Sa 501 s Ivb Auxiliary

8/9/2019 Sa 501 s Ivb Auxiliary

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N A S A T E C H N I C A L N O T E NASA TND-5

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LOAN COPY: RETURNTO

KIRTLAND AFB, N MEXAF!'


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TECH LIBRARY KAFB, N M

Illllllllll1IIIIlllllllllI llllI110333772

INVESTIGATION O F SA-501 S-IVB AUXILLARY

P R O P U L S I O N S Y S T E M F L I G H T A N O MA L IE S

By K. C o a t e s a nd E . D o n a l d

G e o r g e C . M a r s h a l l Space F l i g h t C e n t e rM a r s h a l l , A l a .

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION~

For sa le by the Clear inghouse for Fede ra l Sc ient i f ic and Technic a l Informat ionSpr ingf ie ld , Virg in ia 22151 - CFSTI p r i c e $3.00


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TABLE OFCONTENTS

Page

SUMMARY . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . i

INTRODUCTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . IMISSION REQUIREMENTS. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2SYSTEM DESCRIPTION . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2PRE-FLIGHT TESTING . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5FLIGHTRESULTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8DISCUSSION O F ANOMALIES. . . . . . . . . . . . . . . . . . . . . . . . . . . . . I 1

MSFCTESTRESULTS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23

CONCLUSIONS . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40

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L I S TOF I L L U S T R AT I O N S

Figure Title

i . Saturn V/S-IVB APS Schematic . . . . . . . . . . . . . . . . . . . . . . .2 . I s o m e t r i c of Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

3 . Tapco Engine Cutaway . . . . . . . . . . . . . . . . . . . . . . . . . . . .4. Prope llan t Tempera tu r e . . . . . . . . . . . . . . . . . . . . . . . . . . .

5 . Ullage Engine Pc Trace . . . . . . . . . . . . . . . . . . . . . . . . . . .

6 .

7 .8.

Engine Chamber Pressur e /Tempe ra tu re Tr ace . . . . . . . . . . .

Comparison of Te st and Flight Engine Chamber P re ss ur e . . . .A r e a of APS Module Heated in Simulated "501" Te s t s . . . . . . .

9 . Sequence of Events Tim e Line for Simulated "501" Duty Cycle .10 . Engine I Pc. Injector Temperature. and Oxidizer Supply

Tempera tu re Versus Tes t Time . . . . . . . . . . . . . . . . . . . . .

l i Engine 2 Pc. Injector Temperature . and Oxidiz er SupplyTempera tu re Versus Tes t Time . . . . . . . . . . . . . . . . . . . . .

12 . Engine 3 Pc . Injector Temperature. and Oxidizer SupplyTempera tu re Versus Tes t Time . . . . . . . . . . . . . . . . . . . . .

13 . Facility Schem atic fo r TRW Engine "Sea Level" Te st s andInjector Flow Tes ts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

14 . Contaminated Engine Injector Oxidizer Flange Inlet . . . . . . . .15 . Cleaned Engine Injector Oxidizer Flange Inlet . . . . . . . . . . . .16 . Corros ive Attack on Injector Oxidizer P ass age Outlet . . . . . . .17. Normal Injector Oxidizer Pa ssa ge Outlet . . . . . . . . . . . . . . .

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L IS T OF TAB LES

Table Ti t le Page

I. Simulated ff501ff APS Engine Fir ing Prog ram . . . . . . . . . . . . . 26

II. TRW Single Engine Duty Cycle . . . . . . . . . . . . . . . . . . . . . . 31

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I N V E S T I G AT I O N O F S A -5 01 S - I V B A U X I L I A R YPROPULS ION SYSTEM FLIGHT ANOMALIES

S U M M A RY

Sever al anoma lies occ urre d on the S-IVB auxiliary propulsion systemduring the SA-50i flight. The anomalies of pri ma ry significance we re lowchamber p r ess ure fo r s eve ra l engines beg inning at the s ta r t of the S-IVBpowered f l ight and high propel lant tem per atur es experienced late in flight.An invest igat ion of possible c aus es by theoret ical analys is and hardwa re tes t ingwa s co nduc ted by MSFC.p e r a t u r e w a s cause d by a degradation of th e A u x i l i a r y Propulsion System (APS)module sur fac e paint ther ma l pr op ert ies , which in tur n was cau sed by S-I1stag e retr o-r ock et exhaus t gas plume impingement. The cause of the anomalouschamber p r ess ure s of se vera l eng ines at the beginning of S-IVB powered flightwas verified by testin g to be contamination of the engine injector because of"burp firing" at ambien t conditions at KSC. The investig ation res ulte d in thedeletion of the "burp firing" req uire men t pr io r to launch at KSC and in modifi-cat ion to the APS hard ware to preven t the tempe ratu re r is e of the propellant.

The an alyses indicated that the high propellant tem-

I NTRODUCTION

The SA-501 post flight evaluation rev eale d se ve ra l disc rep anc ies in the

opera tion of the S-IVB Stage APS. During the flight the sy ste m had appea redto be operat ing in a nominal mann er and contro l of the vehicle w a s maintainedwithin acceptable lim its. No effects of the anomalies wer e noted on the stageoperation. The pr im ar y mis sion of the flight was to ve rif y the adequacy of thelaunch vehicle de sign , Service Module design, and Command Module hea t shielddesign. In addi tion, sev er al maneuv ers we re perform ed to test the communica-tion s adequacy. The orientatio n experienced by the stag e during the extrane ousmaneuv ers was not considered to be a normal Saturn V mission requirement .

Since the amount and quality of flight data rec eive d was lim ited , thedegr ee of analysis that was performed w as al so limited and assump tions,

theoret ic al calculat ions , and ground tes t ing wer e relied upon fo r sev era l areasunder investigation.perfor med by the Propuls ion and Vehicle Engineer ing Laboratory and Tes tLaboratory of the Marsh al l Space Fl ight Center.

This document contains the re su lts of the investigation


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MIS S ION REQUIREMENTS

The Auxiliary Pr opulsion System is designed to provide atti tude controland propellant se ttl ing capability for th e S-IVB stage. The propellant settl ing

capability is unique to the Saturn V application and do es not exist fo r the SaturnI-B system. The APS provides rol l control during 5-2 engine powered phase sof flight, and pro vide s pitch, yaw , an d roll control during coas t per iods. Thesys tem is also requ i red to per fo rm seve ra l maneuvers fo r l andmark s igh tings,undocking, docking, etc. The ullage engine is r e q u i r e d t o f i re immediatelyfollowing injection into orbi t to keep propellan ts settled at 5-2 engine cutoffbefore opening the continuous vent valv e.during r ecir cula tion chilldown of th e 5- 2 engine pr io r to restart.

The engine is required to f i re again

The minimum impulse bit (MIB) re quir ed of the engines for atti tudecontrol is 7. 5 l b sec.

SYSTEM DESCRI PTION

The Saturn V/S-IVB APS co ns ist s pr im ar ily of a pressurizat ion sub-sys tem , propellant subsystem , engines , and instrumentation. A schemat ic ofthe system is shown in Figur e I . An is om et ri c view of the component andinstrume ntation locations within the module is shown in Figu re 2. The APS isbuilt in a mod ular concept in which al l subsys tems are enclosed inside thestructural fair ing. Two modules are located on the stage 18 0 d e g r e e s a p a r t on

the af t skir t .sta ge to control the firin g of the engines.

Electr ical s ignals ar e generated from the Instrument Unit and

The pressur izat ion system co nsis ts of a helium tank, a two-stageregulator , quad check valves , a low pr es su re helium module (which containsrelief and solenoid valves) , and helium f i l l check valves. The initial heliumbott le pressure is 3100 ps ig and is regulated to 196 psia . In the event of apri ma ry regulator fa i lure (wide open o r regulat ing to higher l imits is normalfai lure mode) , the secondary regu la to r reduces the p re ssu re to 200 psia.quad check valves reduce mig ration of prop ellant va por s upstrea m f rom the

Th e

propel lant tanks to prevent a mixing of ex plosive propellant vap ors .

monomethylhydrazine (MMH) and the othe r containing nitrogen te troxide ( N2 0 4 ) ,The propellant subsystem cons is ts of two propel lant tanks ( one containing

in e filter, f i l l and dra in valv es, and propellan t distribution lines. Each module

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HELIUM PRESSURE REGULATORA- CHECK VALVES

HELIUM,LOWPRESSUREMODU LE(FUEL) 7

I l l I I I

MODULE, PROPELLANT0x 1DIZER CONTROL

MODULE, PROPELFUEL CONTROL

FIGURE I. ATURN V/S-IVB AP S SCHEMATIC


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HIGH PRESSUREHELIUM SUPPLYTRANSDUCER

.REGULATOR OUTLET PRESSURE TRANSDUCERQUAD CHECK VALVES, FUEL lA67912T

\ QUAD CHECK VALVES, OXID IZER lA 67912HELIUM TANK TEMPERATUREr TRANSDUCER

OXIDIZER LOWPRESSURE HELIU M MODULE 1A49998

OXIDIZERULLAGE PRESSURE TRANSDUCER

OXIDIZER TEMPERATURE TRANSDUCEROXIDIZER PRESSURE TRANSDUCER

OXID4ZER PROPELL ANT CONTROL MODULE lA49422

FIGURE 2. ISOMETRIC O F MODULE


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contains 127 lbm of MMB and 188 lbm of N2 0, t liftoff. The propellan t tankscontain teflon bladders to provide positive expulsion of propellant in zero-genvironment. The f i l l and drain valves are contained in the propellant controlmodule, which also incorporates a filter with a nominal filtration rating of 10micr ons and an absolute ra ting of 25 micr ons.

Thr ee engines in ea ch module are used for attitude control and opera teat 147 lbf th ru st each.engine provides pitch control. Each engine is equipped with quad solenoidvalves , in each propellant inlet line , ar ranged in a series-parallel combination.A cutaway of the engine is shown in Figure 3. Th e inlet to the valve packagecontains a 100 mesh scr een ( 150 micr ons) to prote ct the engine from l ar geparticulate contamination.is equipped with single solenoid valv es in each propella nt inlet line.

Two engines provide ro ll and yaw control while one

The ullage engine is operated at 72 lbf thr ust , and

PRE-FLIGHT TEST1 NG

The mo dule s used f o r SA-501 and SA-502 undergo tests that a r e uniquein that no other Saturn V flight modules a r e hot fire d before installation on theflight stage at KSC.inert f luids at the Santa Monica facility.functional and elec tro -m echanical checks.

A l l APS modules a r e functionally acce ptance te sted withThese tests con sis ts of lea k and

Followi ng the Santa Monica tes tin g, the SA-501 m odul es we re shippedto the Sacramento Tes t Center (STC) , where a pre-firing checkout w a s con-ducted. Following the checkout, a confidence firing w a s performed, consis t ing

of a tota l of 217 pu ls es on the attit ude con trol en gines and 2 puls es on the ullageengine. During the fi ri ng of module I, it w a s noted that the performance ofengine 2 w a s marginally low.checkout fo r use in another test module, a fuel valve stuck open.w a s then shipped to Santa Monica for a failure analysis. During the post firingchec kout of mod ule 1 a faulty oxidizer propellant control module and leakingengine re circulation manifold was revealed. Both components were replaced.The pre -fi rin g checkout of module 2 revea led a sticking fuel valve in engine 2and the engine w a s replaced. The post firing checkout of module 2 rev eal edexcessive leakage through the pr imary refer ence pre ss ur e pa rt of the regulatorand exce ssi ve leak age through both check valv es of the fuel propellant controlmodule. Both compone nts w e r e replaced.

The engine was later replaced, and, duringThe engine

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7 F l ) r E L IN LE T P ORT


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Following the confidence firi ng , both SA-501 modules we re flushed withF reon M F to clean the module fo r handling purpos es. The modules w e r e thendisa sse mble d and components that had been exposed to propellants were cleanedby vacuum baking or purging. Defective and obsolete-configuration p ar ts w e r erep laced during the reassemb ly process . A final post-firing checkout on eac hAPS module w a s then performed.

Following the disa sse mbly and cleaning, it w a s discovered tha t se vera lfuel valves in the engines w e r e stuck. An analys is revea led that the modulefuel components w e r e not pro pe rly cleane d by the proc ed ure of flushing withF reon M F followed by vacuum baking. A ha rd r e s idue w a s found on the enginefuel val ves and in oth er components of the fuel syst ems . A l l of the engineswer e rep lace d and recl ean ed by the engine vendor. Five of th e six enginesuse d to repla ce the origin al 501 module engines had been previously used fo rthe SA-502 APS modu le confidence firi ngs . The SA-502 modu les had ex pe rie nc edthe sa me contamination as SA-501 modules. The engines valv es we re removedand replaced with new ones and the inject ors cleaned before reas semb ly into

the SA-501 modules. The fuel propellant con trol modul es w e r e replace d andall oth er fuel sys tem components recleaned.

A f t e r comple tion of the delayed rea ss em bl y, the module s we re shippedt o KSC, where a pre-flight checkout w a s per form ed. This consisted of leak,

KSC fo r approximately one year. A f t e r stor age the modules underwent pre -limi nary checkout and were ass emb led to the stage. The modules were thengiven a preloa ding checkout. During loading op er ati on s an oxidiz er tank blad derin module 2 was inadvertently overpressurized, and subsequently w a s replaced.

. functional, and electro- mechan ical checks. The modules were then stored at

A "burp" fi ri ng was conducted on the APS engin es on November 3, 1967,six days before launch.cle ari ng bu rs t, followed by two 65 milliseconds pulses on each attitude controlengine with each ullage engine fired for approximately 430 milliseconds. A l lexcept engine 3 of module 2 exhibited nominal performance .chamber p r e s su re trace indicated a slow and uncharacteristic buildup. A f t e rexamining the possible fai lure m odes, it was concluded that the anomaly w a scau sed by instrumen tation , and the system w a s judged to be adequate for flight.It w a s noted d uring the post flight evaluation that the th ru st decay as denoted byP c w a s slo wer on engine 4 module I than engine 4 module 2. A re-examinationof burp-firing dat a rev eale d that the sam e result oc cu rre d on the pad.

The bur p firi ng con sist ed of one 200 millis eco nds

The engine 3

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FLIGHTRESULTS

Flight data indicated that atti tude control f or the S-IVB was maintainedwithin pro gra mme d lim its throughout the flight in addition to perf orm ing al lrequire d maneuvers . During the f light the pre ssu riza t ion sys tem operated ina nominal man ner and the usage of propellant and pr es su ra nt w a s less than hadbeen predicted. If the mission had been flown without system evaluation instru-mentation, the flight would have been classified as a complete success.instrum entation perfor med the function fo r which it w a s intended and indicatedthe following anomalies:

The

H i g h P r o p e ll a n t Te m p e r a t u r e

Both modules exper ienced propella nt heating, which exceeded expecte d

amounts beginning at approximately I1 00 0 second s into the mission . Thetem pe ra tur e of the oxidizer in module I exceeded the limits of the tem pe rat ur et r a n s d u c e r ( 5 9 i " R ) . Tra ces of the t empera tu re are shown in Figu re 4.

Slow Chamb er P r e s su re Decayof Ullage En gine

Engine 4 of module 1 exhibited a slow er than nominal Pc tailoff f or eac hf i r ing (F ig . 5).f rom .the ful l value to IO ercent value within 35 d s e c fr om cutoff signal. Thisvalue is based upon the close-coupled tra ns du ce r used for engine manufacturetesting and with no s ur ge supp res sion circuitry.

The specif icat ion r eq uir es that the thr us t level shal l decay

Chamber P r e s su re Os c iI a t i ons

Engine I f module I exhibited oscillations in chamber pr es su re ofapproximately 34 p s i a peak-to;peak with a f requency of 400 hertz .c h am b e r p r e s s u r e w a s 80 psia.

The nominal

Chambe r P r e s su re Decay

Engines I and 3 of module 1 decreased in chamber p re ssu re f rom the i rinitial values to a s low as 55 psia during the miss ion period fro m I 1 000 secondsto 20 000 secofids of flight.

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I0x1 DI ZER TEMPE RATU RE

/ - P O D U L E

0

591.7 R

ULLAGE ENGINESI

5 15 20 25

6o01FUEL TEMPERATURE-I I

I IM ANE UVER TO L OC AL H O R l Z O N T A L l

ULLAGE ENGINESI

20

\-FIRST BURN OF ULLAGE ENGINES1 I

5 10 15

540rTIME FROM LI FT OF F (1000 SEC)

FIGURE 4. PROPELLANT TEMPERATURE

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150

Y 100W

zm50

E

n

0

150

Gin100

W9?

zm

50n

0

~ -APS ULLAGE ENGINE CHAMBER PRESSURE TRACE-MODULE2

f

NORMAL PRESSURE

_ _ - - - - - - - - - - - -

- - - - - - -

=q+DECR EAS E

FIGURE 5. ULLAGE ENGINE Pc TRACE

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Low C h a m b e r P r e s s u r e s

Engines I and 3 of module I ere approximately 15 psia low in chamberpressure f rom the f i r s t f i r ing s th roughout the miss ion .module 2 also exhibited lower Pc in flight than that shown during bu rp firing.Engine I exhibited values f ro m 5 to 10 perc ent low during the mission . Engine2 Pc w a s consistently 8 to 10 pe rce nt low throughout the flight. Engine 3 Pcw a s also lower than expected ( < 5 percent) throughout the flight, but was withinsystem tolerances.and did not caus e mis si on degradat ion during the flight.

Engines I an d 2 of

The anomalies l i s ted w ere noted in the post f l ight evaluat ion

D I S C U S S I O NOF A N O M A L I E S

H i g h P r o p e l l a n t Te m p e r a t u r e s

The des ign of th e APS is intended to maintain the propel lant tem per atur eswithin acceptable lim its by means of passiv e control. Th is is done by m ea ns ofthermal insulato rs placed in conduction paths to cr i t ical components , by insula-tion placed aroun d pro pel lant tank s, by polishing of propella nt lin es , and bypainting the s ur fa ce of th e module with a ther mal ref lect ive paint which keeps

the absorbtivity to emiss iv i ty ra t io - o I. 0. Accord ing to McDonnell Douglas

Corporation (MDC) personnel, the effectivity of t his coating appeared to b edegraded (f ro m the ant icipated values) during the e art h orbi tal operat ion.During development testing , the APS was subjected to extensive the rma l environ-mental t est ing in th e MDC high vacuum fac ility at Huntington Beach. The modul ewas subjected to both hot and cold temp era tur e cycling with a normal sur facecoating and with a black surface. The tempera -tu res in fl ight could have r esul ted from one o r a combin ation of the following:

aE

N o problem s were experienced.

I. ~~~~~Degraded Surface Paint at Liftoff. The APS modules a r e normallySince the paint is a si l icone

The modules wer e handled extensively during theThe paint was degraded

painted at the Santa Monica manufacturing facility.type, it is easily rubbed off.confidence firing s and subsequent checkout operati ons.to such aneextent that it w a s deemed necessary to repa in t the modules at KSC.The effect of the repain ting operation is unknown.

2. Degraded Surface Paint Because of Boost Heating. The effe ct onthe sur face proper t i es of the paint because of the tempera tur es exper iencedduring boost is not normally considered to cau se any degradat ion. The SaturnI-B APS modul es are t rea ted in the sa me manner and undergo a s imi l a r b o os t


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situation, and no proble ms have been noted to date. The repaint ing proces smu st again be noted as rep res entin g an unknown effect.

3. Degraded Surface Pai nt Due t o Retro-Rocket_ _ Exhaust PlumeL m p i n g e m e n t m m a l tes ting at Huntington Beach included the black paintcoating in an effort to produce the wo rst effect caused by plume impingement.

It has been noted, however, that the worst case is rea l ly some surface p r o p e r t i e sbetween those rep res ente d by the s i lvery color and black co lor where the -ra t io is grea te r than I. 0. Thus the effect caused by retro-r ocket plume im-pingement re pr es en ts another unknown.

_ - -

a!

c

4. Severe Miss ion Heating Profii. A s sta ted previously in the intro-

The orientation of the stag e to the sun w a s changedduct ion, sev er al maneuv ers were performe d on experimen ts that are not normalto the Saturn V application.at M 10 800 second s which produced a higher hea t flux on the oxidizer side ofmodule I and fuel sid e of module 2. The tempera tu re rise began to occur at

thi s point.

A f t e r examining each of the possible ca us es , the mo st logical explanationfo r the anomaly is a combination of ( 3 ) and ( 4 ) . H e a t transfer calculations byMDC and an exam inat ion of the flight da ta do indicate a degra dation of sur fa cep r o p e r t i e s in ea rt h orbit . Since the Saturn I-B modules have been repainte d inthe pa st before flight and the s ur fa ce coating did not deterio rate , cause ( 1) m u s tbe eliminated. Cau se ( 2 ) can also be eliminated becau se of similar i ty toSatu rn I-B, and the fact that the mos t severe degra dation of the paint, if causedby boost heating, would have oc cu rr ed on the no se of the module with ve'ry littlehea ting on the af t end.

module s towards the su n indicates just the opposite effect.tion of ( 3 ) an d ( 4) re pr es en ts the mos t probable ca use of the anomaly.

The heating c au se d by orie nta tion of the aft ends of

Thus the combina-

Slow Chamber P re s su re Decayof Ullage Eng ine

A s shown in Figu re 5, the p res su re decay fo r eng ine 4 f module I w a sslower than specif ication l imits . The data received rep res ent s the first flightenvironment chambe r pr es su re data ev er received fo r this engine. The decayc h a r a c t e r i s t i c s are a function of the distance of the pr es su re detection devicefro m the so urc e of the pre ss ur e and the size and volume of line leading to the

pr es su re trans duc er. The amount of time fo r decay of the engine from cutoffto 10 p e r c e n t Pc w a s less than I second.engine w a s repeatable fo r both starts in spac e, and app ears to be s imi lar in

The cutoff characteristic f o r the

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t im e to the tailoff in the se a level burp firings. The data wer e di scussed withRocketdyne ( engine vendor) and their opinion w a s that the engine performedsatisfactorily and such traces had been noted in som e of the ir sea level testing.However, the flight trace was different from that nominally expected. Pos sibl ecauses f o r the slow tailoff would be:

I. Slow Engine ~ Valve Shutdown. An exa min atio n of th e trace indicatedthat the Pc decayed in a norma lly expected mann er down to 25 perc ent offull value, and then the rate w a s decreased . Normal operat ion for a stickyvalve would be initial slow movement ra th er than slow movement at the endof the str ok e. The slow mov eme nt of a valve poppet would be caused by bindingof the poppet within the bore be cause of t he rm al expansion, tole ran ce mism atch ,o r contamination. Since the val ue did move rapidly to nea r full close condition,th is mode of ailu re would ap pe ar to be unlikely.

2. Contamination on Valve Seat. Since the valve initiall y did moverapidly to cutoff propellant flow , but ap peared to not seal immediately, apossible explanation could be a particle imbedded in the poppet seal material .Since the sam e c har act er i s t i cs a r e noted for both starts and the propellant isflowed through the valve in a steady state operat ion, any contam ination on thes e a t holding the valve poppet off the seat would have to be imbedded to keepfr om being washed away. Sincethe burp f i r ing a lso indicates a long er than usual tailoff, the possibility ofche mic al contam ination buildup will be eliminated. To have cau sed slow tail-off in th re e instanc es, the parti cle would have had to be in the seat before sealevel firing.

Possib ly this could cause such a Pc tailoff.

3. Chamber Pressure Instrumentat ion Cha ract er is t ics . The length ofl ine connect ing the pre ss ur e t ran sdu cer to the engine chamber is longer thanis norma lly design ed f o r th is type of engine. However, si nc e the engine is nota pulsing engine and operates only in a steady state firing mode, the steadystat e value indicated by the pre ss ur e t ransd uce r is independent of the length ofl ine. This is not tr ue fo r buildup and shutdown ch ar ac te ri st ic s, however. Theindication of Pc tailoff would be lengthened by the amount of tim e re qu ire d fo rthe vacuum environment to empty the line of any gase s. It mus t be noted, how-e v e r , that engine 4 of module 2 is instrumented in the sam e manner and the Pctailoff was significantly faster.blocking the Pc l ine , this type of tailoff trace would be expected.

F u r t h e r , if any kind of particle w a s partially

Several o ther poss ib le cause s , such as valve seat swelling, were ex-amined and eliminated.contam ination (which' was imbedded in the teflon) on the valve seat. Theuniqueness and mino r effect of the anomaly does not indicate that any kind ofse r i ous prob lem ex is t s , o r tha t it should wa rra nt any fur the r s tudy.

The mo st probable caus e w a s determi ned to be a slight

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Chamber Pressur e Oscillations

Durin g the pe riod beginning at 16 000 seconds and extending pa st 20 000seconds the chamb er pre ss ur e oscillations of engine I e r e greater thannominally experien ced by the attitude control engines. Oscillations of x 35 ps iapeak-to-peak we re noted at a frequency of 400 hertz. By closely examining allof the attitude con trol engine Pc traces, it is noted that al l engines exhibit somedegree of rough combustion. This is a characteristic of the particular enginedesign. Oscillations at nearly t he s am e frequency have been noted during se alev el testin g and in prev ious flights. The magnitude of the previ ously notedoscillations have been in the 5 to 10 per cen t ps ia peak-to-peak range. Thi stype oscillation ha s not been the cause of any problem in the pas t even thoughit is undesirable. Possible cau ses for the osci l la t ions a r e as follows:

I. nstrumentat ion. The pr es su re t ransd ucer normal fai lur e mode isa shift in calibration o r los s in reading. Pr es su re t ransducer fai lure indicatingoscillations of the type indicate d in flight have not been observ ed in any previou stesting.puts f rom the rate gyros verif ied this , indicates that the t ransduc er was per-formi ng normally.perform until the end of flight.

The fac t that the engine chambe r pr es su re was low initially, and out-

The data indicates that the t ransducer w a s continuing to

2. High Prope llant Temp eratu re. The oscillation occ urr ed during theperio ds of highest propellant temperat ures.tempe ratur e we re investigated in testing by th e engine vendor and in syste mtest ing at MSFC.in previ ous ground testing. Because of propellan t manifold fluid dynamics , theoxidizer can theoretically vaporize in the lines when the tem per atu res a re highand propellant inlet pr es su res drop during engine start operations. I t is prob-able that these conditions existed during the ch ambe r pre ss ur e oscillations.The high propellan t tempe rat ure i n conjunction with low propellant in let pre ss ur eha s been s imulated in MSFC test ing, and chamber pr es su re osci l lat ions ver ysi mi la r to the flight oscillations were obtained.

The effect s of high prop ellan t

No oscil latio ns of the type ex perie nced in flight w e r e noted

3. Reduced _ _ _ _ropellant FW. The engine that ex perienced oscillationsw a s al so running with reduced cham ber pr es su re . A possible cause of lowcham ber pr es su re could be reduced propellant flow caused by a res t r i c t ion inthe propella nt flow path. Th is would indicate a lower than nominal propellant

inlet pr es su re ups tream of the engine injector o r increased injector resis tance.The propel lant pre ss ur e drop necessary to cause a 15 percen t drop in enginePc issimulated in tests at MSFC.section.

55 psia. The combination of ( 2 ) and reduced pr opellant flow wasThe result of the tests will be discu ssed in a later

14


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C h a m b e r P r e s s u r e D ec ay

The decay in chamber p res sur e of engines I nd 3 of module I ccur redduring the period of high propellant tem per atu res ( Fig. 6 ) .t empera ture was the obvious suspect after reviewing the data and noting thecorrelat ion between the chamber pr es su re decay and oxidizer temp eratu rerise. The res ul t of high engine injector and oxidizer tem pera ture and norm almanifold pr es su re osci l la t ions would be a reductio n of flow ca used by vapo ri-zation of th e propel lant in the feed system.caus e of the anomaly, and the test sect ion w i l l cover the results of the MSFCtests to ver ify the cau se of the anomaly.

The high propell ant

This was determined to be the

Low C h a m b e r P r e s s u r e s

Degraded perf orm ance w as noted in engines I nd 3 of module I an dengines I nd 2 of module 2.

low in Pc ne ar the beginning of AP S operation and continued at those levels o rlower throughout the miss ion.were 5 percent low in Pc would not normally be considered an anomaly sinceinstrumentation, engine tolerance, etc. , would allow fo r this. However, forthis mission, the low Pc values may indicate a t rend. A comparison of testand flight engine chamb er pr es su re is shown in Figure 7.

Four engines were approximately 10 to 15 percen t

The remaining two attitude control engines that

Burp f i r ings of the engines s ix days before f l ight ver if ied that eachengine w a s performin g in a nominal manner .verificatio n of nom inal integrity .remained in the module fo r flight.

p ressur ized to nominal operat ing pr es su re in a routine manner.indications at liftoff were that the SA-501 APS w a s read y f or flight.

This f i r ing represen ts the lastThe sam e propel lants used in the burp f i r ing

Immediately before liftoff the A P S w as

Thus al l

Considering the low Pc at beginning of flight the following poss iblecauses w e r e examined:

Ins trumenta t ion e r r o r

Change in thro at area

Engine valve malfunctions

Particulate contamination

Chemical contamination

These possible caus es are disc usse d below:

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TIME FROM LI F TO FF (1000 SEC)

MODULE1 OXIDIZERtiiMODULE1 FUEL560

55 0

5400 2 4 6 8 1 0 1 2 1 4 16 18

TIME FROM LIF TO FF (1000 SEC)

5

26

FIGURE 6. ENGINE CHAMBER PRESSURE/TEMPERATURE TRA CE


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h

4-1oc

Y

WK3v)v)

WK

KW

n

m3E 9c

8(1

202 50i

501

I I I I160 180 200 220 240

MANIFOLDPRESSURE (PSlAd

FIGURE 7. COMPARISON OF T ES T AND FLIGHT ENGINECHAMBER PRESSURE


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Instrumentat ion Er ro r. The first logical candidate was instrumentation.Since thi s ishe magnitude of the decreased chamber p ressu re was M 15 psia.

essent ial ly sea l eve l ambient p ress ure , a sh if t in re fe rence pres sur e of thetra nsd uce rs would be suspected. The last rec ord ed bench calibration of thetra nsd uce rs recor ded i n the module log book was about the middle of 1966,approximately one and a half ye ar s before flight. However, even though thein te rnal re fe rence pre ssu re is sea level ambient , the loss of the re fe re nc epre ssu re would increase ra ther than decrea se the chamber p res sur e read ing.A s tudy of rate gyro and position indicator data by MDC al so indicated that thesuspe ct engines were operat ing in a degraded manner. Thus, it is coneludedthat instrum entation was not the cause of the anomaly.

Change in Thro at Area. An enlargement in throat area caused by throate ros ion or loss of thr oat i ns er t would re su lt in a reduction in chamber pressure.However, thi s type of engine failur e has not occ urr ed sinc e earl y in the enginedevelopment progr am. Fur the rmo re, the re was no fir ing o r condition that wouldcause th is to occur d uring the period between burp fi rin g and the first f i r ings inflight.under the flight conditions is so remo te that this mode of fai lure is not worthy offurth er discussion.

The possibility that a s e r i e s of throat failures would occur in one flight

Engine Valve Malfunction.- The attitude control engines u s e quad re-These va lves a reundant val ves in each propellant line to co ntrol the flow.

designed to elim inate a single point failure in e ith er a fail-to-open o r fail-to-close mode. Valve failu re mode testin g w a s perfo rmed on these engines by thevendor during the development and qualification progr am.simu latin g both failure-to-open and failure-to-close to det erm ine the effectupon the engine performa nce.ser ies-paral le l valve arrangement caused a Pc de creas e of 5 psia.mode would not provide the characteristics noted in flight.exhibited nearl y the s am e degradation ear ly in flight, the engine valv es wouldappea r to be an unlikely source.

Tests were run

The c ase of a fail-closed valve in one le g of the

Since four engi nesThis fai lure

The probability of valve failures of the sam e natur e occu rring at nearthe s ame t ime to produce the same resul t in four engines is highly unlikely.Thi s reas oning would place any type of engine prope llant valve malfunction inthe same category as an improbable fail ure candidate.

Parti culat e Contamination. Precaut ions a r e taken to ensu re that noparticulate contamination i s allowed to ent er the sy ste m both dur ing buildup andtes tin g and by m ean s of fluid s loaded into the module.cleanl iness procedures imposed for al l operations, and use of filters both within

These precaut ions are


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the modules and in the GSE. In spi te of the pre cau tio ns , however , particulatecontamination has been found in the sys tem . Metal par t ic les have in a numberof instances e i ther caused leakage pa st a valve seat o r cau sed binding of avalv e poppet in the bore. Thes e problems occu rred pri mar i ly during develop-me nt testi ng. In none of the cases was a degradat ion in engine performa nce ofthe magnitude experienced i n SA-501 flight noted.

Examining the syste m (F ig. 1) it is noted that only propellant feed

The l ine s ize f roml ine s , a filter, propellant valves, and engine injector constitute the flow pathfr om the tank outlet to the engine combustion chambe r.the tank outlet to the fi l ter ent rance is 3/4 inches in diamete r.f e eds al l engines, n6 flow res tr ic t io n occu rred in this sect ion.leve ls fo r the f i l ters are 10 mic ron s nominal and 25 microns absolute.ret ical ly, the maximum siz e part icl e that could pa ss through the filter wouldbe 9. 84 xThe clogging of the filter by particu late contamination mu st als o be ruled outsin ce the f low fo r all four engines in each module pas se s through the f i l ter , and

eff ect s of clogging would be s ee n on all engines alike. The size of the prima rypropel lant feed l ine from the filter to the engines is 3/8 inches in diameter.The a r r angemen t is such that the prop ellant flows through the filter into acavity at th e filter base in the propellant control module where it is fed fromtwo manifolds to the engines. Thus, i f flow w a s r e s t r i c t e d in the prope llantcontrol module cavity, it would be se en by all engines but i f it w a s r e s t r i c t e din the manifold s feeding engines 1 an d 3 , then engines 2 and 4 would be unaffectedo r vice ve rs a. Stated s imply, the problem occ urr ed downstream of the pro-pel lant control module, either in the manifolds feeding the eng ines, o r in theengines themselv es .

Since this lineThe f i l t rat ion

Theo-

inches in diameter. The length of s uc h a par t ic le is indeterminate.

Again reviewing the sys tem , it is noted that even i f the main propel lantfeed l ines to the engines are blocked upstream of the engine flange the enginescan still recei ve propel lants through the s ma lle r ( 1/4 nch ID) ecirculat ionl ines. Test ing at MSFC indicates that even with the primary propellant inletlin es completely blocked, the engines still opera te with a Pc of approximately85 psia. If the manifold to two engines w a s res tr ic ted , the r es ul t should be ' thesa m e on both engines.propel lant flow res tr ic t io n to be in the engines themsel ves unless the rec i rcu la -tion line is a lso res t r ic ted .

This was not the case. This fact would indicate any

A cutaway view of the eng ine is shown in Figure 3. The engine flow pathcontains a 10 0 mesh ( f i l t ra t ion leve l of 150 mic rons ) s c r e e n filter at the valveen t r anc e ( r ec i r cu l a t i on l ine is downstream of fi l ter) , the solenoid o perat ed quadva lve s , a flow control orifice at the valves out let , 12 capillary feed tubes on the

19


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oxidizer s id e, one feed tube fo r fuel , and the engine injector.valves open to a distance of 0. 040 f 0. 003 inch es off their seats. If contamina-tion blocked the opening and reduced the s tr ok e of the poppet, leakage wouldoccur. However , i f the contamination had built up behind the poppet , thepropellan t flow area would be reduce d in the valve open position, and no leakage

would oc cur in th e closed position. The reduc tion in movem ent of the poppetwould r educ e the Pc buildup rates due to restr ic t ion of flow. In or de r to causePc reduction s of the or de r experienced in flight, at least one valve in eachpara l le l leg of one propellant would have t o be affec ted in ea ch engine. Partic-ulate contamination would not ap pear t o be a logical condidate to c a u s e at leasteight valves ( two in each engine) to be re str ic ted. It is noted that the back sid eof th e valve poppet is essent ial ly a ffdead-endedff olume. Part icul ate contami-nation would be mo re likely to oc cur som ewh ere along the flow path, and buildup at a f low res t r i c t ion ( f i l t e r, o r i f i ce , etc. ) ra th er than in the poppet bore.

The propel lant

The flow co ntrol or ifice s have a minimum diameter of 0. 060 inches.

The s ize of part icle req uir ed to clog the orifice would have to be of such a s i z ethat it would not be able to pas s through the f i l ter s and valves unless i t wasfluid in natu re ( i . e. a chemica l ge l ) .of the flow restriction.

Th is would not be the mo st logical so urc e

The oxidizer is fed into the engine injector by 12 capi l lary feed tubesthat a r e nominal ly 0. 0355 inches in diamete r and 3. 18 3 inches i n length.fuel feed tube is nominally 0. 1190 inches in di ame ter and approximately 2. 5inches in length.inches in di ame ter while the twelve fuel injec tor holes are 0. 0283 inches indiameter. Again a single part icle of parti cula te contamination requir ed to clogany of th e openings would not logically be a ble t o pa ss through the filters up-s t ream. Unless a part ic le is generated within the valve package i tself , apart ic le l arge enough to cause the flow re str ic t ion nece ssar y for the P c decayshould not be the sou rce of the problem. The probability that an accumulationof smal ler par t ic les is deposited in a flow path in four en gines is unlikelyunless all cleanlines s precautio ns we re completely ignored. The burp firin gsshowed this not to be the case before lift off.

The

The twelve injector oxi dizer holes are nominally 0. 0354

Because of filters, sizes of lines and openings in the flow path s, ando t h e r f a c t o r s , it is concluded that particu late contamination (ex ter nal o rinternally induced solid particle s) is probably not the ca use of the performan cedegradation.

Chemical Contamination. _ _ Chemical contamination for this discussionis defined as contamination that is generated by re actio ns within the syste m

20


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between propel lants , syste m hardw are, and the ambient atmosphere. Thefollowing types of c hem ica l contamination w e r e examined:

N2O4 flow decay

Propellant /moisture/hardware incompatibility.

N204 Flow Decay. The N2O4 flow decay phenomenon is one in whiche i t he r a gel- like m a s s of propel lant re st r i ct s the opening to a flow passa ge, o ra ma ss of mater ia l is deposited on the w a l l of f low restr ict ions (orif ices,valves, etc. ) .f e r rou s n i t r a te [ ( No) Fe NO,)^ ] that is forme d when the propel lant has beens to r ed o r has been in contact with ferr ou s materials . It mu st be noted thatnot a single case of flow decay o ccu rrin g in a propulsion sys tem during asimulated mission propulsion test ha s been documented.have always been in te st facilities where an at tempt was being made to induce

the phenomena o r a long dura tio n steady flow of pr opell ant w a s being passedthrough a faci l i ty to test som e sp ecific component.mo st knowledgeable on the s ubj ect , flow decay occ urs when a dec rea se i ntempe ra ture occurs in the prope llan t at a place in the s ystem where a p r e s s u r edrop occurs .re qu ire d to initiate flow decay ha s not been thoroughly defined. Te sts indicatetha t sm al l changes in tempe ra ture ( < 5' F) can caus e the decay to begin occur-r i ng at a flow restriction.decay to disappear.

The mate ria l deposi ted has been determined to be crystal l izable

The instances observed

According to person nel

The magnitude of the temp eratur e and pr es su re dro ps that is

Increa sing propel lant tem per atu re caus es the flow

The conditions experienced by the APS on the launch pad before flight

and during the boost pha se of flight a re not considered to be conducive to flowdecay sin ce the re w as no propel lant f low during this period.of the contaminant out of the propel lant while in the qu iescent s t ate on thelaunch pad would tend to se ttl e into the bottom of th e prope lla nt tank, anddurin g flow would be expected to collect on the prope llant control module filter.Any re st ric ti on of flow through the fi lt er would be ref lected on all engines alike.Th is did not occu r. The amount of contaminant that could for m in the prop ellantfeed sy stem while on the pad would be s o smal l that it should not clog any com-ponent in the sys tem. Based on the conditions that e xiste d in the APS, it wasdetermined that N2O4 flow decay wa s not the cau se of t he anoma ly.

Any precipitation

Propel lant /Moisture/Hardware Incompatibi l i ty. React ions can occu rbetween the oxidizer ( N 2 0 4 ) and w a t e r to produce ni t r ic ac id, which at tacks 'many meta l s to som e degree. Since no previous problem s had been noted inground test in g o r f l ights , it was a ssume d that no problem existed in the basic

21


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29/49

Another input which indicated that this phenomenon w a s the problemw as the r e su l t s of a se r ies of tests conducted by MDC at the i r Sacramentotest facility on a Saturn V APS module. The objective of the program w a s todetermine the adequacy of the syste m to withstand extended hold pe rio ds,s imulat ing launch aborts . Burp f i r in gs were conducted on the engines a t thestart of the 9O-day program and at i5-day intervals thereafter.

Two engines began to exhibi t low chamb er pr es su re after the first 15-day peri od and continued to d ec r e a se at succeeding f i r ings. A meeting betweenpersonnel from MDC and MSFC r esu lte d in the conclusion that the problemcould be a reduction of th e o xidiz er flow caused by a buildup of th e nit ri c acid/hardware cor ros ion products in the engine injecto r. It was decided that MSFCshould conduct a series of tests on a single engine to simulate the launch padconditions to ver ify the validity of the as su me d probable cause. Resu lts of thetests (which a re pres ented in the next sect ion of t his document) indicated thatthe assumpt ion was cor r ec t .

MSFC TESTRESULTS

A spec ia l test p ro g ra m w a s conducted at the MSFC T es t LaboratoryStorable Propel lant F aci l i ty to provide an experimental b ase f or defining theAPS Pc anomalies o bserved during the SA-50i f l ight .dir ect ed toward repr oduc ing the SA-501 flight effect and identifying the con-tr ibut ing cau se( s) .five specific phas es involving single engine as well as complete module test in gat both sea level and simulated altitude conditions.effort a re disc ussed below.

This test effor t was

To accomplish this , the test program w a s divided into

The resu l ts of this test

A P S Module Testing

An initial test ser ies w a s conducted on the Saturn V Pha se IV T es t APS

In these tests the APS engine f i r ing densi tyModule in an at te mpt to rep roduce the Pc decay phenomenon obse rved i n thelatter port ion of the AS-501 flight.was approximated and signif icant module temp era tur e conditions experiencedduring the SA-501 flight w e r e simulated. Te sts wer e conducted in the MSFCTe s t Labora tory i2 - foo t d iam eter a l ti tude cell at a s imula ted pr es su re a l t i tudeof 120 00 0 feet.

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The simulated "501" APS duty cycl e, as shown in Table I, ons isted of

Each attitude control529 puls es on engine I, 1088 puls es on engine 2 , 529 puls es on engine 3 , an dtwo steady state b u r n s totaling 420 seconds on engine 4.engine pulse wa s 65 mill isec onds in duration. Before beginning the pre-pro-gramme d f i r ing cycle , 3-second f i r ings wer e conducted on each engine toestabl ish the s teady state Pc operat ing level. Similar f i r ings were conductedfollowing the duty cycle. To provide the the rm al simulatio n, the oxidizercontrol module , consis t ing of the f i l l valve , recirculat io n valve , filter , an dsupply manifold , wa s heated beginning at 12 000-seconds into the duty cycleusing a 2 . 5 ampere in f ra red lamp. A photograph showing the heating appara tusand the a r e a h e at e d is shown in Figure 8. The significant events for the 19 800-second duration duty cycle are depicted in a t ime line shown in Figure 9.

In a ful l s imulat ion test, the APS module was loaded 3 t o 5 days beforea test and w a s maintained under a 50 psig gaseous hel ium blanket pr essu re.Nominal day-night t em pe ra tu re excursi ons of 30" F were experienced duringthi s period. The purp ose of the hold w a s to s imula te KSC prelaunc h groundholdconditions. A certain de gree of G H e saturat ion in the prop ella nts would beexpected duri ng the hold. In one test the module was loaded and the duty cycleconducted the day of the test to minimize the GHe s aturat ion to determ ine theeffect, if any, of satu rate d propellan ts on th e Pc decay phenomenon.

The test data revealed a significant decay in engine Pc with inc rease inpropellant and engine temperature. However, for the initial 12 00 0 seconds ofthe tes ts , there w as l i t t le var iat ion in the nominal Pc operat ing levels .exceptions to th is were the times that 3- to P m i n u t e h ol d p e r io d s w e r e i n c u r r e din or der t o change the F-M recording tape.this hold exhibited reduced Pc values. However, the nominal Pc values wereres to red a f te r two or thre e pulses. It was noted that the engine temper ature s"soaked outf' at higher values during thes e hold periods. The dense port ionof the fi rin g cycle was begun at 10 900 seconds, followed by the beginning ofheating of the ox idiz er cont rol module at 12 00 0 seconds. P c decay began atthis point ( 12 000 seco nds) and continued throughout the re ma ind er of th e dutycycle.contain curves of engine Pc , engine injector wal l tempe ratur e , oxidizer inlettemp erat ure, and oxidizer control module ( PCM) in let t empera ture versus testt i m e f o r a typical test. The maximum experienced Pc decay in this test pro-gram was 32 psia . Note that the Pc values retur ned to nominal during the

post-test steady state f i r ings.

Th e

The in itial engine pul ses following

The Pc decay is bes t i l lus t ra ted in F igures 10, 11, and 12, which

2 4


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FIGURE 8. AREA OF A P S MODULE HEATED IN f3IMULATED '*50$" TESTS .


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NOTE: SEE TABLE I FOR TH E ENGINE FIRING RATE S FOR SEQUENCES 1,2, AND3.

START PULSE SEQUENCE3

START HEATING OXIDIZER PCM

ULLAGE ENGINE FIRE FOR 330 SEC

PULSE SEQUENCE1 START PULSE SEQUENCE2

ULLAGE ENGINE FIRE FOR90 SEC

0 660 SEC 10 90 0 12 000START 11 160 12660

FIGURE 9. SEQUENCE OF EVENTS TIME LINE FOR SIMULATED"501" DUTY CYCLE

TABLE I. SIMULATED 17501" APS ENGINE FIRING PROGRAM

Time~~

0-10 900 sec*

660-750 see

1 0 900-12 660 sec"

I 1 160-11 490 sec

12 660-19 80 0 sec"

Engine I

2 pulses/100 sec

- -

14 pulses/100 sec

I ulse/I 1 0 sec

._

Engine 2

8 pulses /100 sec

~. .

8 pulses/100 sec

1 pulse/95 sec

~~ .

~ -

Engine 3

I ulse/250 sec

25 pulses/10 0 sec

I pulse/66 sec

I19 800

CUTOFF

~

Engine 4

'$ Pulse Width = 65 mil l isec onds

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g 100v)v)Wpcn

h

LL

0

wpc3c4n!Wn

v

3c

160

TEMPERATURE20

80

40

ELAPSED TEST TIME (SEC)

FIGURE I O . ENGINE 1 Pc, INJECTOR TEMPERATURE, AND OXIDIZER SUPP LY

TEMPERATURE VERSUS TEST TIME

1


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120h

iv)

nY

W

5 00v)v)

WeLn

80 J'

ENGINE2 P c

I

PNGINE 2 OXIDIZER INLET TEMPERATURE

. _ - - - I

ELAPSED TST TIME(SEC)

FIGURE 11. ENGINE 2 Pc , INJECTOR TEMPERATURE, AND OXIDIZER SUPPLYTEMPERATURE VERSUS TEST TIME


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120A

1E

Y

w

i ooY,wG!n

80

h3W

240

- 200

160

A

LL

120,wPL

2aWn5 80

TEST C-008.534; JUNE 11, 1968 I IPHASE IV SATURN V APS MODULE , IMODULE LOADED5 DAYS PRIOR TO ENGlNk 3 INJECTOR WALL TEMPERATURE

/

0

OXIDIZER INLE TTEMPERATURE I

TEMPERATUREI

I

f t- START SEQUENCE3BEGIN HEATING I

f- NkNE 4 FIRED FOR 330SEC1

START SEQUENCE2 I CUTOFFL- 4 MINUTE HOLDTO CHANGE F- TAPE

h I 12000 4000 8000 10 000 12 000 14 00 0 16 000 18000 2 0 (

00

ELAPSED TEST TIME (SEC)

FIGURE 12. ENGINE 3 Pc, INJECTOR TEMPERATURE, AND OXIDIZER SUPPLYTEMPERATURE VERSUS TEST TIME

DO


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Th e results f rom the M S F C tests substantiated the theory that the causeof the decay in engine Pc was vaporizatio n of the oxid izer in the injecto r andinjector feed tubes due to elevated propellant and/or engine tempe ratur e.hydraulic resi sta nce c reat ed by the pres enc e of ga s in thi s area r e s t r i c t e doxidizer flow, ther efor e resulting in the Pc drop. The cau se of the N2O4vaporization, as state d above, w a s a combination of hot har dwa re and hot pro-pellant. With the hot injector and inje ctor tub es, a condition aggravated by highpropel lant temp erat ure, the temp eratu re of the N2O4 was elevated to the pointthat it gasified before reaching the injection orific es.Pc values were r est ore d during the post-test s teady state runs as the propellantinlet tempera ture dropped and , subsequently engine hard ware temp eratu re w a slowered beca use of the continuous flow of c ool er propellant through the inje ctorand injector feed tubes , fur the r substantiated oxidizer gasification as the causeof Pc decay.

The

Th e fact that nominal

Gaseous helium saturation in the propellant showed no effect on the Pc

decay phenomenon.engine pe rfo rm anc e in the form of low frequen cy comb ustion instability of*8 to 1 0 psia ampli tude, 30 0 he rt z frequency. Additional disc ussi on on theeffect of satu rat io n is contained in the following sub-section.

The primary influence of GHe saturat ion w a s in oscillatory

Single Engine Altitude TestingTo mo re specifically investigate propellant te mp era tur e effect on engine

Pc, s ix tests w e r e conducted on one TRW attitude control engine in the MSFC3-fOOt di am et er alt itud e cell as a follow-up to the p reviously di scus sed APS

module tes ts. The objectives of this test se r i es were to isolate the effects ofsatu rated propellant , engine temper atur e , propel lant temperature , and thecombination of thes e var iab les on engine perform ance .

The engine firing cycle as outlined in Table I1 w a s utilized in each run,an d a 3-second firing w a s conducted immediately after each duty cycle to obtainpost-test steady state performance. The ini tia l thr ee tests used unsaturatedoxidizer at temperatures of 80" F , 140" F , an d 160" F. The last th ree testsused G H e saturated oxidizer at these same tempera tures .

The eff ect of GHe sat ura tio n on engine P c decay wa s not significant.

However, the data indicated that the saturation contributed to unstable per-formance with nominal Pc oscillations of * 5 to * 8 per cen t of rate d Pc. Alsoduring the ini t ia l pulse s , the P c value d ecrea sed to as low as 70 p s i a as g a spassed through the engine.

30


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TRW S ing le E ngine Sea Leve l Tes t ing

Another ser ies of tests was conducted to invest igate the ste p functionPc drop noted in the 501 APS engines.six days earlier, and the engines, containing the res idua l propel lants , we reexposed to the at mos phe re during the hold period.conducted by MDC, Sac ram ento , Calif orn ia, resulted in similar cas es of Pcdecay. Consequently, the test program was initiated at MSFC to determine therelation of the Pc decay phenomenon to the interactio n between propellantsremaining in the engine at shutdown, a ir mois ture , and engine mate r ial s . Insupport of t his objectiv e, hot fir ings we re conducted on single TRW engines andthe Saturn V APS module, and NzO4 flow tests w er e conducted on two TRWengine injectors.

Pre-flight bur p firings wer e conducted

"Sea level-hold" tests

The sing le engine testing is discussed in this section.

Two TRW engines were hot f i re d in a series of 16 tests at ambientpressure condit ions.tubing schematic as shown in Figure 13.engine inle ts under 20 psig blanket pr es su re durin g hold periods.duty cy cle fo r these tests consiste d of thr ee pulses: 250 milliseco nds on -500 millisec onds off, 65 milliseco nds on - 500 millisec onds off, and 65 milli-seconds on.

The engines were mounted in a "sea level" stand with the

The standardPropel lan ts were maintained at the

The initia l engine test ed (TRW SN 585) exhibited a P c of 99 psia an dO/F ratio of 1. 630. After a 5-day hold period, the engine w a s tested again,using the ident ical inlet pr es su re s as th e first run.with an O/F rat io of 1. 003. The fuel flowrate rem aine d nominal. Followingthe second test, the engine was dis asse mbl ed and inspected by the Mate rial s

Division of P&VE Laboratory. Crystal l ine ma tte r was noted at the entrance tothe twelve 0. 035-inch diam eter oxi dizer injector feed tubes. Although the tubear e a reduct ion caused by the mate r ia l did not app ear to be la rge , the effect ofthes e part ic les on the tube surface roughness in the form of increased fr ic t ionfact or and reduced flow ar ea was such that flow res tri cti on could be experienced.

The Pc value was 70 p s i a

A s e r i e s of tests wa s then begun on a second TRW engine ( S N 526). An11 per cen t reduction in Pc..was noted duri ng the init ial 6-day p erio d, and thePc d e c r e a s e d t o a low of 81 psi a fro m the i niti al valu e of 99 psia ( 18 p e r c e n t ) .Seven tests o v e r a 21-day period we re requ ired to reac h this level. A s withthe previous engine, the Pc decay was proport ional to the decrease in oxidizer

flow rate. In th i s case the O/F sh i f t was f r om I. 631 to 1. 300. Fu el flow raterem aine d constant. Fr om the low of 81 ps ia the Pc began to incr ease againand stabilize at 94 ps ia a f te r 4 e s t s i n a 14 day period. A 10-second steadystate run was then conducted and the Pc ret urn ed to the nominal value of 99 ps iaa f t e r I. 5 seconds burn t ime.

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ww

MAROTTA MODEL MV563 K VALVE -3/8 IN.NOTE: OXID IZER SYSTEM SHOWN FOR ENGINE SETUP.

FUEL SYSTEM IDENTICAL.150 LB THRUST TRW ENGINE

TRW INJECTOR

3, TO FACILITY BLEED LINE GHeSUPPLYII

10 MICRON NOMINAL FILTER+1/2 IN .

VENT VALVE

TO SCRUBBER

I26 GALLONRUN TANK

1/2 IN.

TO 3 FT. DIAMETERALTITUDE CELLPREVALVE FLOWMETER

TO APS MODULE

FIGURE 13. FACILITY SCHEMATIC FO R TRW ENGINE "SEA LEVEL" TESTSAND INJECTOR FLOW TESTS


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TRW InjectorFlowTesting

The particulate matter discovered in the oxidizer injector feed tubes inengine 585 prom pted a series of injector flow tests. The injector w a s removedfrom a TRW engine previously f i re d at MSFC n APS module per form ance test-ing, and special conne ctors we re fabricated to ma te the injector oxidizer inlet

fitt ing to the facility oxidizer supply line. The engine valves we re re place d bya single facility solenoid valve to eliminate the engine valv es variable. Nitrogentetroxide ( N 2 0 4 ) w a s then flowed through the oxidizer injec tor feed tubes ,through the injector, to the atmosphere. A protective dust cover w a s placedover the injector following each test.

A 5-test ser ies w a s conducted over a 16-day period on one injector( SN 106) , nd this injector w a s then rem oved from the stand and delivered toP&VE Laboratory for contaminant analysis. Microscop ic examination revealedthe presence of contamination at the entr ance to the oxidizer tubes in the formof a green ish, jelly-like substance.percentage of nickel and lesser perc enta ge of i ro n and chromium i n the contami-nation. Although the inje ctor w a s open to the atm osph ere for a day o r m o r eduring the la bora tory examination, liquid w a s noted in the inje ctor . The liquidw a s postulated to be nitri c acid and litm us paper tests proved it to be acidic.Photograp hs we re mad e and views of the contam inated portions of the i njec tora re shown in Figu res 14 o 17. F igure 14 shows the con tamination in theoxidizer tubes entr ance flange. Figu re 15 shows the same view of a flange thathad been cleaned. The oth er two photographs i llust rate the difference in amountof c or ro si ve attack on the front face of a n injecto r at the outlet of the oxidizertubes for two injectors, one which w a s used in this test ( Fig. 16) , and one usedin ear ly system tests ( F i g . 1 7 ) .

Spectrograp hic analysis indicated a high

A P S Module "Hold" Testing

The Saturn V Phase IV Te st APS Module w a s used for a series of teststo obtain "system'' test data on the "sea leve l firing-hold" induced Pc decayphenomenon. Initially engine numb er I w a s fir ed through a duty cycle consistingof 3 puls es of 250, 65, and 65 milliseconds duration at ambient pre ss ur e condi-tions. A P c value of 99 ps ia w a s observed. A 9-day hold period w a s incur redand the module was fired at a simulated altitude pressure of 100 00 0 feet.Engine numbe r 1 Pc decayed to 84 psia. The oth er two attitude control enginePc's w e r e within the nominal value tole ran ce, although fr om 2 to 3 ps ia low.The altitude cell remained evacuated fo r s ev er al minutes following the altituderu n, thus resu lting in the evaporation of resi dual propell ants r emai ning

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F I G U R E 15. C L E A N E D ENGINE INJECTOR OXIDIZER F L A N G E INLET


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FIGURE 16. CORROSIVE ATTACK ON INJECTOR OXIDIZERPASSAGE OUTLET

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Engine f i r in gs, even after seve ral days hold period, that fol lowed pri orsteady state f i r ing s of seve ral seconds ( I O o r m o r e ) at ambient pres sur e condi-tions, indicated no Pc deca y. Apparently th e engine bec ame hot enough du rin gthe extended run to vaporiz e any residual propel la nts .

CONCLUS ION S

The APS anomalies experienced during the SA-501 flight were:

1. High propellant tem per atu re

2. Slow cham ber pre ssu re decay of ul lage engine

3 . Chamber p re ssu re osc il l a tions

4. Chamber pressure decay

5 . Low chamber pressure.

A f t e r investigating the anomalies and perform ing a series of tests toverify the assumed causes of the problems, it was concluded that th e highpropel lant tempera ture and low chamber p re ss ur e anomalies warrantedremedial action to preclude possible future mission problems.

The testing at MSFC verified that the 501 flight APS gradual chamberpr es su re decay and oscillations, noted in the lat ter portion of the mi ssio n,w e r e a dir ect r es ul t of the high propellant te mp er at ur e. Although the highpropellant temp erat ure was in part a re sul t of the unique 501 miss ion profile(not typical f or Apollo) , flight data analysis showed that the p assive the rma lcontrol for th e APS was less than expected. Thus , a need f or additionaltherm al isolat ion of the propellant feed syste m is needed. The solution,implemented by MDC, is the addition of: ( 1) new radiat ion shield s in thepropellant manifold area inside of the module, ( 2 ) f iberg lass cove rs overthe servi cing disconnects located externally on the aft end of the module toshield them from d irect sol ar heating, and ( 3 ) new brackets that attach thepropel lant l ines to the pr imary s truc ture which will ser ve as the rmal insu la torsto prevent conduction of heat from the st ru ct ur e into the propellant lines. Thedesign changes we re based upon the re su lt s of analytical studies performe d byMDC and should correct the problem.

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. . .

The low APS chamber pre ssu res observed throughout the 5 0 1 flightrep res ent d egradation of engine performance, although possibly one the systemcould endure without serio us miss ion consequences. Thi s performa ncedegradatio n ha s been repro duced i n post flight ground test and found to be ares ult of clogging in the oxi dizer flow pass ages . It appears to be a functio n ofthe number of ti me s that the engine is burp fired , the period of exposure to sealevel conditions, and othe r facto rs such as temp erat ure of the hardware,rela tive humidity of the a ir , orientation of the engine, and vacuum exposure.The degr ee to which each independent fact or affects the engine performa nce isunknown and would require a significant amount of testing in a str ingent lycontrolled environment to dete rmin e. A s stated in previous sections, theproblem ha s been experienced in other sys tem s and the effect of a combinationof N z 0 4 and wat er on Inconel ha s been studied in the past. However, becauseno s im il ar anom alies had been experienced in the S-IB Flights o r ground te st s ,the problem was not considered to be ser ious . The reason for this inherentproblem not being identified earlier i s probably twofold.fo r forming the obstruct ing residue were negated during most s ystem testsand single engine tests by eit her the fact that en gines we re purged with GN,following test and/or the test firin g cycle was such that sufficient heat w asgenerated within the engine to vapori ze all resi dual propellan ts. Secondly,t h e r e is evidence that the obstructing resid ue is relatively soft and easilywashed out at sea leve l conditions; howeve r, und er vacuum conditions th isresi due may become crysta lline and extr emel y difficult to dislodge. The secondreason is considere d the possible key as to why the low chamber p res sur eproblem was peculiar to the Saturn V and not experienc ed during Saturn IBflights.V compared to 2 minutes fo r Saturn IB.by the AP S pr ior to initial f i r ing i n f l ight i s different for Saturn IB.resu l ts to determine the effect of th ese vslriab les on the clogging phenomenon a r enot yet available, but testing effort is underway.

Firstly, conditions

The time line fr om lift-off to in itial AP S f i r ing i s 8 minutes for SaturnIn addition, the level of vacuum seen

Te s t

The conclusion is that the low chamber pre ssu re was caused by chemicalcontamination; specific ally, reac tion s between nitric acid and engine inje ctorhard ware caused the oxidizer flow passages to be clogged. The nitr ic acidwas formed by allowing the ox idizer to co me in contact with mois ture in theambient air on the launch pad.probable cau se of the flight anomaly.

The test ing at MSFC has ver if i ed th i s to be the

The solution to the prob lem is to not expose the engine inje ctor to N204

and wateq and to not hold at sea level condition fo r sustain ed period s of tim e(days) . The method to accomplish thi s is the deletion of the burp firing re-quirement before launch. This is being implemented and should eliminate

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fur the r prob lems of this nature.checkout, loading, a nd launch pro ced ure s so that the deletion of the burp f iri ngdoes not cause any significant dec rea se in confidence that the syste m is readyfo r flight.

Enough confidence ha s been gained in the

The other anomaly was the slow chambe r pre ss ur e decay of the ullageengine. The probable cause was attributed to a sm al l par tic le of contaminationimbedded in the engine valve seat. This is considered to be a random occur-ren ce, and only caref ul adherence to cleanliness requ irem ents would be any aidin eliminating the problem. N o effect on the mis sio n w a s noted due t o theoccurrence, and it is not conside red to be worthy of any fur th er study except towatch for any future oc currences.

In conclusion, it is cons ider ed that with the modification of h ard war e

F u r t h e rand launch proc edu res being implemen ted, that the SA-50i APS flight anomali eshave been identified and their caus es eliminated fo r futur e flights.tes t ing to provide a mo re definitive explanation to som e of the assum ptions isstill in pro gr es s and should be useful in futur e designs.

George C. Mar shall Space Flight CenterNational Aerona utics and Space Administration

Mars hall Space Flight Cent er, Alabama, J anua ry 1969904-24-03-0000-50-529

42

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NASA-Langley, 1969 - 8 M 3 5 5

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Sa 501 s Ivb Auxiliary