Apollo Experience Report Lunar Module Reaction Control System

8/8/2019 Apollo Experience Report Lunar Module Reaction Control System

1/23

T E C H N I C A L N O T E NASA TN D-6740

-Chester A . Vuughun, Robert Villemarette,

itulij Kdrukulko, and Donald R. BlevinsSpacecrufi Center

77058

T I O N A L A E R O N A U T I C S A N D S PA CE 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 1 9 7 2


2/23

1. Report No. 2. Government Accession No.NASA 'I" D-67404. Title and SubtitleAPOLLO EXPERIENCE REPORTLUNAR MODULE REACTION CONTROL SYSTEM7. Author(s)Ches te r A. Vaughan, Robert Villemar ette,Wita lij Karakulko, and Donald R. Blevins, MSC9. Performing Organization Name and Address

Manned Spacecraft CenterHouston, Texas 7705812 . Spo nsoring Agency Name and AddressNational Aeronautics and Space AdministrationWashington, D. C. 20546

3. Rekipient's Catalog NO.

5. Report Date6. Performing Organization Code

March 1972

8. Performing Organization Report No,MSC S-315

1 19 . Security Classif. (o f this report) 20 . Security Classif. (of this page) 21 . No. of Pages 22 . PriceNone None 23 $3.00

10 . Work Unit No.914-1 -10-EP-72

11 . Contract or Grant No.

13. Type of Report and Period CoveredTechnical Note

14 . Sponsoring Agency Code

15. Supplementary Notes The MSC Director waived the use of the International System of Units (SI) orthis Apollo Experience Report, because, in his judgment, use of SI Units would impair the usefulnessof the report or result in excessive cost.16. Abstract

The design, development, and qualification of the reaction cont rol sys tem fo r the Apollo lunarmodule are descr ibed in this document. The lunar module react ion control sy ste m used manyof the components developed and qualified for the s ervi ce module react ion control sys tem . Thesyst em was qualified fo r manned flight during the unmanned Apollo 5 mission on January 22and 23, 1968, and has operated s atis facto rily during all manned lunar module flights includingApollo 11, the first manned landing on the moon.

17. Key Words (Suggested by Author(s))* Rocket Engines - Propellant Systems* Attitude Control - Propellant Pressurizat ion* Lunar Module * System Development- Apollo Experien ce * System CleanlinessReaction Control * Spacecraft Propulsion

~ ~~18. Distr ibution Statement

Fo r sa le by the Nation al Technical Informat ion Service, Spr ingf ie ld, Virg in ia 22151


3/23

CONTENTS

Section PageSUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1REQUIREMENTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2DESIGN PHILOSOPHY. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3DESIGN . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

5Installation of the System . . . . . . . . . . . . . . . . . . . . . . . . . . . 6

DEVELOPMENT AND CERTIFICATION . . . . . . . . . . . . . . . . . . . . . 678

Flange Heater and P re ss ur e Switch . . . . . . . . . . . . . . . . . . . . . .

Preproduction System Development Test Program . . . . . . . . . . . . . .Production System Development Te st Prog ram . . . . . . . . . . . . . . . .Design Verification Development Test Program . . . . . . . . . . . . . . . 9Production Clus ter Environmental Tes t Pro gram . . . . . . . . . . . . . . . 10Lunar Module Production Cluster Firing Test Program . . . . . . . . . . . 1 0Integrated RCS/APS PA-1 Tes t Pr og ra m . . . . . . . . . . . . . . . . . . . 11In-House LM RCS Te st Pr og ra m . . . . . . . . . . . . . . . . . . . . . . . 13Heater Integration Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14Engine Valve Tempe ratur e Tes ts . . . . . . . . . . . . . . . . . . . . . . . 15

15UNAR MODULE RCS CHECKOUT OF FLIGHT VEHICLES . . . . . . . . . . .LUNAR MODULE RCS FLIGHT PERFORMANCE . . . . . . . . . . . . . . . . 1 6Lunar Module 1 (Apollo 5 Mission) . . . . . . . . . . . . . . . . . . . . . . . 17

Lun ar Module 3 (Apollo 9 Mission) . . . . . . . . . . . . . . . . . . . . . . 18Lunar Module 4 (Apollo 10 Mission) . . . . . . . . . . . . . . . . . . . . . . 18Lunar Module 5 (Apollo 11 Mission) . . . . . . . . . . . . . . . . . . . . . . 1 9

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19iii


4/23

FIGURES

Figure Page. . . . . . . . . . . . . . . . . . . . . . . . . . . 2Apollo lunar module

345614

2 Reaction control sy stem installation . . . . . . . . . . . . . . . . . .Reaction control system schematic . . . . . . . . . . . . . . . . . . .

4 Propellant manifold and distribution syst em sc hemat ic . . . . . . . .5 Development and flight schedule . . . . . . . . . . . . . . . . . . . .6 Engine safe operating regimes . . . . . . . . . . . . . . . . . . . .

3

iv


5/23

APOLLO EXPERIENCE REPORTLUNAR MODULE RE ACTION CONTROL SYSTEMBy C h e s t e r A. V a ugh an , R obe rt V i l l em a r e t t e ,W i t a l i j K a r a k u lk o , a n d D o na ld R . B l e v i n sM a n n e d S p a ce c ra ft C e n t e r

S U M M A R YThe lunar module reaction control system was patterned ve ry closely after the

se rv ic e module reaction control syste m. Components common to the se rv ice modulereacti on control system and the lunar module reaction control sy ste m were used wherepossible. Where components could not be common, common technology was used in thedevelopment of the lunar module reaction control system. The exper ience gained fro mGemini miss ion s and the command and se rv ic e module reaction control sys te ms in theareas of sy st em fabrication, checkout, and testing also w a s applied to the lunar modulereaction control system. The system reliability requirements were achieved throughsy st em and component redundancy. Two independent operat ional luna r module reactioncontrol s ys te ms were provided.Th e development and certi ficat ion consisted of nine ma jo r ground test programs :

(1)preproduction system development, (2 ) production system development, (3 ) designverification development, (4) production cluster environment, (5) lunar module produc -tion clu ste r firing, (6) integrated reaction control system/ascent propulsion s yst emPA-1, (7 ) the NASA Manned Spacecraft Center in-house lunar module reaction controlsystem test, ( 8 ) heater integration, and (9) engine valve temperature.The checkout of the lunar module reaction control flight sy st em s wa s divided intofour basic categories of tests: (1)component, (2) module, (3) sys tem level, and (4 ) ve-hicle integration checkout.The perf ormance of the lunar module reaction contro l sys tem on Apollo missionswas satisfactory. Several minor probl ems occurred, but solutions were found for all

pro ble ms encountered.I NTRO DUCTI ON

The Apollo spacecraft is composed of the command module (CM), the servicemodule (SM), and the lunar module (LM). In Ju l y 1961, the NASA Space Task Groupreleased the first statement of work for the CM and the SM . Included in this statement


6/23

of work was a description of the reaction contro l syst em (RCS). The experience gainedduring the initial development of the CM RCS and the SM RCS was applied in the develop-ment of the LM RCS; and, as a resul t, the qualification prog ram was simplified greatly.This technical note is concerned only with the development and evolution of the LM RCS.

REQU IREMENTSThe Apollo missions required that the LM (fig. 1) maintain various attitudes withrespect to its flight path and maneuver in three axes to achieve a successful mannedlunar landing and re tu rn t o the command and se rv ice module (CSM). Specifically, theLM was required to be stable during all phases of flight and to have three axes of translation available for CSM separat ion, fo r CSM docking, and f o r var ious translat ionalmaneuvers during the lunar-orbit rendezvous. In addition, X-axis longitudinal tr an s-lation was required to provide propellant-settling th ru st f or the descent and ascent pro-pulsion systems .

Arcenl

Figure 1 . - Apollo lunar module.

A wide spectr um of operational re-quirements for vehicles that varied in massand moment of in er ti a by a fact or of 10 , cou-pled with the nece ssa ry vehicle acce ler a -tions, established a requirement for rocketengines capable of producing high sustainedthrust as well as low impulse. The initi aldesign criteria wer e to provide rocket en-gines of various t h r u s t levels to satisfy thevariety of requir emen ts. However, a closeexamination of the rocket engine that w a sbeing developed for the SM RCS revealedthat the high thrus t and low total-impulsecapabilities could satisfy all the transla-tional and rotational requirements of theLM mission. After this engine was select-ed , the rem aind er of the sy ste m was pat-terned very closely after the SM RCS; thedifferences were primarily in the propellant-load capability and sys tem geometry as dictated by the requirements of the LM mission.The similarity was extended to the use of common components wher ever possible. Be-cau se of the common-use philosophy and the common-technology approach, componentdevelopment and qualification testi ng fo r the LM RCS wer e simplified greatly.

The environmental constraints for the LM RCS generally were less severe thanthose of the SM RCS; therefore, the experience gained with the SM components in theareas of vibration, shock, the rmal vacuum, compatibility with propellant, and suscep-tibility to contamination could be applied di rect ly to the LM design. Two speci fic areasin which environmental conditions differed significantly were the vibration and the coldsoaking of the four LM engine clus te rs . Also, because LM propellant tanks wer e con-siderably longer than the SM tanks and the helium tanks we re lar ger in diamet er, thevibration test experience with the SM tanks could not be applied directly to the LMhardware. In these instances, the components wer e subjected to environmental testingdictated specifically by the LM environments.

2


7/23

DES GN PHILOSOPHYTo en su re reli abl e sy st em performance, the design of the LM RCS was based onsystem and component redundancy - imi la r t o the Mercury and Gemini spacecraftand to the Apollo CSM. Two independent and operationally ident ical LM RCS sy st em s,

each capable of providing attitude contro l and positive and negative longitudinal tr an s-lation, were provided (fig. 2). The RCS propellant supply (nitrogen tetroxide oxidizerand Aerozine-50 fuel) consisted of predetermined quantities for the lunar descent andascent maneuvers . The tankage of eac h sys tem was sized to contain one-half the RCSpropellant required fo r descent, plus the total RCS propellant required f or ascent. Inaddition, a contingency propellant supply was provided through an interconnect arrange-ment between the ascent propulsion system (APS) propellant tanks and the RCS mani-folds. The interconnect arran gemen t originally was meant t o be used only in anemerg ency situation. However, the interconnect arrangement was used as a normaloperating mode during the powered-ascent phase to conserve RCS propellants for dock-ing contingencies.Within each pressu rizat ion syste m, redundancy wa s used fo r such components asregulators , check valves, and explosive pres surization valves. The explosive valveswere in a parallel configuration because the primary failure mode wa s in a closed posi-tion. The regul ators wer e in a se ri es arrangement because the prim ary fa ilure modew a s in an open position. The check valves were arranged in a ser ies -pa ral lel configu-

Helium pressureregulating package

\Thrust chamberisolation valves

L R C S manifold

propellant transfe r -+ycrossfeed andascent-enginevalves - X

Axes orientationystem A1 System BOx. = oxidizer

Figure 2. - Reaction control systeminstallation.

ration (1) because the failure probability inan open o r closed position was consideredto be about equal and (2) because the weightpenalty associated with this part icula r com-ponent was minimal. Thus, no single func-tional failure could impair the control ofthe spacecraft or jeopardize crew safetybecause of propellant shortage.

The common-use philosophy was ap-plied throughout the system design, usingdeveloped CSM components wherever pos-sible. Whenever the component could notbe used directly, but could be made usableon the LM with minor modification, acommon-technology approach was followed.The manufactu rer of the SM pa rt was givent h e task of modifying hi s product t o make itusable on the LM. Because thi s approachpermitted the use of the sa me test proce-dures, test equipment, and personnel em-ployed fo r the SM, the r es ou rc es andlearning tim e required to produce a givenpiece of hardware were minimized and,t h u s , significant cost savings and increasedconfidence in reliability resulted.

3


8/23

DES IGNThe LM RCS consisted of two independent sys tems , A and B. Each system pro-vided the vehicle with attitude control and X-axis translation when used independently.When used together, Z - and Y- a x i s trans lation could be obtained also. The two sys temswere identical in all re sp ec ts other than the engine locations and thrus t vecto rs.Each system had an independent helium-pressuri zation module, propellant tanks,

and propellant manifold (fig. 3 ) . The helium-pressurization module consisted of thehelium storage tank, two parallel initiating explosive valves, a filter that protected theregulator, an orifice that acted as a damping device, a pressure regulator that reducedthe initial storage pressure of 3000 psia to an operating level of 180 psia, a check valveassembly for each propellant (oxidizer and fuel) a n k , and a relief valve assembly fo reach tank. The function of the check valves w as to prevent backflow of fuel and oxi-di ze r vapors, which later could condense in a common area and r ea ct to cause a localpressure and temperature rise. The function of the relief valves was to protect thepropellant tanks from overpressurization in case of an increase in propellant tempera-ture o r a regulator malfunction. Each syst em included a number of servicing andcheckout tes t ports .

System B System A(PQMDI sensor (2 )q:ropellant-quantity-measuring device Helium pressure vessel 121

p -Pressure transducer (81-Helium initiating valve (4 )

Note:Valve positions (N.O., -Helium filter (21

Ip m n Iwitch (161-Figure 3 . - Reaction control system schematic.

4


9/23

The propellant t anks were of the positive -expulsion configuration. Each propel-lant w a s contained inside a Teflon bladder that was in turn placed inside a titaniumshell . Helium ga s entered the ar ea between the titanium and the Teflon, forcing thepropellant out of the bladder and into a perforated standpipe connected to the tank outletport.The propellant manifold and distribution system is shown schematically in fig-

ure 4. The position of the main, interconnect, fuel, and oxidizer valves controlledthe propellant distribution to the engines.The valves were developed specificallyfor the LM RCS applications because nocomponent was available that could sat-isfy the requirements of every location.Each valve was a latching solenoid type,which required power only for openingand closing. After the des ire d position * y s t e m Awas established, the poppet w a s held inposition by permanent magnets. Thesystem conditions but were underdesignedfo r dynamic conditions; theref ore, thecrew men had to verify cor re ct valve posi-tions during critical phases of flight. Withbe used for any dynamic environment with-out creating a major syste m problem.

Engine isolation

valves wer e w e l l suited for nominal, static ;:iii:r stem B oxidizerfuel tankOxidizer crossfeedvalverop er cr ew procedures, the valves could vaIve

Figure 4. - Propellant manifold and distri-bution system schematic.In-line propellant f il te rs originallywere located downst ream of the isolationvalves. However, because the isolation valves were part icul arl y sensitive to contami-nation (as discu sse d in detail later), the fi lt er s were relocated upstream as shown infigure 3 .

A redundant set of interconnect valves was provided. The valves ensured that,upon completion of the APS-interconnect operation during lunar ascent, the RCS pro-pellants would not be tra ns fe rr ed to the empty APS tanks and that pr es su ra nt ga s fro mthe APS tanks would not be ingested into the RCS. Thes e interconnect valves were nor -mally in the open position, and they were closed only if a malfunction was detected du r-ing the nominal APS-interconnect termination procedure.

F la n g e H e a te r an d P r e s s u r e S w i t c hTwo components developed specifically for use with the LM RCS engines were theengine flange hea ter and the engine chambe r pre ss ur e switch.Engine flange heaters. - The engine flange heaters were required to maintain thete mp er at ur e of the engine combustion-chamber flange above 120" F. This temperaturelevel, which w a s determined during the engine-requalification program described later,was required to ensure safe operation of the engines with the Aerozine-50 fuel duringall phases of the mission. Two heat ers pe r engine we re provided to ens ure redundancy.

5


10/23

Both heate rs we re operated by the automatic system , and one heat er had a manualon/off override.

1966 1967

111111 ,,11

111 II11111II

11111111111

11111 111111111111111

Engine chamber pr es su re switch. - The engine pre ss ur e switch w a s used in con-junction with the failure-detection sy st em of th e LM. The electrical signal fro m theguidance, navigation, and contro l sys tem to the engine valves was compared electron-ically with the output of the p re ss ure switch; and, if the two did not match, the enginefai lur e indication was displayed to the crew. Corre ctive action such as engine iso la-tion o r troubleshooting in other sy st em s then could be perf orme d by the c rew.

1968

111

11111'IllI1111I

Ins ta l la t ion of t h e Sy st emThe modular grouping of the propellant and helium storage components was madef o r the purpose of simplifying the checkout and re pa ir procedures. Both operationscould be perfo rmed "on th e bench" without interfering with the ove rall vehicle op era-tions. The sys tem was installed in two bay areas and on four outr igger booms. Thetankage modules (helium, fuel, and oxidizer) were ins talled on the left- and right-hand

s ides of the LM directly above the APS tanks. The engines wer e installed in cl us te rs offour on the outriggers, which were located around the periphery of the ascent stage at4 5 " t o the orthogonal (pitch and roll) axes. Two of the four engines in each clu ste rwe re fed from eac h propellant supply.

DEVELOPMENT AND CERT l F I CAT1 ONThe development and certification of the LM RCS consisted of nine major groundtest programs. An overall RCS development schedule that includes the time phasing ofthe various programs and of the first six flights is shown in figure 5 . A brief discus-

sion of each major test prog ram follows.Taskor flight I 1964

~

iR-3 breadboardiR -3 Droduct ion svstem. I R -~besign ver i f ib t io n testi R - li R 2upplemental engine qualificatioi'A-1 (series 21'A-l lseries 81'A- l (ser ies 11and 121n-house LM R C S test p r q r a m'roduction cluster:nginelheater integration:ngine valve high temperature\erozine-50 margin tests: m e n propellant testsW l cluster thermal testsLM-3 IApollo 91LM-5tApollo 111

LM-1 IApOllO 51LM-4 IApollo 101LM-6 IApOllO 121LM-7 IApollo 131

11111111111965

11111969

I

I

I' I I

1970

Figure 5. - Development and flight schedule.6


11/23

Pr e p r o d u c t i o n Sy s te m D e v el op me n t T es t P r o g r a mThe preproduction sys tem development test prog ram , o r breadboard test, wasthe first test in which the proposed configuration of the LM RCS wa s hot fir ed. Thet e r m "configuration " is emphasized here because the geometrical configuration of thepropellant-feed plumbing, which was the most important tes t item, w a s about the onlyaspect of the test hardw are that resembled actual flight hardware. With the exceptionof the ve ry e ar ly prototype engines, the breadboard s ys te m was composed enti rely ofcomme rcial ly available industrial-type components.In early 1964, when the breadboard wa s being assembled, most flight componentswere still in early stages of development and were ra ther s ca rc e. Most of the compo-nents were being developed under a common-usage agreement, and those few partsmade available to the LM pr ogr am we re being installed on the LM-1 vehicle and onmajor test vehicles.The breadboard testing was conducted by the RCS engine developer between Au-

gust 1964 and May 1965. One of the primary objectives of the test program was to in-vestigate the dynamic ch ar ac te ri st ic s of the propellant -feed sys tem . The maximumpulse-frequency req uir eme nt fo r the engines was 25 pulses/sec at a pulse duration of10 milliseconds. This high pulse frequency caused concern tha t the res po nse of therela tivel y long propellant f eedlines of the LM system was too slow to maintain adequatep ressu res at the engine inlet s during tran sient flow. Other objectives wer e to evaluatepropellant -manifold p riming procedures and engine per for mance during multienginefir ing s. Another importan t aspec t of the program was to "shake down" the new LMtest facil ity and data-acquisition equipment, which had been instal led specifi cally fo rL M RCS development testing.The firing progr am consisted of single and multiengine f irin g mat ri ce s coveringa wide rang e of pulse widths, pulse frequencies, and predicted flight duty cycle s. Theeffects of pulse firin gs on the steady-state performance of another engine in the samesys tem also wer e investigated. Almost 15 000 separ ate fir ings totaling approximately8000 seconds of burn time were accumulated on the eight-engine system.Although the breadboard configuration hardly r ese mbl ed a flight system, it pro-vided valuable dat a much earlier than would have been possible if the test program hadbeen dependent on the availability of flight-type hardware. The breadboard test pro-gra m was completed almost a ye ar before the first flight-type configuration was avail-able for testing.One of the m os t significant findings of the tes t pro gram wa s that the feed-pr es su re fluctuations during the short-pulse high-frequency firing were mo re se ve rethan had been predicted analytically. In certain pulse modes, the transient engine-inlet pr es su re s dropped to levels as low as the propellant vapor pr es su re , in whichca se no thru st w as produced by the engine. On the other hand, peak tran sient p re ss ur egenerated at engine valve clo sure approached the proof -p re ss ur e req uir eme nts of som ecomponents.Fortunately, the detriment al effects of these se ve re fee d-p res sur e fluctuationswer e recognized in tim e to avoid s er io us effect to the LM progr am. Thi s informationled to a complete reevaluation of the control system r equi reme nts for the LM and

7


12/23

helped to define the interface between the guidance sys tem and the RCS. The signifi-cant result of the reevaluation was merely a change of the maximum pulse frequencyfrom 25 to 7 pulses/sec.Another significant conclusion fr om the breadboard test ing was that the plannedtechnique f o r filling o r priming the propellant manifolds resul ted in excessiv e trans ien tpressures, in some cases greater than the design burs t pr essure fo r some flight com-

ponents. The bur st- pre ssu re requir ement for most propellant-feed sys tem componentsw a s 550 psia. During priming, with nominal operating tank pressures, transient pres-s u r e s of almost 1100 psi were recorded. Thes e dat a result ed in modification of theplanned flight-activation procedure fro m priming with full sys tem operating p res su reto priming at tank pad pres su re, before activating the helium-pressurization syste m.

P r od uc t i on S y s tem D ev e l opm en t T es t P r o g r amThe second major test program, the production system development test pro-gram, w a s conducted from August to November 1966, during the development of theLM RCS. Thi s program was pe rhaps the most significant in the LM RCS development.The test ri g was designated HR - 3 P . With the exception of addi tional ins trumen-tation and minor modifications necessary to facilitate ground test operations, the con-figuration of the HR-3P test r ig w a s almost identical to that of the RCS on LM-1, thefirst flight vehicle.The basic objective of the test program was to determine if the system couldachieve fundamental design requ ire ment s. Most of the components in the test r ig wereprequalified or qualified models that had undergone extensive development tests as in-dividual components. However, HR - 3 P was the first ri g that was assembled and testedas a complete system .To facilitate efficient test operations, the progr am w a s divided into a series ofshort tests, each with a specific objective relating t o the various environmental o roperational conditions that a flight sys tem might experience during a lunar landing.Also, anticipated flight conditions we re s imulated whenever possible within the opera-tional limitations of the test facility and the imposed schedule requir emen ts. All testswere performed at local baromet ric pres sur e, which was approximately 12.5 psia. Alarge number of specific test objectives included the investigation of the following items.1. Helium -pressurization -syst em activation and propellant sect ion priming2. System performance during simulated vehicle control modes and missionduty cycles3 . The effects of high and low temperatures on the performance of the system4 . System performance during cros sfeed and simulated interconnect operation5. Effectiveness of system-malfunction procedu res and component redundancy

8


13/23

6. Component compatibility with propellants7. Decontamination techniques and fluidsThe HR-3P test rig also provided valuable experience in helium and propellantservicing. Thi s experience w a s used in the design of the s pace cra ft ground-support

equipment used at the launch site.The firing program consisted of single and multiengine firings under nominaland off -nominal conditions of propellant-feed pr es sur es , engine valve voltages, andpropellant tem per atu res . Approximately 57 000 separate engine firings f or a total ofapproximately 8300 seconds of burn time wer e accumulated on the syst em engines dur -ing the pr ogram.Except f or the r at he r routine operational problems with instrumentation and sup-port facilities that are usually expected during a complex and large-scale test opera-tion, the progr am ran ra th er smoothly and was completed within a reasonable periodof time. All test objectives were accomplished; that is, the capability of the systemdesign to me et fundamental re qui rem ent s was demonstrated and no insurmountable de -ficiencies wer e uncovered. The unique environment resulting from sys te m operationdid discl ose salie nt ch ar ac ter is ti cs of some components that wer e not compatible with

all planned system operational modes. An outstanding example w a s the discovery thatthe propellant latching valves would unlatch and shift position when subjected t o the highflow rates o r pr es su re sur ges that occurred during initial filling o r priming of the pro-pellant manifolds. Th is problem led to a very comprehensive investigation into the de-sign char act er is ti cs of the latching valve and revealed the true limitations of the valve.As mentioned previously, thi s valve problem was solved for flight by requiri ng the cre wto ascert ain c or re ct valve positions during criti cal phase s of flight. Other componentproble ms discovered wer e (1) ransducer diaphragm incompatibility with propellantcombustion r es id ua ls and (2) an inadequate seal design in the ground half of thepropellant -servicing quick-disconnect couplings.

The te st p rog ram also substantiated the contention that t he contamination contr olreq uir emen ts and proced ure s fo r the s yst em were incompatible with the designs andreliability re qu iremen ts of some components. Almost every sys tem component andmany facility components experienced leakage failu res caused by particulate contami-nation. The propellant latching valve, because of a very narrow (0.006 inch) seat, wasfound to be particu larly sensi tive to contamination. The large number of leakage f a i l -ures caused by particulate contamination provided support for a broad-based contami-nation control program, which is discussed in the section of th is r ep or t entitled "LunarModule RCS Checkout of Flight Vehicles."

D es i gn V e r i i c a t o n D ev e l opm en t T es t P r og r amThe third system-level test pro gra m was a design verification test (DVT) exe-cuted during February and March 1967. Whereas the objective of the previous produc-tion system test pro gra m was to determine the performance of the sy ste m under vari ousconditions, the objective of the DVT program was considerably br oade r in scope. Notonly was acceptable operation of the system demonstrated, bu t other factors such as

9


14/23

manufacturing and checkout procedures , contamination contro l techniques, and propel-lant decontamination pr oce dures used on flight sy st em s wer e verified. All componentsin the DVT sys tem were fully qualified models and were assembled into the sa me con-figuration as that of LM-3, the first manned vehicle. Also, the DVT sy st em underwentthe same manufacturing and checkout operations as the flight systems.The DVT program was essenti ally a repet ition of some pa r ts of the production

syst em test progra m. High and low propel lant- temperature tests, crossfeed opera-tion, and simulated failure-mode operation tests we re conducted. Approximately23 200 engine fi ri ngs wer e made, totaling approximately 3800 seconds of engine-burntime.

Production Cluster Environmental Test ProgramThe production cluster environmental test program demonstrated the structuralintegrity of the engine cluster and vehicle mounting hardware. Extensive shock andvibration tes ts in al l maj or a xes wer e conducted on a complete production flight-typeengine clust er and boom assembly.With the exception of the fai lure of a chamber pres sur e transducer bracket, thecluster design withstood all the mission-level random and sinusoidal vibration loads towhich it w a s subjected. Ov er st re ss vibration leve ls of up to 200 percen t of spec ifica -tion requirements also were imposed on the cluster without causing any significantstructural failures. Failure of the transducer bracket resulted in a bracket redesignthat was retrofitted on LM-1, the first flight vehicle.

Lunar Module Production Cluster Firing Test ProgramThe firings of a complete, flightworthy LM engine cluster (four engines) undersimulated altitude conditions took place during the production cluster firing test pro-gram that w a s conducted in April 1966 . The prim ar y objectives of the prog ram wer eto evaluate engine performanc e under mor e r eali sti c flight conditions (particul arly lowambient press ure ) and to determine the heat-transfer char act eri sti cs of the clusterduring steady-state and pulse-mode duty cycles. The the rmal data were to be inte-grated into the cluster thermal tests that we re being conducted in a the rmal vacuumfacility that simulated the space environment.The firi ng pro gra m consisted of single and multiengine fi rings that simulatedselec ted portions of expected mission duty cycl es. In all, 1807 seconds of burn timeand approximately 8500 firi ngs wer e accumulated on the four engines in the clus ter.On April 27, 1966, during the low-temper ature mission duty cycle par t of theprogram, the combustion chamber of the upfiring engine w a s destroyed by an explosionthat occurred during the start of a pulse. Thi s failur e resulted in the immediate ter-mination of the program and the initiation of an extensive failure-analysis effort.The analysis revealed that the f ailur e was caused by a combination of conditionsra th er than by one single cause. The upfiring attitude of the engine, low engine te m-peratu res, helium saturation of the propellants, short-pulse firi ngs, and relatively

10


15/23

high test-cell ambient pr es sur e wer e some factors found to contribute to the accumula-tion of ni tra te compounds, which could caus e high ignition ove rp re ss ur es and injector -manifold explosions.The engine fa ilu re brought about an extensive engine -requalification program butdid not resu lt in any engine design changes. For the CSM application, the engine was

qualified with monomethylhydrazine fuel. For the LM application, the requal ifica tionwas done with Aerozine-50 because of the requirement for u se of the sa me fuel for theLM RCS and APS. When the engine flange temperature was maintained above 120" F,engine failu re was unlikely. The clust er heater design was changed to ensure 120" Fflange temperature.

I n t e g r a t e d R C S l A P S P A - 1 T es t P r o g r a mThe PA-1 test rig was a flight-weight ascent-stage structure with only the RCS

and APS installed. It was built principal ly fo r development tes ting of the APS, and theRCS was included pri mar ily fo r evaluating the interconnect-propellant-feed mode.The configuration of the PA-1 RCS was simil ar to that of test r i g HR-3P and flightvehicle LM-1 except that solenoid valves we re installed between the propellant tanksand the helium module to prevent migration of propellant vapors into the regulato rs andcheck valves during extended downtimes. Also, additional feedline acc es s port s wereinstalled between the cluster isolation valves and the engines to aid in drainingpropellant.All testin g was conducted at the NASA White Sands Test Facili ty (WSTF) duringSeptember and October 1966. A series of 11 runs was made under the altitude condi-tions (88 000 to 140 000 feet) attainable with the WSTF vacuum pumping system. Ap-

proximately 3000 firings and 515 seconds of f iring time were accumulated on 12 engines;the fou r upfiring engines were disabled t o prevent possible injector-manifold explosions,which could occur at the relatively high test-cell pres sur e. The high test- cell pr es su rewas not a realistic simulation of space vacuum.Analysis of the da ta indicated that neither the RCS nor the APS experienced anydetrimental effects during the interconnect-feed operation. A very minor decrease inRCS engine thrus t, estim ated to be 1 o 2 percent, was observed while propellant wasbeing supplied fr om the APS. Thi s condition was attributed to slightly lower engine-inlet pr es su re s resulting from increased pressure los s through the longer feedlines.The propellant pressure transients generated during RCS engine pulsing were found tohave little o r no influence on the per formance of the APS engine.As in earlier system test programs, the lack of adequate contamination controlwas the only system-oriented problem that arose during the testing. Almost all thepropellant latching valves in the sys tem experienced inter nal leakage; and, before thetesti ng began, all engines had to be returned t o the manufacturer f or cleaning.A demonstration te st of a proposed flight procedure f o r venting the RCS propellantmanifolds of the nitrogen pad p re ss ur e by opening the engine valves at high altitude wasunsuccessful. Gr os s propellant leakage through either the interconnect valves or themain shutoff valves resul ted in hot firings during the attempted manifold venting. As a

11


16/23

result of t h i s test and numerous leakage failures of the latching valve, the LM-1 mani-fold venting operat ion was eliminated.The RCS testing on the PA-1 configuration was resumed with series 8, 11, and12 tests, start ing in July 1968 and ending in May 1969. The maj or emphas is duringPA-1 series 8 testing was to test the APS and RCS in support of the first manned flight(LM-3). The RCS engines wer e operated in normal-feed, cross feed, and interconnected-

feed modes. The syst ems installed in the PA-1 ri g wer e essentially the LM-3 configu-ration. All tests we re conducted at simulated altitude conditions. The test-cellpressure w a s maintained below the 0.2-psia red-l ine value based on engine test expe-rience; therefore, the upfiring engines could be fir ed. Testing consis ted of base-lineengine performance tests with var ious propellant-feed modes, high-altitude start test-ing, and selected firing ma tr ices designed to evaluate the integration of the RCS and thecaution and warning elec tro nic s assembly (CWEA).The prim ary conclusions from series 8 tests were as follows.1. The RCS performance in the vario us feed modes was acceptable.2. Safe RCS start capability and acceptable engine performance at high altitude(220 000 feet) were demonstrated.3. The NASA John F. Kennedy Space Center servicing procedures should bemodified to include a continuous powering (open) of the main shutoff val ves dur ingpriming of the RCS manifolds. All other priming and pressurizat ion procedures werefound to be acceptable.4. Hydraulic interactions between the APS and RCS during interconnect feedcaused minor fluctuations in the ascent-engine chamber p re ss ur e.5. The integrated RCS/CWEA per formed sat isfactor ily.In series 11, RCS engines wer e operated in the interconnect-feed mode (APS pro -pellants) throughout the series to simulate the a scent portion of the l unar missi on.Testi ng consisted of a shakedown firing, an off -nominal lunar-landing-mission dutycycle to evaluate the extent of RCS-induced fluctuations on the ascent-engine chamb erpressu re, and subsequent heat-soakback t es ts to simulate vari ous failed RCS engineconfigurations.Series 11 testing indicated a potential problem with pres su re rise of trapped pro-pellant in the inl et manifolds as a res ult of th erm al soakback fro m a hot engine. Inthese tests, thr ust er- pair isolation valves were closed and the RCS heat er was turnedoff after RCS firing activity. In several cases, engine heat soakback and the concomi-tant thermal expansion of the trapped propellants resul ted in inlet pr es sur es up t o themaximum allowed (700 psia in this test). The rate of pressure buildup showed that700 ps ia would have been exceeded had the thr us te r-pa ir isolation valves not beenopened for relief; consequently, the Apollo malfunction procedures incorporated pres-su re relief steps- ir ing of one of the two isolated engines aft er isolation of an enginepair.

1 2


17/23

Ser ies 12 testing consisted basically of a shakedown firing, a simulated ullageburn (in support of the LM-3 APS anomaly investigations), an off-nominal lunar-landing-missio n duty cycle (high-frequency RCS pulsing), and subsequent heat-soakback tests.It was concluded fro m series 12 testi ng that high-frequency pulsing (up to 11 pulses/sec)of the RCS does not initiate CWEA thrust chamber assembly (TCA) failure indicationsin eit her the nor mal - o r interconnect-feed modes. Fur the rmo re, high-frequency puls-ing of the RCS doe s not seriou sly degra de engine performance, although the effect isgr ea te r in the interconnect mode than in the normal-feed modes.

I n-House LM R C S Test ProgramA complete LM RCS test was conducted at the NASA Manned Spacecraft Center(MSC) in Dec emb er 1967. The prim ar y objectives of the tes t were to define the gene raloperation al cha ra ct er is ti cs of the LM RCS under simulated altitude conditions and to

obtain per for man ce da ta on individual subsystem components. This syst em te st wasthe first t o be conducted at simulated altitude conditions in ex ces s of 100 000 feet.The test art icl e included all qualified components except the combustion-chamberpr es s ur e switches. Most sys tem components and all propellant lines had been usedpreviously in tests on the HR-3 DVT system at the subcontractor's Magic MountainTest Facility. The HR-3 DVT configuration w as modified and updated as required tosatisfy specific test objectives and to incorporate the latest changes to the flight sys-tem . The most important of these changes were as follows.1. Propellant-quantity-measuring devices were installed in each helium tank.2. One flight-type th ru st er heater w a s installed on ea ch engine.3. A propellant filter was installed in each engine in jector valve.4. A pre ss ur e switch (not flight configuration) was installed in ea ch of 16 engineinjector heads.5. Flight-type arc -su ppr ess ion circuitry was installed on eac h engine.The test program included (1)pretest operations, (2) base-line performance dutycycles, (3) simulated LM-1 and lunar-mission duty cycle s, (4) spec ial duty cycle s de-signed to accomplish specific test objectives and evaluate s yst em pe rform ance whensubjected to worst-case" duty cycles, and (5) post-test checkout and decontamination.The prima ry objectives of the test were satisfied ; that is , dat a on the generaloper ation al ch ar ac te ri st ic s of the LM RCS and of the individual components were ob-tained. In the test prog ram, th re e types of anomalies we re observed. These anoma-lies, which were investigated and resolved, were propellant latch valve leakages,p re ss ur e switch failu res, and engine injector cooling below the 120" F lower limit.The propellant latch valve leakage was found to be caused by particulate contam-ination; there fore , the need f o r sys tem cleanliness w a s emphasized . The switch fail-ure s were of two types- ailed closed and failed open. The failed-c losed conditionwas traced to the contamination of the switch mechanism by the semiliquid combustion

13


18/23

products. Such a fa il ure would simply eliminate the use fulness of the switch but wouldnot lead to a ser iou s problem. The open failure w a s caused by a design deficiency thatwa s corrected on the flight configuration. The injector cooling problem was tra ced t othe engine duty cycle. The probl em was resolved during later tests when it w a s shownthat the normal miss ion duty cycl es would not produce the cooling effect observed.Another significant finding was that, in general, the HR-3 DVT system compo-nents performed within specification limit s aft er tes ting at the subcontractor's plant

and storage fo r s ever al months at MSC. Also, following the test program, the systemperformance was adequate to complete a subsequent test program (LM-1 nomaly inves-tigation) after a 4.5-month exposure to an unknown and uncontrolled concentrat ion ofpropellants .H ea t e r I n t e g r a t i o n Tests

After the engine was requalified with Aerozine-50 fuel at 120" F flange tempera-ture, a complete clus te r of fo ur engines and eight engine flange heat er s was subjectedto numerous fi ring duty cycles. In the test progr am, cer tai n combinations of sho rtpulse s caused the engine flange to cool fa st er than the he ate rs could war m it. In somecases, the cooling effect was so se ve re that the flange temper atu re dropped to below100" F, and an engine explosion finally resulted. Additional test run s were performedt o identify the flange cooling regimes . It was recognized that many of the duty cyclestested were much more seve re than the duty cyc le expected in flight. However, be-cause the actual duty cycle in flight is highly unpredictable, it was necessary t o estab-li sh the safe operating regime. The resul t was a map defining safe and unsafe engineoperat ing duty cycles (fig. 6). Concur rent with the engine tests, additional missionsimulations were performed to obtain a bette r est ima te of the flight duty cycle. Ascan be seen in figure 6, the miss ion operat ing envelope (automatic mode) was found tobe well within the safe region. The cooling effec t was not a problem in the manualcommand mode because short-pulse combinations were not possible.

F d i l pintY2M1

tllu -

80 -

Figure 6. - Engine safe operating regimes.1 4


19/23

Engine Valve Tem perature TestsDuring the LM-3 mission, the temperatur es recor ded on the engine cl us te rs ex-ceeded not only the predicted values but also the upper limit of the cluster instrumen-tation. A series of tests was conducted in April 1969 to define a maximum temperatu reto which the engine valves could be subjected without degradat ion of performance. The

engine was teste d at ever-increasing tempe ratur es starting at 275' F and ending at3 7 5 " F. The tests were terminated at 375" F when no degradation in performance wasexperienced. An instrumentation change (increase of upper limit to 260" from 200" F)also was made to accommodate the expected operating tempera ture of the clusters.LUNAR MODULE RCS CHECKOUT OF FLIGHT VEHICLES

The checkout of the LM RCS was divided into four basic categories of tests, whichwere conducted during the var iou s stages of vehicle buildup. The se we re componenttests, module tests, system-lev el tes ts, and vehicle-integration-checkout te st s. Ingeneral, all tests conducted on individual components at ea ch level of checkout wer ethe sam e. Also, most test and s ucc ess cri te ri a were patterned after the componentpredelivery acceptance tests, which were derived fro m the original procurem ent spec-ifications fo r the component. The various checkout tests performed in the co urse ofthe vehicle buildup provided increased confidence in the system and permitted the track-ing of component perfo rmanc e fro m one te st t o another so that any degradation could bedetected easily.

Component tests were conducted at two locations, at the plant of the componentmanufacturer and at the point of assembly j u s t before installation into the vehicle. Thepreinstallation test (PIT) was essentially a repetition of the checkout by the componentmanufacturer with some of the less important tests eliminated. A s the program pro-gressed and mo re exper ience and confidence were gained with the hardware, the scopeof the PIT w a s decre ased gradually; and, in some ca se s, the test was eliminatedentirely.The PIT was informal and consisted of pro of-pressure tests, external and inter-nal leak tests, clean liness verifications, and functional checks. Each test was con-ducted according to a test plan or test outline that did not ca ll out each individual st epof an operation, and these documents did not requi re the de gr ee of quality control andinspection required by the higher level test procedures. This lack of fo rm al testpro ced ure s was the most significant flaw in the PIT operation and resulted in the in-troduction of particul ate ma tt er into some components, rupturing of num ero us helium-

relief-valve bu rs t disks , and shorting of position indicator switches on the propellantlatching valve. However, the PIT did eliminate many faulty units before they could beinstalled on the vehicle.The second level of checkout -t he module level - as conducted on as sem bli esconsisting of two o r mor e components. Interconnecting braz ed joints and plumbingwe re proof and leak tested, and the stand ard checks on individual components we re re-peated. The module-checkout concept increased the efficiency of the checkout o per a-tions by substantially reducing the complexity of the pro ced ure s and the test equipment.

15


20/23

For example, the elaborate and time -consuming precautions required to prevent flexu reof the propellant tank bladders during regulator and relief valve checkout were com-pletely eliminated in the helium-module checkout because this checkout was performedwithout the tankage module attached. The module -test concept a ls o allowed flexibilityin the checkout flow; that is , the va riou s modules could be checked out independentlywithout constraining oth er checkout functions. Other ite ms checked as modules werecl us te r isolation -valve/f ilter ass emb lie s and the propellant manifolds. Tankage-modulecheckout consisted of bladder leak checks, unlatch cur re nt measu reme nt and leak testsof the main shutoff valves, and proof and leak tests of the br azed joints connecting thecompletely assembled and tested helium module to the tanks. Final instrumentationche cks produced a complete tankage module ready for installation on the vehicle.

Concurren tly with tankage -module checkout, the prope llant manifolds were in -stall ed and verified clean by a liquid Freon flush. This clea nline ss test was conductedin two stages. The first pa rt verif ied the plumbing between the ou tlet of the tankagemodule t o the inlet of the c lu st er filter, and the second par t verified the cl eanl ines s ofthe e nt ir e manifold up to the engine inlets. It should be pointed out that th is liquid-flush cleanliness verification was incorporate d only after a gaseous verification ap-proa ch proved to be totally inadequate. The effec tiveness of the flushing proc edurewa s enhanced fur the r by simultaneous low leve l vibration of the plumbing. Leakagefai lur es of engine valves and propellant latch valves wer e reduced d rasti cally by theliquid -flush techniques.

System-level tests consisted of manifold integrity tests, engine leak tests, andengine gas flow tests. Some helium-module components, which were susceptible t odegradation by other functions o r by time, also were retested. A rat her unique testtool called a I fcalibrated feather'' was used to verify, qualitatively, ga s flow througheac h individual propellant orifice in the engine injector head. Until th is device wa sdeveloped, no method w a s available to dete rmine if the small injector orifices wer eunobstructed.Vehicle integration tests, usually conducted just before vehicle shipment, verifiedthe inter faces between the RCS and the other s ys tems . Engine valve wiring and re-sponse, clus ter heate r cur rent draw, and instrumentation wer e checked during the fina lphase of factory checkout.

LUNAR MO DULE RCS FLIGH T PERFORMANCEThe LM RCS performance on all Apollo flights was satisfactory. Sever al minorproblems occurred, but satisfactory solutions were found for all prob lems noted. The

L M - 1 and LM-3 flights indicated that the upper temperature limit of 190" F on the en -gine cl uster was exceeded on numerous occasions with no delet eriou s effects. As pre-viously discussed, additional vendor te st dat a demonstrated that the engine valves couldbe safely operated at a much higher tempe ratur e. Consequently, the limit was deletedon L M - 4 and subsequent vehicles.Pre ssu re switch "closed" failu res occurred during the LM-3 , LM-4 , and LM-5flights, but these failure s had no significant effect on the flights because the only con-sequence was the lo ss of capability t o detect an engine f'off" failu re. The single

16


21/23

pr es su re switch "open" failu re that occurred intermittently on LM-5 wa s considered t obe a se ri ou s problem because an errone ous TCA failu re indication resulted. Thisproblem was solved on subsequent missions by briefing the crews about the possibilityof an erroneous TCA flag. The switches were kept in the system p rimar ily to aid flightcontr oller s in the rea l-ti me analysis of co ntrol problems and RCS engine operation du r-ing the criti cal powered-descent mission phase. They wer e considered to be the bestswitches available f o r this function but were not as reliable as the engine. A brief dis-cussion of the LM RCS performance on the Apollo 5, 9, 10, and 11 missions is providedin the following sections.

Lunar Module 1 Apol lo 5 Mission)The unmanned Apollo 5 spacecraft, including the first flight-configuration LM, wastested successfully in e ar th orbit on January 22 and 23, 1968. The pri mar y objectivesof the Apollo 5 missi on were to flight verify the LM ascent and descent propulsion s ys-te ms and the abor t staging function fo r manned flight. These objectives were accom-plished. The LM-1 RCS configuration differed from subsequent LM RCS configurationsin several areas of feed syst em design and instrumentation design.During the mission, the RCS perfo rmance and operation wer e nominal until con-t ro l of the spacecr aft was switched to the guidance, navigation, and control systemafter abort staging (intentional separation of the ascent and descent stages of the LM).At that time, the vehicle m as s in the digital autopilot was configured fo r control of atwo-stage, fully loaded vehicle; consequently, the sys tem was commanded t o deliverpropellant at a rate approximately 10 000 t imes greater than actually required. Thisanomaly caused the RCS to operate in several off-limit conditions and resulted in f a i l -ures in the sys tem. Within 3.1 minutes, the system A propellant was depleted to27 percent, and that sys tem was isolated to conserve propellant. System B continued

at a rapid duty cycle until propellant depletion 5 minutes later, at which time heliumsta rte d leaking through the collapsed syst em B fuel bladde r. Satisfactory vehicle rateswere restored by the system B thrus t reduction (resulting from propellant depletion)and by the isolation of syst em A propellant tanks. While system B w a s operating withtwo-phase oxidizer and helium-ingested fuel, the quad 4 upfiring engine failed. Whensyst em A was r eactivated, the syst em A main shutoff valve on the oxidizer sid e inad-vertently closed. The ascent propellant interconnect valves wer e later opened, return-ing operation of the engines to nor mal until the interconnect valves were closed. Thedepletion of all propellant during the last minutes of the second ascent-engine fir ingallowed the spacec raf t to tumble. Each of these specific RCS anomal ies (i. e. , thebladde r, the engine, and the oxidizer main shutoff valve fai lures) was duplicated whena ground test sys tem was exposed to s imi lar duty cycle and environmental conditionsafter the flight.The tot al propellant consumption from t he RCS tanks w as approximately600 pounds. An additional 230 pounds of propellant w ere used f rom the APS tanks dur inginterconnect operations. It is estimated that the RCS engines accumulated 16 000 fir-ings during the miss ion. All inflight LM-1 RCS problems were considered to be a re-sult of the operational anomaly that caused the RCS to operate in se ve re off-limitconditions. Theref ore, no syst em design changes wer e made as a result of theseproblems.

17


22/23

Lunar Module 3 (Apol lo 9 Mission)The Apollo 9 mission wa s the second to include the LM and the first to include amanned LM. The successful earth -orb ital miss ion lasted approximately 241 hour sfr om launch on Marc h 3 to splashdown on March 13, 1969. The objectives of the mis -sion were as follows.1. Evaluate LM systems performance2. Evaluate LM functional capability3. Pe rf or m select ed CSM/LM opera tions (rendezvous and docking)

These objectives were accomplished.The LM RCS performed satisfacto rily throughout the mi ssion. The only problemnoted was a failed-closed thru st chamber p re ss ur e (TCP) switch, which was used tomonitor the quad 4 upfiring engine.The total propellant consumption fr om the RCS tanks w as approximately353 pounds as measured by the onboard propellant-quantity -measur ing devices. Anadditional 99 pounds were used from the A P S tanks during interconnect operations. It

is estimated that the RCS engines accumulated a tot al of 1250 sec onds "on" ti me and20 000 firings during the mission.A significant decr ease in the natu ral frequency of the LM RCS fuel and oxidize rmanifold-pr essure fluctuations was noted duri ng interconnect-feed operations a ssoci -ated with the APS burn to depletion. The decrease apparently was caused by eith erfree helium enter ing the RCS manifolds fr om the APS or a higher saturation level ofAPS propellant rel ative to RCS propellant. In any event, the condition was not detri-

mental to RCS operation.Lunar Module 4 (Apol lo 10 Mission)

The Apollo 10 mission w as the t hird to include the LM and the second to includea manned LM. The successful lunar-orbital mission lasted approximately 192 hoursfrom launch on May 18 to splashdown on May 26, 1969. The objectives of the missi onwere to demonstrate the satisfactory performance during a manned lunar mission ofthe crew, the space vehicle, and mission support facilities and to evaluate the L M per-formance in the cislunar and lunar envir onments. Thes e objectives were accomplished.The LM RCS performed sati sfac tori ly throughout the mission. The only problem notedconsisted of five failed-closed T C P switches.

The total propellant consumption fro m t he RCS tanks w as approximately557 pounds; however, only 276 pounds of the tot al wer e used du ring manned ope rations.An additional 42 pounds of propellant wer e used f ro m the APS tanks during interconnectoperations, It is esti mated that the RCS engines accumulated a tot al of 1640' seconds"on" ti me and 20 000 firings during the mission.

18


23/23

Lunar Module 5 Apollo 11Mission)The Apollo 11 mission was the fourth to include the LM and the third to include amanned LM. The successf ul lunar landing mission lasted approximately 195 hoursfr om launch on July 16 to splashdown on July 24, 1969. The pr im ar y purpose of themission was to perfo rm a manned lunar landing and to ret urn the cr ew safely to ea rt h.Th is objective was accomplished.Pe rf orman ce of the LM RCS was satisfactory throughout the mission. The onlyproblem noted involved two TCP switches. One of them was a failed-closed o r "on"condition similar to the exper ience on previous flights. The total propellant consump-tion f ro m the RCS tanks was approximately 319 pounds. An additional 69 pounds ofpropellant w er e used fr om the APS tanks during interconnect-feed operations associatedwith lunar lift-off. It is estimated that the RCS engines accumulated a total of 1060 sec -onds "on" time and 1 2 000 firin gs during the mission.During an 18-minute period j u s t before termin al phase initiation, the quad 2 aft-fir ing engine switch failed to respond to seven consecutive minimum-impulse com-

mands. Thi s situation resulted in a master alarm and a TCA warning flag, which werereset quickly by the crew. Engine opera tion was nominal, and the switch failu re hadno effect on the mission. Subsequent cr ew s were briefed that an erron eous TCA flagwas possible and that, theref ore, they should not abort unless the engine fai lur e wasverified by vehicle dynamics o r so me other means.

C O N C L U D I N G REMARKSSucces sful completion of the ground-tes t program, coupled with excellent flightperfo rmance of the reaction control syst em, proved the syst em to be highly reliable.To a large extent, th is high degree of reliabili ty can be attributed to the commonalityphilosophy applied to the command and service module and lunar module reaction con-tr ol sys tem components. Another significant factor was the sys tem and componentcleanliness levels that were maintained by flushing and by providing in-line filters up-st re am of cri tic al components.The luna r module reaction control sys tem did not experience significant designchanges during development, qualification, o r flight. This fact can be attri buted pr i-mar ily t o appli cation of the common-use and common-technology philosophy.The engine injector valves proved to be extremely reliable. In the total progra m

(both lunar module and service module), no engine injector valve leakage, as a resultof engine operation o r malfunction (explosion), was observed. The engine isolationvalves, which wer e incorporated into the sy ste m to deal with the leaking engine valves,thus b ecame of little value. It w a s concluded that the valves we re not e sse nti al to cre wsafety, and the decision was made to delete them fr om the syst em in later flights torealize a 25-pound weight saving.Manned Spacecraft CenterNational Aeronautics and Space AdministrationHouston, Texas, December 17, 1971914- 11-10-EP-72

Date post:	09-Apr-2018
Category:	Documents
Upload:	bob-andrepont
View:	228 times
Download:	0 times

Apollo Experience Report Lunar Module Reaction Control System

Documents