sand had an engine firing life of 500 seconds.Ignition occurr ed by means of hypergolic reaction in the th ru st chamber. Propel -lant flow to the thrust-chamber assembly w a s controlled by a redundant s et of series-pa ra lle l ball valves that we re actuated by pneumatic p re ss ur e that was controlled byelectr ically operated solenoid valves. Gaseous nitrogen, stor ed in redundant tanks,provided the pneumatic pre ss ur e.A flight combustion-stability monitor (FCSM) system w a s designed to monitorSPS engine vibrations and automatically commanded engine shutdown when unacceptablecombustion-excited vibration conditions occurr ed. This system contained engine-mounted ac cel er ome ter s and associated monitor circui ts and an electronic summingcircui t, which acted to shut down the engine. An indica tor lamp in the command mod-ule (CM) was illuminated when a shutdown was commanded. Manual control allowedoverriding the automatic system to re st ar t the engine. A combustion-stabilityverification-test progr am was conducted to show the inherent stability ch ar act er ist ic s

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I

Figure 3 . - Service propulsion sub-sys tem engine assembly.

of the Block IT engine configuration. Suc-cessfu l completion of thi s tes t prog ram al-lowed deletion of the FCSM.

Block I ConfigurationThe two major changes in the Block IISPS we re the establ ishment of the operatingpropellant ratio as 1.6 pounds of oxidizerpe r pound of fuel, to provide higher enginespec ific impul se, and the reduction of thetotal onboard propellant quantity as a resul tof vehicle-trajectory changes combined withthe improvement in engine performance.The changes faci litated making the propel-lant tanks smaller; and, because thepropellant-density rat io was 1.6: 1 (oxidizer-

to-fuel), the oxidizer and fuel tanks weremade identical in size. The cylindrical sec -tion of all four tanks was shortened by ap-proximately 11 inches . The Block Idiamete r and the hemispherical-head de-sign we re maintained. An additional advan-tage of the redesigned tanks was reali zed:the sump and storage tanks wer e installedin adjacent bays in the service module (SM) ,ra th er than on opposite si des (as in Block I),for a more desi rabl e location of the centerof gravity. The Block II propellant tanksalso w er e designed to have a l imit pressu reof 225 psia, which was a reduction fro m t h eBlock I value of 2 4 0 psia . Th is change facilitated fu rt he r reduction of the tank wallthicknesses. To reduce the tank limit pre ss ure to 2 25 psia and yet maintain the Block Ipermissi ble ullage-pres sure-ris e design limit of 213 psia, a narrow-range relief de-vice was developed for us e on Block 11 spacecraft.Because l e ss propellant w a s expelled by the Block 11pressuri zation equipment,le ss helium was required. The design loading pr es su re w a s reduced to a value thatcorresponded to the new helium quantity, and the wall thickness was reduced to thatthickness which was required fo r the lower loading pr es su re ; thus, an additional

weight saving was achieved.DEVELOPMENT AND QUAL1F I C A T I O N PHASE

The time-sequenced planning and resul ts that we re used in the testing of eachmajo r component and assembly of the subsystem a r e discussed in this section.

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Com ponen t Deve lopm e n a n d Qual i cat io nComponent development was conducted at the plant of each subcontractor. Com-ponent compliance with design criteria and the determination of optimum componentconfigurations were the purposes of the effort.Fo r most components, the flight phase s tarted before the development and quali-fication phase was complete. This fact made i t necessary to qualify hardware specifi-cally for the first flight and to delay the more general overst ress-t ype te st s until finalhardware designs were completed. A chronological descr iption of the development andqualification of each component in the engine assembly, propellant-supply assembly,and pressuriza tion assembly is given in the following sec tions.Engine assembly. - The SPS engine (f ig. 3) design and development effort wasst ar te d by the subcontractor April 9, 1962. The effort that was contracted originallycovered the design and development of t he SPS engine, preliminary flight-rating test s,and the delivery of two mockups, five prequalified engines, seven qualified engines,sp ar e pa r ts to support the delivered engines, and some ground-support equipment (GSE).The init ial design effort was concentrated on layouts of the overa ll engine conceptand on the interf aces with the spacecraft. By August 1962, the design definition hadprogressed sufficiently to per mit t he s ta rt of fabr ication of the ha rd mockup enginecomponents. The design review of thi s mockup wa s completed, and the mockup wasdelivered in November 1962.Throughout the engine development and qualification phase, many configurationchanges occurr ed in both the Block I and Block I1 engines as a resu lt of knowledgegained in the test pr ogr ams. Engine-development testing was conducted at the con-tr ac to r sea-level test facility at Sacramento, California, and at the simulated-altitude

te st facility at the Arnold Engineering and Development Center (AEDC), Tullahoma,Tennessee. One of the more significant changes resulted in the incorporation of abaffled injector to reduce the r isk of combustion instability. Both baffled and unbaffledinjectors we re used in the development program. The baffled injector is shown in fig-u r e 4 . The requirements for the engine to damp pr es su re oscillations and the poor t estexperience with unbaffled inj ectors necessitated the us e of a baffle to damp pr es su reoscillations i n the combustion chamber. Many te st s included the use of explosivecharges i n the combustion chamber during engine fir ings to verify the damping capa-bil ity of the engine.The Block I test program, performed at the contractor facility, consisted of ex-tensive firing of the injector as a component and fir ing i n conjunction with other engine

components. The sea-level qualification-test program consisted of fi ring the engineassembly 56 ti me s. The altitude testing at the AEDC consi sted of t hr ee te st phases:development, prequalification, and qualification testing.The first firea ble tes t engine (SN 003) was shipped to the AEDC in May 1963 f o rinitiation of the phase I program. Two additional SPS engines (SN 004 and SN 008) we reshipped to the AEDC in September and November 1963. Simulated-altitude testing beganwith engine SN 003 on June 26, 1963. An engine fi ring of approximately 5 seconds re -sul ted in shutdown because of the col lapse of the titanium nozzle extension, which was

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Figure 4 . - Service propulsion subsystem engine baffled injector.caused by excessive back pr es su re in the test cell . High-frequency cha mbe r- pr ess ureoscilla tions, which occ ur re d during the acceptance testing of engine SN 004 at the sub-contractor plant, we re attributed to incomplete air removal f ro m the engine oxidizercircuit. To ensur e removal of all gas fr om the oxidizer circuit, the engine-air-bleedsys tem on the engines w a s increased in size fr om 0.25 to 0.50 inch, and the requiredair-bleed times wer e lengthened for the phase I testing.

Two engines (SN 009 and S N 011) were used for the phase I1 prequalification test-ing at the AEDC . The f i r st altitude-test-cell fi ri ng involved engine SN 009; the testwas completed i n December 1964. In January 1965, because of inadvertent operat ionof the te st stand during stand maintenance, th e engine sustained nozzle and chamberdamage. No co rr ec ti ve action regarding engine design or operation w a s required be-cause the damage w a s attribu ted to test-stand problems. The second engine (S N 011)was equipped with a baffled injector that was designed to minimize combustion instabil-ity . By June 1965, the two engines had completed 101 te st s, fo r a total fi ri ng time of2581 seconds. The tes ts were used to evaluate engine operation and per formance overa wide range of chamber pr es su re s and propellant mixtur es. Simulated-altitude gim-baling of the engine was accomplished in March 1965 on engine SN 009 and was repeatedsuccessfully on engine SN 011 in April 1965.

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Engine SN 009, retrofitt ed with the fi rs t pneumatically actuated valve, was tes tedat simulated altitudes; the engine completed 27 te st s (408 seconds of firing time) inJune 1965. The chamber forward flange failed during the last test, which resulted inthe lo ss of the chamber and nozzle extension. The chamber-flange fai lure occurredagain on the second engine (SN 011) in July 1965. The fai lur e occurr ed on the 27th te staf ter 333 seconds of firing time. Both engines sustained extensive damage and were re-turned to the subcontractor. The fa ilur es were caused by shrinkage of the chamberablative lin er; shrinkage occ urred during cooldown after exposure to hot firing condi-tions. This shrinkage caused the opening of a gap at the chamber-to-injector flangejoint, allowing circulation of the combustion gasses in the joint, subsequent cha rringof the chamber, and overheating of the chamber-to-injector flange. Several firing cy-cl es with the associated heating and cooldown of the chamber and flange joint causeddegradation of the joint to the point of failure by means of separation. The chamberwas redesigned by step machining the outer liner surface and by adding a mechanicallockring. This eliminated the susceptibility t o thermal cycling. Burn-time and coast-time limitations were imposed on flight engines that were not retrofitted with the newchamber.

Thre e engines completed the phase 111 simulated-altitude qualification test ing;130 fir ings were made for an accumulated duration of 3599 seconds. During the f i rs tte st ser ie s of the third engine, se ve re chamber damage occur red that was attributed tovacuum-grease contamination of the fuel. This contamination caused overheating andwarpage of the injector flange, which resulted in burn through at the injector-flange-to-chamber joint on the second test se ri es . The injector was replaced, and the last twotest s er ie s were completed satisfactorily . The contamination problem was eliminatedthrough cleanliness controls during subsequent testing.Concurrent with the phase I to 111 AEDC tes t program, extensive developmenttesting was accomplished at the subcontractor sea-level t es t facility to establish com-patibility of engine components and for selection of the individual components to be used.Engine SN 023 completed acceptance testing in M ay 1965, and engine SN 021 completedacceptance testing in July 1965. As a result of the chamber burn-through problem atthe AEDC, the engines we re retrofit ted with the redesigned cha mber. Qualificationtesting was resumed, and the fi r st sea-level qualification-test ser ie s was completed onengine S N 022 in November 1965. Two additional test series were cnrnpleted by Decem-be r 28, 1965; this marked the sati sfac tory completion of the sea-level qualificationtest ing. A total of 56 tes t s was performed, representing 1518 seconds of firing time.Block I flight engines, which were of the sam e configuration as the qualificationengines, were processe d through a standard test cycle that consis ted of component ac-

ceptance tests, engine assembly acceptance firing, and postfire testing.The Block I1 engine-tes t prog ram consisted of 392 development-test f irings at thesubcontractor sea-level facility and three te st phases performed under simulated-altitude conditions (a total of 663 firings) at the AEDC. In May 1966, the first Block I1development test ing under simulated altitude was conducted (phase N). Forty-ninetests were performed, and 810 seconds of firing time we re accumulated. Phase IVtesting involved two engines; 265 te s ts were performed, and 6704 seconds of fi ring timewere accumulated.

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Phase V , which w a s .simulated-altitude qualification testing that consisted of sixtest se ri es , began November 18, 1966, and was completed in February 1967. A secondengine was used fo r th e last two series; 108 tests were performed and 4521 seconds offiring time were accumulated without unscheduled interruptions. Engine operation wasevaluated over the extreme range of thrus t-chamber pr es su re s, propellant ratios,fir ing durations, and propellant temperatures. One significant problem was encoun-tered: leakage of the ball-valve s ea ls was noted aft er testing. Subsequently, the sekl swere redesigned to provide a second seal fo r each ball, and the seal material waschanged from TFE Teflon to a glass-filled Teflon (BF-1 Blue Teflon) seal.

In addition to the ball-valve seal-leakage problem that was encountered duringphase V testing, overboard leakage of gaseous nitrogen through the actuator piston w a spossib le because of shrinkage of the Delrin pistons at temperatur es below 30' F. Toeliminate t h i s problem, the piston materi al was changed to aluminum. The phase VItesting, which consisted of 222 firings, w a s designed to qualify the design changes thatwere made to eliminate these problems and to evaluate the magnitude of chamber-pr es su re overshoots by the use of high-response instrumentation.The ball-valve seal-leakage problem was not eliminated completely by the rede-sign, but the resulting total leakage rate was determined to be acceptable for flight use.The Delrin piston had been replaced with an aluminum piston and a Delrin sliding sur-face; the replacement was satisfactory. The chamber-pressure-overshoot evaluationwas indicative that excessive pres su re spikes may re su lt fro m igniting the engine byopening both the redundant flow paths simultaneously. A revised operating requir e-ment, the use of only one flow path to st ar t the engine, was incorporated for flight use.The primary bas is for qualification of the engine was the simulated-altitude test ing con-ducted at the AEDC. The tests accomplished on the final (Block II) configuration a r esummarized in table I. The maturity and reliabili ty of the engine design were estab-lished through the number of fi rings conducted and the conditions simulated.Propellant utilization and gaging sys tem. - During the development and qualifi-cation testing of the PUGS, se ve ra l design changes were needed to eliminate discrep-ancies. However, other problems associated with the system also resulted fr ominteractions between the PUGS and the SPS tanking ar rangement .In the first generation of cont rol unit s, the connector panel was integrated withthe control-unit housing. Thus, technicians had to solder approximately 400 wires ina limited space. Solder inspection was difficult because only one row of solder cupsper connector could be seen at any one time and some solder cups had more than onewire. To eliminate thi s problem, the control unit was redesigned and the connectorpanel w a s separated fr om the housing. Therefore , th e connectors were wired in a

separa te subassembly, allowing adequate inspection.The original design for overload protection involved the use of fu se s packaged inthe modules. This method w a s costly because the module had to be unwired, depotted,repotted, and reinstalled when a fuse was blown. As a result , a design change wasmade so that all fus es wer e relocated on an accessible terminal board. This reducedthe time and cost required to repair units with blown fuses.Several other operating problems were noted during subsystem development test-ing at the NASA White Sands Te st Fac ility (WSTF) and during the flights. A propellant


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bias that resulted in off-nominal fue l flow was detected by means of the PUGS duringthe early flights. Initially, it was believed t o be a gaging problem, but la ter i t was es-tablished a s an increased fuel-flow rate over the rate that w a s predic ted based onground tests . This phenomenon could not be reproduced in helium-saturation groundtests. The propellant-utilization ( PU) valve was used i n the increase d oxidizer flowposition to eliminate most of the flow imbalance until engine reor ificing eliminated thebias.

Another bias i n both fuel and oxidizer flow w a s observed on the spacecraft 001(SC 001) tests at the WSTF and on early flights. The propellant level in the cylindricaltube, which houses the gaging probes in the sump tanks, was lower under flow condi-tions than was the level of the bulk propellant in the tanks. This inequality resultedfrom a Bernoulli effect that w a s caused by propellant flowing out of the bottom of thetank. Flow dividers were added in the retention res er vo ir to eliminate thi s problem.

Two other e r r o r s associated witin the PUGS have been identified mc!a r e accountedfo r i n preflight predictions of propellant usage. One er ro r resulted f ro m absorption ofhelium from the ullage of th e sump tank, which allowed the propel lant level to riseabove the top of the gage. This er r o r was compensated fo r operationally. The othere r r o r resulted fr om an offset calibration of the oxidizer-storage-tank probe, which wasnecessary to eliminate a residual signal fr om the empty storage tank. The probe offsetcalibration is incorporated during gaging system checkout pr io r to servicing. Thise r r o r is also compensated for operationally.

Flight experience also was indicative that approximately 25 seconds of engine op-eration ar e required to sett le propellants i n the gaging probe housings. This circum-stance caused activation of the caution and warning sy ste m, indicating a criticallyunbalanced condition between the remaining fuel and oxidizer quanti ties dur ing theApollo 9 mission. The design est ima te of settling time required was approximately4 seconds. The caution and warning sys tem was designed to be activated approximately5 seconds after ignition, and the sy stem compared the quantity of fuel and oxidizer thatremained. The caution and warning activation system was revi sed to delete the PUGScomparison and thus eliminate erroneous act ivation of the caution and warning syst em.

The auxiliary gaging sys tem al so failed on sev er al flight vehicles during groundcheckout. Leakage of conductive fuel vapor into the sealed electronic package of thefuel probe resulted i n shorting of the point-sensor electronic cir cuit s. This problemw a s not eliminated because of the significant cos t and schedule impact and the lack ofcriticality of the auxiliary portion of the gaging sys tem. The auxiliary sys tem was usedonly i f the primary system failed. Also, other methods were available to establish on-board quantities such as acceleration, helium usage ver su s burn time, and predictedpropellant flow rat e times burn time.Propellant-distribution assembly. - The Block I propellant-tank qualification tes tbegan in October 1963. Four fuel tanks and four oxidizer tanks were used to qualifythe propellant tanks. The major problems that were encountered during qualificationtesting ar e described as follows.In October 1963, a fter p roo f- pr ess ur e testing at 320 psig for 30 minutes, thenumber 1 fuel tank had a local meridional c ra ck in the lower dome section just belowthe weld joint. The crack in the number 1 fue l tank was caused by st re ss corrosion

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that resulted from a localized contaminant. Manufacturing pr oc es se s were revised toeliminate the us e of materials containing halogens fr om pr oc es se s at temperaturesgrea ter than 500" F. The cracked lower dome and cylinder were replaced; subse-quently, the tank w a s tested successfully.

In March 1964, a par tia l vacuum that w a s applied inadvertently to a fuel tankcaused the tank to buckle. The tank retur ned to acceptable dimensions after the vacuumwas release d. The tank w a s returned to the qualification program; no damage wasnoted after a thorough inspection. The tes t stand and procedures were modified to pr e-vent a re cu rr en ce of the problem. The Block I qualification testing w a s completed inJanuary 1965.Subsequent s t r e s s corr osio n fai lures in RCS oxidizer tanks and in one of thespacecraft 01 7 oxidizer tanks requi red additional SPS tank-compatibility tests. Thesefailur es were caused by s t r es s corrosion resulting fro m a pr oc es s change in the manu-facturing of ni trogen tetroxide (N204 ) . It was determined that the use of additives to

incr ease the fr ee nitrous oxide content of the oxidizer resulted i n satisfactory compati-bility with the tank material. A l l of the N 0 that w a s used in the Apollo veh icles had2 4to meet the ni trous oxide content requirement. A Block I SPS propellant tank was sub-jected to an 80-day compatibility test to demonstrate satisfactory results.

During the flight of spacecraft 009 (SC 009)February 26, 1966, he transfer linein the reser vo ir f ailed, causing helium to be ingested into the SPS engine, which re-sulted in loss of thr ust . The re se rv oi r was modified by strengthening the standpipesupport brackets and weld joints and was recertifi ed by means of vibration testing dur-ing the Block I1 propellant-tank qualification. After completion of the Block I1 qualifi-cation test , the propellant-retention re ser vo ir had cr ack s in each of six welds thatconnect the outlet po rt s to the main body of the re se rv oi r. Tru ck transportat ion of thetanks with the retention re ser vo ir installed w a s done only on the qualification-test tanksbecause the r ese rv oi rs were installed at the prime contractor facili ty. The manufac-tu re r indicated that the c ra ck s could have occurred during transportation of the unit;however, the retention re se rv oi r had not been disassembled for inspection between in-termediate s tages of testing, and the cr ac ks could not be definitely attributed to tra ns -portation. Therefore, a Block I and II retention-reservoir qualification ret est wasconducted to verify the integr ity of the crossover-tube welds in the environmental con-ditions that the retention r es er vo ir would undergo during boost and space flight. Bymeans of these te st s, the st ru ctur al integrity of the retention re se rv oi r was shown.

Helium-pressurization assembly. - Qualification testing of the first helium tankbegan in October 1963 and w a s completed i n April 1965. A major problem was encoun-ter ed when qualification-tank units t hr ee and four bur st at less than the design bur stpr es su re . By means of a design review, it was determined that the cause of fai lu rew a s excessive stress concentration in a heavy girth-weld bead. Removal of the weldbead allowed the weld joint to work in unison with the membrane during cycling. Twotanks f rom which the weld bead was removed were added to the qualification-test pro-gra m; therefore, the program w a s completed successfully.Qualification testing for Block I regulators w a s performed fro m July 1 to Novem-ber 3, 1965. Although the two units met the qualification-test requirements, the

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primary stage of the class IV regulator leaked excessively after endurance-cycle tests.A higher internal-leakage-specification limit w a s accepted because the leakage throughthe redundant secondary stage w a s le ss than the requir ed limit. Block 11 regulatorscompleted the qualification-test p rogram satisfac torily .The helium check valves were redesigned after the Block I test program to reduce

high leakage char acte ris tics . The pr ima ry change w a s the deletion of Teflon poppetsea ts and the incorporation of Resis tazine 88 as the poppet mater ial on the oxidizervalve to comply with leakage requireme nts of 1.08 scc/h r . Qualification testing w a scompleted on the helium-isolation valves without major redesigns.Subsystem Tests

During the development of the SPS, a comprehensive ground-test program w asconducted at the WSTF. A high level of confidence in the rel iabili ty of the basic, simpleconfiguration of the SPS resulted from these te sts . Various portions of the test programa r e summarized according to the test vehicle.From September 1964 to September 1965, t es ts were conducted at the WSTF usinga Block I te st fixtur e. The test ri g consisted of a boilerplate configuration that simu-lated the spacecraft propellant-line si ze s and routing and that had the neces sary i nstr u-mentation and safety provisions that were required for static-test operations.The initial test ser ies w a s conducted using a preprototype engine and off-the-

shelf hardware i n the propellant and helium-pressur ization sys tem s to establish opera-tional procedures for fluid servicing and to evaluate system-operating cha ract eri stic suntil flight components were available. Late r, the helium-pressurization syst em w a supdated to the flight configuration. The engine w a s updated continually throughout theprogram; a Block I qualification engine w a s installed fo r the final test seri es.Two injecto r fai lu res occurred during the fourth series of tests. In both cases,pos ttes t inspections were indicative that the hub of the injec tor baffle had separatedfrom the baffle. This problem w a s caused by afterburning of propellants in the injec-tor . *The njector w a s purged with nitrogen in all subsequent SPS ground test s to r e-duce the thermal stress caused by burning of res idual propellants i n the baffle.In preparation f or Block I1 testing, the Block I test fixture w a s modified exten-sively to a Block I1 configuration. The major di ffe rences between the flight configura-tion and the test configuration are listed as follows.1. Propellant tanks : Test -ar ticl e propellant-storage and sump tanks wer e boil-erplate, but the wetted surfaces were flight configuration.2. Helium-storage tanks: The helium supply for phases I and I1 and for series Iand 11 of phase 111w a s provided by nonflight-configuration external GSE. Series I11and IV of phase III t a d flight-configured Block I helium-storage tanks.The Block I1 SPS test program that w a s conducted at the WSTF began in Novem-ber 1966 and was completed in April 1969. The program consisted of sea-leve l t es ts

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that were conducted in th ree phases. Phase I consisted of eight tes t s er ie s; the objec-tives were sys tem verification and performance demonstration under nominal, off-nominal, and malfunction conditions. Phase I1 consisted of seven te st s er ie s; theobjectives wer e to show lunar-mission performance at normal and var ious abnormalconditions. Another objective w a s to evaluate the improved (double sea l) SPS enginebipropellant configuration. The final test phase, phase JJI, consisted of four test series;the objectives were to investigate SPS performance under ext reme off-limit conditions,to investigate flight anomalies, and to perform an additional evaluation of the enginebipropellant-valve double-seal configuration. The Block I1 SPS test progr am consistedof 650 test firings that had a total firing tim e of 20 478 seconds. A sum mar y of th isprogram is contained in table 11.

TABLE II.- BLOCK II WSTF TEST SUMMARY

Test phaseand date

Phase INov. 15, 1966 oJune 7, 1967

Phase IlJuly 13, 1967 oApr. 3, 1968

Phase IIIMay 2, 1968 oApr. 30, 1969

Objectives

System evaluation and characterizationat nominal, off-limit, and malfunc-tion conditions)emonstrate system operation forCSM 101 (Apollo 7)mission dutycycle

Demonstrate lunar-mission perf orm-ance at nominal, off-limit, and mal-function conditionsEvaluate ball valve double-seal ( 1 4 )configurationEvaluate PUGS perfor manc e at nominal

Demonstrate sy stem operation forand off-nominal conditions

CSM 101 (Apollo 7)

[mrestigate SPS operation under ex-treme off-limit conditions and multi-ple malfunctionsPerfo rm additional lunar andSC 101mission simulationshraluate ball valve (1-E) onfigurationInvestigate PUGS flight ano malie s

Testaeries

I

11IIINVVIVIIVIII-IIm.NVVIVII-IImnr-

!lumberof tes tfirings3647305645365428

33215171817175626

%66

--

1750

5134

a152b850--

Firingtime,se c

825.51917.84805.47905.50869.63896.82

1016.09975.19

a7 212.051 193.691210.411016.901 207.751 210.62947.45553.03

a7 339.851497.60909.07

1569.351 950.72

a5 926.74b20 478.64

Test conditionsrempfrature,F

7040 o 70

7050 o 10050 o 7070 o 10045 o 7070 o 100

707070

70 o 10045 to 70

7070

7070 o 10535 to 7040 o 70

Oxidizer andfuel ullageJressure, psia175175175

175 o 205179 o 187

179100 o 179

179~

179173197

170 o 190179

70 o 179142 o 179

68 o 179179

165 o 179179

aTest-phase total.bTest-program total. . .


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I ntegrated Systems TestsThe SPS subsys tem demonstra tion te s t s were conducted on SC 001, which includeda flight-type S M , at the WSTF fr om Febr uary 5, 1965, to September 7, 1966. The SPSthat was installed i n SC 001 was ident ica l functionally to the Block I flight s yst em s ex-cept for minor modifications that were required for ground testing and a more detailed

instrumentation sys tem. The SC 001 SPS was updated as requi red to maintain theBlock I configuration.A special tes t w a s initiated to investigate the flight anomaly that was observedduring the SC 003 flight. Before thi s te st , the oxidizer sump-tank standpipe was modi-fied to simulate the failed tr ansf er line i n SC 009. The res ul ts of this t es t led to astructura l improvement in the prope llant -transfer lines. A spacec raft 011 (SC 011)mission duty cycle was conducted to evaluate sy stem c haract erist ics that res ulted f ro mthe reworked transfer lines.System operational characteri sti cs of the SPS during nominal and off -limit condi-tions were demonstrated successfully by SC 001 te st s. These te st s included firings thatranged in duration fr om 23 to 600 seconds ; dual- and single-engine bipropellant-valveoperation; engine fir ings at minimum expected propellant and hardware temperature s;propellant-depletion fir ings; engine opera tion with a high engine-valve actuation pres-su re (190 psig); engine st ar ts with propellant only in the retention res er vo ir s; rapidre st ar t firings; chamber-pressure-decay firings at different propellant loads; engineoperation with and without postf ire in jector pur ges ; sys tem operation with tank ullagepre ssur es of 225 psia; simula ted fa ilures of individual engine-valve banks duringsteady-state operation; P U valve cycling during steady-state operation; firings duringwhich "zero" propellant imbalance w a s maintained by means of cycling the P U valve;system operation using fuel-cell power only; and sinusoidal and step gimbaling duringengine-start transients and steady-state operation.Before the flight of AS-201 (SC 009), the SM was static fired at the NASA John F.Kennedy Space Center (KSC) launch pad 16 in November 1965. Before the firing, theoxidizer sump-tank standpipe leaked propellant into the tra ns fe r line and storage tank.It was concluded tha t a maximum of 230 pounds of propellant would leak into the storagetank. A shift of th is magnitude would not cause a significant change in center-of-gravitylocation. The condition was waived because the engine opera ted sat isfactori ly duringthe test. This w a s the only Block I flight vehicle that wa s subjected to a static firing(see discussion of SC 009 flight).A static firing of the first Block 11SPS (spacecra ft 101) had been planned. To ex-pedite the launch schedule of spacecraft 101 (SC 101) by approximately 30 days, it was

decided to static fire spacecraft 102 (SC 102) ins tead, decontaminate it, and retur n itto the production cycle. This was considered acceptable and desirabl e because the twospacecraft were identical, and SC 101 would not be degraded because of propellant ex-posure. The general objective of the SC 102 test program was to verify that the Block I1SPS was ready or flight. To accomplish this, ce rt ain specific objectives wer e manda-tor y and we re completed successfully.

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F l i g h t - S i m u l a t i o n T e st sUnder simulated mission conditions, the spacecraft 008 (SC 008) SPS was evalu-ated during the thermal-vacuum-test program. The program contained provisions forverification that the subsystem could withstand the environments to which it would beexposed during the flight phase. The program was conducted in the Space Environment

Simulation Laboratory at the NASA Lyndon B. Johnson Space Center (JSC), fo rmer ly theManned Spacecraft Center (MSC), with simulated propellants . Data we re gathered todetermine the heat balance and the equilibrium-temperature distributions and to evalu-ate the ef fects of cyclic heating and cooling on the operation of the subsyst em.In August 1966, the fi r st two te st s were conducted but were invalid fo r the SPS.Flight-type h ea te rs on the engine bipropellant valve wer e not connected, and the he at er son the engine-gimbal-ring brackets were connected so that the primary and redundantheaters we re operated simultaneously. These discrepancies we re correc ted for thethird te st , and additional changes were made to bring the SPS configuration up to thespacecraft 012 (SC 012) configuration (Apollo 1). These changes consisted of insulatingthe sy stem feedlines, adding flight-type heat ers to the system feedlines, and placingtemper ature se ns or s in the locations that wer e planned for u se on SC 012.Similarly , in June 1968, spacecraft 2TV-1 was used to evaluate operation of theBlock II CSM in a simulated thermal-vacuum environment. Two te st objectives wereapplicable to the SPS. The first objective w a s to determine the SPS engine-temperature-decline ra te during a cold soak. The second objective was to establi sh the ability of theSPS engine the rm al control subsystem hea te rs to maintain temperatur es satisfactorilywhen exposed to a simulated side sun. The data fr om these te st s were used to updatethe computer th er ma l model, which was used to predict SPS engine tempera tures f orthe Block 11 missions. Subsequent flights resuted in verification that the SPS th er ma lcontrol system sati sfacto rily maintained cri tical points within the established tempera-

tur e limits.SUBSYSTEM FL I GHT RESULTS

Up to and including the lunar landing (Apollo l l ) , nine flights had involved the us eof the SPS. Four of the sy stem s were of the Block I configuration; five sy ste ms wereof the Block I1 configuration. All Block I flights were unmanned, and all Block 11 flightswer e manned.

B l o c k I F l i g h t sMission AS-201 (SC 009). - Two SPS firings were planned. Perfor mance was ac-ceptable for the first 70 seconds of the 180 seconds of fir ing that wer e scheduled for thefirst fir ing . At that time, engine-chamber pres su re began a gradua l decay and was ap-proximately 70 percent at SPS shutdown. The shutdown was ini tiated by the control-pr og ra me r backup command based on elapsed firing time. As a re su lt of the degradedperformance after the first 70 seconds, approximately 20 percent of the planned deltavelocity was not achieved. At shutdown, oxidizer-tank pr es su re unexpectedly dropped

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approximately 17 psi. Er ra ti c chamber pr es su re was noted during the st ar t of the sec-ond fir ing; stabil ized burning was not achieved during the planned 20-second burn. Theengine did appear to be recovering fr o m the effect s of helium ingestion before shutdown.Because the fuel-systems performance was either normal o r was as would be ex-pected as a res ult of the change in chamber pre ssu re , the initial malfunction analysiswas indicative that the probable malfunction modes were confined to the oxidizer sy s-tem. Also, it was concluded that helium ingestion into the dngine and two-phase flow(gas and liquid) between the sump and sto rage tanks had occur red . Scale-model testingof the oxidizer tank and retention re se rv oi r at the JSC and a full-scale test at the WSTF(SC 001) were indicative that the probable fa ilu re mode was a leak in the oxidizer-transf er line inside the retention re se rvo ir as discussed previously.Mission AS-202 (SC 011).- Mission AS-202 was the second flight test of the SPS.The primary test object ives were to verify the SPS standpipe fix by means of a minimumfiring of 198 seconds and to demonstrate multiple SPS re sta rts (at least three firings ofat least 3-second duration at 10-second intervals). Secondary tes t objectives were todetermine long-firing (approximately 200 seconds) SPS performance (including shutdowntransient cha racte ristics ) and to obtain data on SPS engine-firing stability. All objec-tives were met.Apollo 4 (SC 017). - The Apollo 4 mission included the thi rd flight tes t of the SPS.The primary SPS te st objectives were to demonstrate a satisfactory st ar t without anRCS settling maneuver and to determine SPS performance during a long-duration burn.Both objectives wer e met. The SPS operated nominally during both fir ing s. Duringpropellant c ros sove r, the engine-inlet pre ss ure and chamber p res su re increased aswas expected; the pre ss ur es wer e steady throughout the firing. The general effectsof propellant crossover were as expected.Apollo 6 (SC 020). - Essentially, the SPS mission objectives were the same as for

the Apollo 4 mission.' Because of the inabil ity t2 res tart the S-IVB stage, the SPS wasused to transfe r the CM fr om an earth-parking orbit to the highly elliptical earth-intersecting orbit that was needed to satisfy entry conditions fo r the CM heat shieldentry tes t. The object ives of sat isfactory SPS operat ion and a no-ullage start were ac-complished. In addition to these objectives, this firing was the fi r st in which the SPSdemonstrated the firing-duration capability that w a s needed to i ns er t the CSM into alunar orbit.Engine performance w a s satisf actory except for an overshoot in chamber pr es -su re during engine s ta rt . A l l other engine-transient cr it er ia were met. For theApollo 4 and 6 missions, the chamber-pressure transducer had been mounted on a2-inch adapter to reduce the th er mal effects that had caused an e rroneous indication

of chamber-pressure drift. The overshoot noted with the new adapter was significantlyhigher than with previous adapt ers . A specia l se r ie s of t es ts involving the us e of high-resolution instrumentation was scheduled a t the AEDC to determine i f the overshootwas caused partially by instrumentation error. It was determined that thrust-chamber-pre ss ure overshoots were reduced significantly if the engine firings were initiated usinga single bank of ball valves. It became standard operating procedure to st ar t eachfiring i n the single-bank mode. If the burn w a s scheduled to be longer than 6 seconds,the redundant bank was opened approximately 3 seconds afte r ignition.

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Block I I FlightsFive manned Apollo Block I1 missions were accomplished up to and including theApollo 11 lunar-landing mission. The SPS performance on these missions is discussedas follows.Apollo 7 mission. - The Apollo 7 mission was the first mission on which theBlock 11 SPS w as flight tested. As w a s planned, there wer e eight firings of the SPS en-gines. The four pri mar y objectives related to the SPS were minimum-impulse burn,perf ormance, primary /auxil iary propellant- gaging syst em, and th er ma l control. Allobjectives wer e met.Apollo 8 mission. - The Apollo 8 mission was in jeopardy after the firs t SPS burn,which was a midcourse correction. A momentary drop i n chamber pre ss ur e was ob-served at the start of the f i r st burn. This drop w as attribu ted to the presence of a gasbubble in the oxidizer feedline. The gas was determined to be trapped helium that re-sulted from a n inadequate engine-oxidizer bleed during preflight servicing. After ex-

tensive d iscussions and analysis by ground-based personnel, the decision was made tocontinue the mission. Two key it em s were important in the rea l-t ime decision to con-tinue this mission, which was the first attempt to leave ea rt h and orbit the moon witha manned spacecra ft. Telemetry data fr om the AS-201 miss ion, in which the enginealso had sustained s ome helium ingestion, and data from ground te st s, in which heliumingestion occurred , were compared with telemetry data to establish the signature tra cein chamber pres sur e of a sm al l amount of helium ingestion. Recordings of voice-tracktapes fr om the Apollo 8 loading operat ions at the KSC were indicative that the enginefeedlines had not been bled properly to remove the helium i n the high point of the line.With th is evidence, the cause of the. chamber- pres sure anomaly was established, andthe required confidence that the engine was not damaged was provided.During the early portion of t ranslunar coast, a drop of approximately 7 psi wasnoted in the SPS oxidizer-tank pr es su re . It is believed that the pr es su re de crease wascaused by helium absorption; the decrease stopped when the oxidizer apparently becamesaturated.The engine was sta rte d on all maneuvers by the us e of only one of two redundantse t s of valves in the engine bipropellant-valve assembly. This procedure was institutedto de crease the initial chamber -pr essure and thrust-level overshoot, which a r e charac.-ter is t ics of a start with both valve sets open. A noticeable d ecr eas e in the overshootmagnitude was achieved. During the lunar-orbit- insert ion (LOI) and trans ear th-injection maneuvers, the redundant-valve se t was opened approximately 3 seconds afte rignition to in cr ea se the operational reliability f o r the remainder of the firing in case one

of the valve sets should close.Apollo 9 mission.- The Apollo 9 miss ion involved both the CSM and LM . Thefifth SPS fir ing was made after a docked L M descent-engine firing of approximately372 seconds duration. Pref light analyses indicated that , when a descent-engine firingwas perf orme d with the spacec raft docked, a negative acceleration gr ea te r than20 .1 ft/sec would result . This accelera tion could cause depletion of the propellantcaptured by the SPS sump-tank retaining screens. Although the retention rese rvo irwould still re ma in full, some helium could be trapped and ingested into the engine

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during a subsequent SPS fir ing . However, af ter the docked descent-engine fi ring, allSPS firings were normal and smooth, indicat ing that no significant quantity of heliumhad been ingested.Apollo 10 mission. - One of the most significant changes in the Apollo 10 SPS wasthe addition of st ri p heat er s in the propellant-distribution lines that ra n fr om the tankoutlets to the bipropellant valves. The str ip heat er s provided a method of maintainingpropellant tempe ratures above the 30" F minimum allowance in a deep space environ-ment to prevent freezing of prope llants i n the feedlines. The SPS performance wassatisfac tory during each of the five maneuvers, and the total firing time was 545 se c-onds. The longest engine firing was the 356-second burn on the lunar-orbit- inse rtionmaneuver. The fourth and fifth SPS maneuvers, which occurred after depletion of thestorage tanks, were preceded by a plus-X RCS translation to sett le the propellants.A l l firings were conducted under automatic control.Apollo 11 mission.- The SPS performance was satisfactory during each of thefive maneuvers; the total firing time w a s 531.9 seconds. The longest engine fir ing w a s357.5 seconds during the lunar-orbit-inser tion maneuver. The fourth and fifth SPS

firings were preceded by a plus-X RCS translat ion to se tt le propellant, and all firingswere conducted under automatic control. The steady-sta te performance during allfirings was satisfactory. The steady-state pre ss ur e data wer e indicative of essentiallynominal performance; however, gaging-system data were indicative of a propellantflow ratio of 1.55: 1, ra th er than the expected propellant-ratio range of 1.60 : 1 o1.61: 1. The lower than expected mixture r at io decrea sed the amount of propellantavailable for velocity changes. One SPS anomaly occur red during the LO1 burn. Anabnormal pr es su re decay was noted in the secondary gaseous nitrogen (GN2 ) supply,indicating a leak. The decay ceased at engine shutdown, and no additional decay w a snoted. The leakage w a s attributed to contamination on the seat of the GN2 actuatorsolenoid valve. Fi lt er s were incorporated to prevent recu rr en ce of th is problem onsubsequent spacecraft.

CONCLUDING REM ARKSThe program that was developed to meet the se rv ice propulsion subsystem re-quirements lasted 92 months, f rom the time that the pr ime contractor w as given con-tr ac t go-ahead until the Apollo 11 manned lunar landing. During thi s period, nineflights were completed successfully, and all of the original requir emen ts we re met.The program necessitated the us e of a pri me contractor, 13 major subcontractors,approximately 500 vendors and suppl ier s, and two ma jor Government-owned te st

faci litie s .Although sever al significant events occur red during the development phase, thepoint at which the program began to reach maturity was on November 21, 1964, whenit w a s decided to initiate a Block I1 (lunar landing) s er vi ce propulsion subsystem con-figuration. By thi s time, most of the technical problems had been identified or at leastwere understood to the point that designs were available fo r incorporation into the sub-system. However, schedules f o r a re se ar ch program i n which the problems a r e un-known usually have to be modified as the prog ram proceeds. The se rv ic e propulsion

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system pro gram was no exception. During 1965, he most c rit ica l period of develop-ment occurr ed when component-fabrication changes reached a peak. Hardware modifi-cations, which are made in ord er to meet reliability requirements, not only delayed thecompletion of the qualification of each component, but necessitated a substantial amountof retest ing with the heavyweight subsystem test rig s to determine if interactions wouldoccur between major assemblies. The schedule also w a s delayed by assembly andcheckout of the first flight se rv ic e propulsion subsystem and its subsequent failure be-cause of the str uctura l collapse of the propellant-retention re servoi r. By 1966, thesubsystem configuration was completed and the program proceeded into the flight phase.

The most significant lesson that w a s learned fr om the s ervice propulsion subsys-tem program w a s the need to develop basic technology f or propulsion sy st em s beforeinitiating full-scale hardware designs. Besides th e anticipated technical problems,such as engine performance and combustion instability, schedule delays were experi-enced during hardware development, and these delays generally were associated withthe high reliabili ty requirements of the Apollo Program and the lack of exper ience withthe propellants and the ir effects on mater ials. Typical of these problems we repropellant-tank s t re s s corrosion, deterioration of seals in the tank doo rs and pre ssu ri-zation components, limited engine bipropellant-valve-seal cycle life, and nitrogentetroxide cor ros ion of aluminum par ts. Although so me studies had been conducted inselected areas, there was a lack of knowledge on the behavior of propellants in zerogravity, on the mechanism of propellant ignition, on the effects of freezing and therma ldecomposition of propellants, on ablative materials suitable fo r use in reliable thr ustchambers, and on data concerning the generation and effects of contamination.

Lyndon B. Johnson Space CenterNational A e ronautics and Space AdministrationHouston, Texas, March 27, 1973914-1 -00-00-72

Documents

Apollo Experience Report Service Propulsion Subsystem