r ,

•

~ORM AL •IS 10 M 11 · 41

U.J,B f.'~ II GALCIT REPORT NO. ___ _

GUGGENHEIM AERONAUTICS LABORATORY

CALIFORNIA INSTITUTE OF TECHNOLOGY

AIR CORPS JET PROPULSION RESEARCH

GALCIT Project No. 1

Report No. 24

WITH SPONTANEOUSLY IGNITING LIQUID PROPELLANTS

.w. :9. Powell N. Kaplan

California Institu~e of Technology July 14, 1944

Copy Uo .

AIR CORPS JET .FROI'ULSIQ:.:; RESEARCH

Report No. 2t~

for the

AA:F MATERIEL CEUTER AIRCRAFT L.A.BOR.!_\.TORY

;;Rzs .. .,..,..6 1 &> 1 hi!i f 1f"'c••rr, • ("·"': ;.,5 ·o:--mltPnn .'\lfTL• •. :rinll 1~..::1 l, .. :-.: ,•· v L • ,. -:: n' t-lC 1 In h•-.I Sc,lttlS,

~--t~Jt:. ~!~~· 1 • -~:(· ~ 'r~ 1

11

r,~n·-~~~:~~¥:,'.~~~t ~~; .r-evcb.c,~,.-.n (· <tii •t.~llfCJI( rn .lt)',' JnoHH\Ct to an wuutho.ti~W peu.:tn Lb prohibit~d by law,.

Claultlcatlon Ch•naed to

UHClASSffJ£0 Aulhorlty

JPL Pu~lkauo

ROCKET l.~OTOR ST.A.RTDTG CH.ARA.CTERISTIOS

WITH SPOlfl'JJIEOUSLY lGNITIUG LIQPID PROPELLANTS

W. B. Powell

U. Kaplan

Chie En~ineer, Air Corps Jet Pro11Ulsion Research Project

"~"" JIJ. Project :, . 1 RES\fRlGf¥0 c ... ' . 1:o . 21~ (Jdr. Lo.b.)

I.

II.

III.

IV.

v.

VI.

•

TABLE OF COiiTEUTS

Introduction and Summary • ••••.....•• • .•••..•...• • ••• • ••

Discussion of' the Problem ...•..............•..•..•.••..

Experimental Data on Ignition Characteristics

Experimental Data on Stnrtins Characteristics

Theoretical Analysis of Startin,~ Prcssur3 Surge and Allowable Initial Rates of Flow .•................•.•.••

Discrussion of Results ................................. .

Page

1

4

7

10

13

16

References • . • • . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . • . . . . . . . . . • • 19

Table I ....................................... ' .............. . 20

Figures 22

'I

GALC I T Project ~o . 1 RESTRICTED PD<:e i i

Re ... . 1;c . 2'' (~\~r . L~.J . )

LIST O]f FIGURES Paee

FiQJ.re

1 . Throttle valve asnenbly drawinG •......••..••.... . •...•.• 22

2 . Section of throttle valve .... . . . .... .. ... . .. . ..... . ..•.. 23

3· A~~aratus for cu~ ir,nition tests ... • . ... . .... .. . ...... .. 24

1: . Cup iGnition test of 6-~~ RF!lA and aniline . . . . . . . . . . . . . . . 25

5·

6.

Open injector ie;nitiou test setup .. ... . .. . .. . . .. .......

Open injector i~nition te$t ..... .. ..... ....... .. .... .. .. 26

27

7· Injector wit h open cylindrica l chamber . . ... . ..... . ...... 2g

S. I e;ni tion test witn open chamber •....••. .'. ...... . .. . ..... 29

9. Dasic cirauit diagr am with by- pass startine system ...•.• 30

10 . 1000 lb thrust rocket motor of conventional proportions • 31

• 11 . Rota.tin": cylindrical plun" valve • . . . . . . . . . . . • • • . . . . • • • • • • 32

12. Assa~bly of ~lu:, valves with electric mote~ drive • . ...•• 33

G_:U,CIT Project :Uo . ~ · ::, ''o:dlU!TVt~ Rep . Uo . 21~ (Air . Lab . ) •

I • I NTRODUCTI CU AND SUI.ThiA.RY

The startin~ characteri s t i cs of rocket mrtors using a spontaneously

i'"?litin"' liqu.icl propellant have been investi P,ated. and tentative equip-

me'1t and operation specifications for stnrts wi t hout danger ou.s initial

pressure sur~es are propos ed. This report pr.esents and summarizet? the

ex:)erimental ~d theoretic9l datn whicl serve as a basis ~or the speci-

ficnti ons . Th~ worl: indicates that nr. d::>.ngerous initial pressure surges

(hard stnrts) Tiill occur if the ini ti?..l flo':'l of propellant into the

c<"' ~oustion ch;mbcr is controlled witr.in certr-in prescribed li'Tii ts .

The e:x:perimental studies of the pro_1osed st.'"'lrting t echnique were

performed with the RFJ.T.A- a.niline prc:::ellant. and the rocket motors used

had h charn.cteristic len,~th, L*, ra.nginp; from 45 in to 97 i n and were

desi";ned to e;ive thrusts of 200 to 1000 lb Tihen operated nt chrunber

"Jressures of the narnitudn of 300 tc 550 psin. .

On the ba~is of observations of Ol.)en cup tests and open in,jector

tests, s~ontaneous i~ition of ~TA-aniline mixtures starts in a rela-

tively stagnant bo~v of the mixed liquid pre )ellnnt co~Jonents . Ap-

parently, o.. rel<-tively slm reaction takes :;_)lace in the liqui<l :::}1aze

until the tenperature is reached at whch flA.me ap~1ec..rs; then rePction

proceeds with :1 ratA approacl ing thr-;t of nn explosion. Tl!e ti'lle l ag

from initial injection ancl nixin<-; of the liquid propellant COT'll?onents

until the combustion 11rocess is established has been found to d.epend

en factors d1ich aff'ect the temperature rise of the mixed liquids,

na":lel:r:

l. The rate of hen.t generation, which depends on the propel l?..nt com-

.._'onents and the nixture r.e.tio of the st'lrtinc propellent flow .

U~G Slfl

GALCIT P::-clect Uo. 1 Rep. no. 21.! (Air. L?.b.)

Pace 2

2. The ra.te n.t uhich heat is abstracted from the mixed liquids to the

co!"lbustion cbn.mber walls or the atmosphere . This depends on the I

I

velocity of "-be liquid jets, the de:;ree of at:;itation and atomiz~

tio~ of the mixture, and. the geooetry 'of the co"lbusti on chamber.

3. The tine th~t the o.ccumulated pr opellant can lie in the combustion

chamber before being washed out by the starting flow.

If the time in tre chamber is not long enough for the necessrurr ignition

to be attained at the net heatinc rate obtained fron items (1 ) and (2)

above, irnition r.ill not occur in the chron1H'r . If an excessive qu.."Ultity

of liquid propellant collects in. the chmnber before i~ni tion occurs,

then the reaction may reach the "explosive" stage and a hard start will

result .

In order to lini t the qu..'lllti t·,· of propollP...nt ,-,.hich can collect in

the chronber.during the time intervnl from initial injection to i~ition,

it is necessa.r;; to li"ii t tJe rate of flo·.- of propelle.nt into the r.:otor

durinG the startinG period. It has been found that e,ood starts can be

obtained in rocket meters of conventional design with L• from 45 to 97

in and usine; the 6-~ RFl~A.-aniline propellant when the following condi-

tions are satisfied:

wi l. The ratio of i!litial rnte of flou to chamber volume = . 002 Vc

to • 008 lb/ sec per cubic inch of conbust.ion chamber volume .

2 . ~1e ratio of initial rate cf flow to exhaust nozzle throat area

wi = .20 to . 70 lb/ sec,per square inch of exhaust nozzle throat ft area,

3. The mixture ratio ri of the initial propellc>.nt flow is 3.0 to 3· 5

In this rOll[;e, icni tion is Most proMpt, tho~h stf\rts hnvo been

made at mixture ratios from 1.4 to 5· 5·

-----.-....-..... ~

GALCIT Pro,iect ro. 1 Pa.""e 3 Re , • 1: o • 2' t (Ai r . k't b • )

After a stc:.rt has been na.rle i n this mennor, the trens~ tion to full flow

should take place gradually durill.b a pericd of at least 1 to 2 seconds,

o.nd only after co;nbustion at t he initial rate of flo\'.' hf'.S proceeded for

a fe\7 seconds n.nd the inner P-all of the co1"1"bustion chamber has warmed

u:p.

The use of different ~ro~ellants or a change in tho basic design

of the rocket motors will make additional investigation necessar;r to

determine the opti·:ru.m values of the quanti ties Wi , Wi , and r 1 , but vc ft

the present inform~tion should serve as a [;uide in conducting such ex-

perinental \.'Ork.

It has been found that quantities of oxidizer, fuel, or water

\7hich ma_v lie initially in the combustion chanber do not affect the

nature of the start if the above limits are observed. SucJ-.1 conditions

~i<:l~t occur in practice when the propellant tLrottle valves are located

so·::e rli starrce from the rocl:et motor and the volumes in the circuit be-

t;reen the valves and the motor are large or unequal, >7hen leaky valves

are encountered, or when motors are washed down between runs .

~1e experioents described demonstrate t hat control of the initial

propellant injection rate is the key to the ::;roblem of hard starts .

The by-pass syster:1 of controllinc; the starting r;o.te of flo\'7 and rni:A:ture

ratio has been used extensively, and no hard starts have been obtained

v;hen conditions were '~e1)t v:ithin the Sj_)ecificn.tions given. Over 280

~otor runs ih several different test setUps have been successfully

started through tLe use of this system.

GAL~IT Pro '"'ct :To. 1 Rep . ::o . ?.1 (\ir. h'\b.)

II. DI 3CUSSIOJ! OF TIS PR03L~:

~he ~ ri''4'\:-" concern in thG investi -:ation of SJOntaneous liquiC. pro-

pell.:mts for rocket 'lotors hnd been the d.eter:nination of their ste~~·

st":.Je 2erfo1"'nance cha.rncteristics . Developments of rocket notors nnd

o.uxilial"j' equipnent have been c'hiefl:r motivated b~· the need for better

and rwre complete perfoma.'1.ce data on vn.rious propellant combinations.

The first motors were sturdil~r constructed nnd aole to stand r.ru.ch

necha.nica.l abuse, but thev vre:::-e uncoo l ed, anG. th"' orucirmn duration c-"

a sin~le ~'1. Tias li ited to a~~roxinately 25 seconds. ~1e ,ro~ellant

control· vclve::; v.rere designed. to throttle the i:1itial propellant flo•.;,

but tite ar:wunt cf the thret tlin2: and the O::?eninr: tine of tl-.e v.'l.lve

-'3l'B n:'t d0ter:1i'"',.,d il' n~corcln.nce 1-i.th all'• consistent scher.1e .

Tests ccnductcli vri th this ·br:re of equl~ment and vri th t1:a c:::-ir;innl

lac:: of 1.4.nderstan"linl: 1;e:.·e _ unctuated b-r hard and sometimes destructive

starts. It ap~eo.red tlk'\t when laree quD~tities of propellant collected

b~ the time i....--,i t::.or. occurred, the rapid renction could cause n vory

1 it;l ~resm1.ro surt;c in the co!"lbustion chr:.''lb!Jr of the conventional liquid

propellant rccke~ I:Jotor des~1i te the relatively le..r;;e OIJeninco: ;;resented

by tho exhai.Lst nozzle . Uo instrur.:1ento.ticn vto.s available nhich v:r>.s ca-

pablo of resolvinc trcmsients of .this nature; thus, tLe ilat;nitude e..nd

rate of ti e st£'..rtinc pressl.A.r~ ~ru.1":e are unknovm. Starts are ,jud{;ed by

ex.;_Jeriencerl observers on tl.e basis of their audible imression, and

ranee from a ~ra~ual bu;_ld-up of pressure to a violentlv destructive

hare start.

T11is sto.te ·of affairs existeci. for some ti~1e , .. hile em!'hasis remained

on soluti0.n 0f t~e probl~11s of stendv stat~ perfor.nance of v arious pro-

pellants and rocl~et P.!Otors . ~e _preble.., of hard starts bec::..."le serious

GALCI'I' ProJect ITo. 1 ?-e9. :·To . 24 (Air. Lab.)

P~e 5

r:hen it was no lont;er permissible to build motors strcn{; enou<:h ~d

l:.eavy enoU;:~h to "1 thstPnd excessive startinG pressure surges, <L"ld S:::>Or-

adic attempts were made to determine quantitatively the nature of the re-

quirements for a smooth start and for a smooth trDilsi tion from the

startin{" _;:>hase to the final equilibrium rate of propellant flow.

The experience up to the time that this stud,y was started had been

that some test setups operated satisfactorily, uhereas others uere

plagued v1i th recurrine; st-..rti:n..r; troubles. These troubles could not be

traceO. definitely to the propellant control valves ( Cf Figs 1 and 2 for

the t~'})e of valve commonly used for motors of up to 1000 lb thrust)

because there was no other device available which would control the pro-

pella.nt flow in a knovm manner so that a com.LJarison of starting perform-

ance could be made. The proble!"l was confusecl b-,. the fact that ~od

starts were obtained in some test setnns even v;hen no atter:1I't was made.

to throttle the initial :prOJ.Jellant flow, thou~h analysis showed that

there was sometimes effective throttlin~ present beceuse o1 the time

required to build up to :t'ull feed pressure in the l)rO "'ellant tanks un-

der certain conditions of operation. Injectors {~i vint' different spra_.y

patterns see;-;Jed to have sor:1e effect on t}e n8.ture of the start, and

variations in the water content of the c.cid type oxidizer were like-

wise sus1Jected of influencing the starts.

Very large scale notors of 2000 nnd 6000 lb thrust were operated

repeatecU~r by applying full feed pressure to tht. pro~)ellant and then

actuati~~ quick openine pro~ellant control valves . The starts ob-

tained in this manner were quite abru')t but never destructive (Cf

Ref 1) .

.Motors givine a nominal thrust of 1000 lb were operated in the

-----...... ,.. ....... ~

GALCL Pro'ect tTo. 1 Rep. :o . 2' (.U:·. L:l.u.)

I

Pa.~e 6

SDT.le nanner, but vcr;t often the starts ..-ere noticenbly hard, and sometimes

do~tructive starts wore encountered. Satisfactory starl;s were · obt?,i.ned

L.~ the i.ni ti:ll pro11ellaut flow rate was reduced. by throttlinc cr b~r ap-

pl;vin, e. lo':"t initial feed pressure to the propellant tank::> , 'but there

vras no means of deternini.n.r; quanti k .. tively the ma.r:ni·tuc:.e of the ini ti.al

rate of flo' or tho duration of the throttli.~~ effect .

The experience with 1:10tors of sr.w.ller than 1000 1 'b thrust vm.s quite

rimilar to that, clescribed above; some test setups workad satisfactorily

and others gave troubl&. However, the uncooled motors of approxi~ately

200 lb thrust which were tested were so stur~r that destructive sk.rts

The devt?lopment of lie;ht·.-;-e i e;ht coolecl rr.oto!'s mf'.d.e it necesscu:r- to

develop startinr-; teclmiqu.es which "rrould not subject the r:1otors to the

hir;J.1 ::_Jressure su1·t:;es associn.tecc mth the hare'. start.

The study of the harc1 start problen VI:'LS seva.ratod intc t'\70 r a:r·ts'

1. The :9henonena. of it;nition of span .nnoous liq..:id propelln.nts

and the fn.ctors dfocting the ignition tiue.

2. The .;henomona ncconpan:ring Sl10:'ltanoous ignition of liquid

propellants in the combustion cl~~.1bers OJ. roc;::et motors .

The first phase of t~1e problem wo.s investicatod by me;:ms of various ·

kinds of cup tosts, open in,iector tests, '2nd open ch.3:.1ber tests (no

ex!m.ust nozzle) . The second phase of tr1e 9roblom ::;ee:ne:i to be related

practically to the rate of tl1e ini tin.l propellPnt in,iection i nto the

coDbustion Cl.1a.r'lber . T~1.erefcre, an attoJ:!pt was made to start several

motor.> with a controll0d. initir,l rate of flO\': and to establish a. cor-

relation betveen \'li · - , and. ri and the nature of the start and

G.-\..LO I:' F:.·'"' _; ; r -': He . l Po_J. Ik . 21, (Air. L."J . )

CC' """'I 'JE'":'I t\L

and in several test setups at tl1e ACJP Pro~ject .

III. E.JCPERI.-:::EKTAL DATA Ol! I GUITICH CIIA..P.ACTEPI STICS

A. O~en Cup Tests

Pn. ·o 7

A cuu test (Of Ref 2) affords a ~relimina~r pictur e of the ig-

nition characteristics of a s2ontaneously ic,nitinc ~ropellant .

:Phis test, in '1'7hict a qudnti ty of one ::?rO._)ellant com·)onent is sud-

denly ;cured into tLe other cooponent in an cr_)en vessel, does not

sinrt.ll ate tne ir1rJinginG of high velocity !;itrem:Js. and therefore

pives only em a,?proximate indicF.:.tion of t~ e ignition YJrocess . The

cu:) tests 1·erc ma~.e .wi tl• the appare.tus shorm in Fi"'s 3 .rmd. ~-. in

which tl"e pourin~· vrns accomplished b7 · releasin--; a neight- loaded

~m.lley. In this vrD:;,r, a uniforn test ;roceclurc was insured. ;:otion

picturen at 61+ fra"'les l)er second enabled the deter::1ination of the

time la. .... beh,oen the initial mixing and the ap~')earp.nce of flame .

R.l!""!JA and aniline '-rere mixed in equn.J. volumes ( ra.Yl.ging from 10

cc to 25 cc ) and in widely dissimilar volumes (1 cc of one conpo-

nent v:it:1 10 cc of the other, and 1 cc of one mth 20 cc of the

other) . I n Rlmost all of these tests, ip;ni tion '1'7c.s proJTI!)t , the

time laG bein- of the order of 0 . 013 to 0.12 seconc.s . There was no

inciic:otion ti:u!.t one set o·r conditions (volume ratio of pro·)ellant

co-nponents and. sequence of _)curine) was Si?lificnntly better than

another . I n such ex.!:)eriments, it is probablv i~nl')ossible to es-

tablish, even approxinately, the mixture ratio existing at the

point where it~ni tion occurs .

In otber tests, one com_;onent '"a3 added dro nise to a rela-

GAL.:;L Pre. ~~t 1:~. 1 r"_). 1:"' . 2 ( .t.: .. :c. k.:a.)

00. 'I ~ ~-'LU,

tivelv 1::-r-;e pocl cf t-1'3 nt <>r co _,,..,vmt . Obseryation inlic·.te.::

a pool of acid t:.;-.!1 \rrhen drops of a.cir' ··ere :1dded to ·a ,POOl of

~miline . When drc >G of one ,.}ropcll.··!.nt .:::o!'.)Cnent ~mre ad,1ed to e.

sl.JD.llc-:-· la;,rcr of t:1c other in n. lar--;e eYn. )Or:::l:t!n..,. dish b;r <llloY .. inr:;

the dro_Js to run do•·.-1~' do;m the side of t:1.'3 cli:oh, tLorc \72¥~ nc

the mecl.£'.!l:.s:~ of spnata..11ec~s i ,e;ni tion i:wolv·~s an i1 ;tial pm~":;ial

re:>..cticn .-:l.ic'l ~8J.1.ero..t~s ~:mcu_.:; •• 11en.t to raise t!1e temperat·...:..rc of

t:1e liq',.l,ids tc tLe ;_Joint ''-'here ra. id cc·1austion ca"l. proce:;d, 1Jut

tLat, Yli th this p roce<iure , the heat t ransfer f r om the thin .!.')rO-

pellant layer to t1>e dish was sufficient to L)revent the te'npern.-

t ure of tha _n·o.:;ellant fron re<•.chin,:; t:.0 i :ni t::..on point.

A serio3 c::.~ t...,sto '.'rere x::nde on t~ .e ::;o lb ':1Ctor test unit (Cf

:ief 2) \-:it }-1 en open i·n~iv:in:";-jet in,jectoY' (C ~ Fi"'s 5 and 6). Sev-

tar:·s . I'i1~ vel:;ci t:~ of tl1e fluid s":;re:l:.IS i ssuin.r· from th~ in;tcctor

'17as controlled b:r the press1.trc ap:)liei to the fluids in the t<:>.nks.

I n tlwsG tests, it .... .-as found thc.t i3niticn 'T."l.S v9r;;r di.fficul t

to o'btn.in . I t ro.~ .r· :3'Wlled that l:.ea~ conduc:;tion fro!~ the S::Io.ll

dro_)lecs of prcj_:-el lant to ti1e air ';';".S res·)onsi'bl e for t:1e i r,ni tion

failure .

accu::ru.lati~.- on t~·.c i njector SU!Jport, conbustion persisted.

GALCE' Pro,iect No. 1 F..eJ.J. lie. 2'' (Air. Lab.)

C. Ooen Cha~ber Tests

Page 9

.Anotner serie 8 of t ests were made wi t l an open c;rlind.rical

chamber in front of the injector (Cf Fihs 7 and S ). \'lith :propel-

lont feeci :Jressures ( the pressure ci.roy across tl~e injector wts ap-

proximately equal to tho feed pressure) of t1e order of 50 psi,

ignition nas consistent and. pro:npt , niti1 no observable lag. At

very lov; feed pres:mres , ·the pro_)ellnnt issued. fro:'l the i:r.,j~c'.;or

and accumulated in the bo t tom of the c-rliJ1der, where it ir:nited

"i th probably no r.10 re 'lac than in O})e:o. cu.:_) tests . These results

are confir:necl b;r other investigators (Cr Ref 3) .

At a foed pre::; sure of ap,roximatel;}r l5v psi, the ignition

lee in the open charnber was observed to ·oe of the order of 1 or 2

seconas ( Ci Fig G) . At pressures above 15\l p si, ignition v:as verJ

uncertain, and. nt ~00 psi, ic;ni tion dicl :lot occur at all. In sone

of tHe cases, i i;,"lli tion aA>eare<l to stc~~ t in the spra-.r cloud outside

the chru.aber, and tL.e flar.Je front then oorkecl baclc to the r.:outi.1 of

the cha.r.1oer.

The onl" oovicus difference betr;ef'r ... tl.ese tests and the pre-

vious ones with t1.c unshielded injector is the ~)ossible presence

of larger cl.l'Oplot <> of propellant wLici. had fcr;ned on ti1o v;c..ll of

tLe cha11bor bcfo1·e being carried. out into the at·no~ )}!ere . 'l'hese

ln.r[;or droplets crumot lose as large a proportion of the:.:r heat to

tl!e e.tmos a1ere QS cnn the finer spraY .~:-'articl~)G issuin · fro :1 the

unshielded_ injector, and icnition _.yrobal>ly stn.rts in these lart;e '

cir0ple t s .

At all three mixture ro.tios testeei (1.5, 5 . 0, and 10.0), the

GALC:T Pro,j~ct lio. 1 Ee•J . :·:o . 2l+ (Ai~·. Lab.)

t

P~::;e 10

' results uere a~::;roxinately similar . There was slir:htly less lag

at a r.~ixture ratio of 5.0 than at the other tvro values.

IV. EXPllini.f.EHTAL DATA OU STARTING CH.ARA.CTE...,TS'PICS

In order to obtain data on rocket r.1otor starts VTith reduced initial

rate of flow, it Tias necesse.rv· to decide on a net :bod of controllinp; the

pr01;ellant flow. Two basic methods were considered: (1) to ste.rt ri th

a reduced feed :9ressure apnlied to the propellant in the tanks, And (2 )

tc start with the final va:lue of feed pressure and to have a variable

l1;v-draulic resist:;o.nce in U·e fluid circuit. It was decided to use the

latter method because it offered the possibilitv of a more direct con-

trc~- over propell[.U'lt flov:r; i.e. , vri th larr::e {jtl..S ,rolumes in tl1e r)ropel-

l3.nt tanks, t:te res JOnse to a<;l~justments in tlle feot'i. ~)ressur"'~ TI01.·.ld. be

slo,·· , and becau3e tl":e red:uced feed pressures w1,ich "iiJUlcl he.v~" to be

useCi. \'!ere so lo17 as to nake accurate cont!:·ol inP)OSsible 't':"i ti~ the exist-

i'1r!; equipment .

The r"ost r.esi rable method. of o;:_Jerati!4::; a roc1:et '1otor vr01.'~d have'

been to admit the .:_Jro:t.'ellant in the quo.ntit'r required for- a smooth

st:->rt ancl then tc increase r;raduullJr the rate of flo,., u.ntil tl1e de-

sirnd final equilibritlr.l value >Ias obtained. Thin continunusly varia-

ble t~1rottlinc process requ.ired mechanical devices ,.,},icb \"ere !tOt

available at the ti 1e tr:e test prot;r.?.m was ini ti2.ted, so a 2 stae,e

t-,;_~·ottlin<'•; s::~ste.'l was used instead. The basic circ·o.i t diagr am of the

test seh:..p is sl10Y'11 i 1 Fi<; 9. The runs vrel'e sta:c·ted b-r appl~rinr~ the

full ::."lroJellant feed )ressure to the prooell.?..nt tan1cs and O;Jeninr~ the

main p ropellant line valv•~s . The ini ti2.l flo,u into the motor Pr>S b"-

passed around the seccr:C', pair of line valves and throu•;h a pair ~f

CO:!FIDEHTIAL

GAL:::rr Frc,icct ::c. l ::.--:_) . 1:-o. 21. (Air. Lr;.iJ .)

s:Jall crifice _)lc.tt s . The initial flcvr int o ti.1a ··otor ''P.s ';Overned. by

the sizes of tr:e lwl!:>s in t:C-,e orifice nlates. T~1e cor'!lstrep;l line

valves yJere of the throttling t·r.l.)e p reviously mentioned (Cf Fi/''S 1 ano.

tw• st..,_rtin.~ value and. tiu3 final eq_uill briu.n value dul"in:-" theL:· t:trot-

tlinP; stn.ge, ;;Licrl lasteC. fro.:~ 2 tc 4 seconr].s . T:tc resulting step\'rise

control of .:)rUIJellant flet.' l;roved to be satis:f.:.ctor:" for stuctrin:: tile

startinG ru1ci tr<:' .. :Hi tion 1Jhases in tbe performrmce of liquid !)ropellant

rccket Motors .

A setu.:_J vrn..s made i n Pit 11 B 11 (Cf Ref 4) according to the above de-

scribed syste 1, anu a series of tests ,.;ere r.1n.ci.e in Y~1ici1 starts \7ere

obto..ined <.i th vn.:r,rinc initial 1ixture r<:'tios end rates of flow. The

motor useci was 6.esicne6. to gi VIS 1000 lb thrust at 300 psi a chamber

pressure and had a co~ ... oustioL chamber vohune cf 125 in3 and an exhaust

nozzlo throat aror oi c:: . o9 in2 ( Cf Fir; 9) . THe propellru1t co::l_:Jcnents

/'"";' ( ) use-:'. '\""Sre o-~7v a·~.A aHC:. aniline Cf Ref 5 . It \":as fou.TJ.d that sati sfac-

tcr.; starts could l.JB oot:~ined if the ini tin.l rate of flo,., of prOJ)Blla..'1t

'-''~~s not mor: than :010 1 o/ sec per cuuic inc •. c:f co::1bustio·· chc::.moer vol- .

-u:Je, and the initial mixture ratio was from 2 to l~ . When the ini tiDl

rate of flovr \':<as incren.sed to . 020 lb/sec p0r cuoic i....,cL of cha·oer

vclune, tne start"' oef:P.n to souncl violent . 1.'} c .eqo.lllibriurr re.te of

flon of tle r.wtor '\7D.S D-?_.:)roxir.Iately . ot~g lb/sec _)er cubic incL This ..

first se1·i'=s of t<"Gts enco,-1.!.)"-SSFHJ. a:oc1.1t e. dozen rc_7J.s, sor1e of rrhich

consisted solely of fl .. series of 5 or 6 st£>.rts, and. was conducted fro·J

Jc-.nuary ll to Januar.r 15. 1)1~4.

On the b<'.SlS o:i: tiHs resu.lts out:-.ine:J_ f'ro .. l tl.e abov~ tests, b-~-pass

.,,...... 1',.-,T~ ,..,..l'MT A Y

GA.LCIT Pro,iect Uo. 1 Rep. No . 2l+ (Air. Lab.)

Page 12

starting syste:-1s were insta.lled at the .ACJP Project in Pits ".A" and 110 11

(Cf Ref 4) with completely satisfactor;:~ results .

In beth Pit 11.A11 and Pit 11:811 , some tests ware made in which a qu.an-

tity of acid or aniline was poured into the combustion chaJ!lber to simu-

late the unbalanced. condition which might arise when onl:~ one prol)el-

lant had tc run throu~h a coolins coil bet~een the propellant valves

and the injector . The starts were completely unaffected by this pro-

cedure.

furing the last half of !.'arch, 1944, a "'lore thoroue;h investic;ation

was made in Pit 11 B11 usinf!: two uncooled motors . One of· these motors bad

a combustion ch~ber volune of 121 in3 and an exhaust nozzle throat

area of 1. 25 in2 ; t hus L 111 was 97 in; the otner motor had a combustion

cham~r volume of 30 . 5 in3 and an exhaust nozzle throat are~ of .407 in2 ,

... so that L was 75 in. .Approximately 50 starts and transitions were

made with these t>ro motors uith varying initial rates bf flow and mix-

ture ratios. U1e tests covered a range of values of initial rates of

propellant a~"'lission of Wi = .004 to . 009 lb/sec per cubic inch of Vc w·

combustion chaober volume and a range of -h a .10 to .68 lb/sec per ft

square incl" of exhaust nozzle throat area. The ini tietl mixture ratios

testeQ varied from 1.5 to 5 · 5 · The transition to full thrust was made

by means of a throttle valve which o~ened in approximately 3 seconds,

and whose throttling passaee w~s such that the rate of flow during

transition was a9proximately one third the equilibri~~ rate of flow of

the larger motor . Starts were made wi t h acid, aniline, or water in thJ noted • combustion chanber, and no effect on the nature of the start VJas

.A motor controlled throttle valve was constructed in crder to

CONFIDENTIAL

GALCIT Fro,joct llo. 1 .i.(~ () . 1To. 2l! (Air. Lo.l), )

I PCl.CO 13

I

I

oo~ai11 a e;ratual transition fro:-1 the stRrtinr.; 1lov.r to the final eqo.ili-

briu':l rate of .:.JrO.:_)e1lo.nt flo-rr . The valve consisted of a pair of C'"lin-

drical _;_)11l5 valves mounted on a base with a geared e l ectric motor· drive

( Of Fies 11 anJ. 12 ). The. motor \7aS remotely controll ed so that the

vPlves could be opened 8..Ild closed or sto~ped in a.n;-r intermediate :1osi-

tion. T.bese throttle valves \".'ere installed in Pit "B" in a setu~) for

testing 1000 lb thrust notors, and: a careful flow calibration of the en-

tire circuit was made with water so that tLe rate of flon of propellant

at a ~iven feed pre~sure and valve opening could be com~uted accurately~

Beca:usP of t~.c s;r:mnetrical design of the valve, it was r1ost convenient

to str>.rt with mixture ratios of ayproxir!lately from l. 3 or l. 4, but sv.c-

cessi'nl starts were made with = . 012 and . 0034, the latter being

rro.ch the smoother . Trnnsi tions with this equipment were ver-;1 un.iform,

as the valve required a_1proximately 6 seconds to open frol:l tho throttling

to the full open )Osition. where VI -'VIas equal t o· • o48.

:By- pass t~ rottlin'; systems for ·starting liqnid propellant l:lotors

were installed. in Pits "D11 and. 11 F 11 at fue conclusion of the previously

described tests. At t:l: .H present time, all liquid pro:pelle~.nt test stands

are so equi!)ped, and, in all, over 280 starts have been made without

ex..:_)eriencinr; a.1'ly difficulty.

V. THEORETI CAL .ANALYSIS OF ST~TI:IG PRESSURE

SURGE .A!ID ALLOWABLE IlJI TI AIJ PAT'SS Oii' FLOW

.An approxinate calculation can be made to indicate tl' e ma,~i tude of

the pres3UI'e sure;o which 'llight occur in t • e chamber of a liquid nropel-

la.11t roc"cet motor. The basic assu:e~ptions are as follows :

1 . The tota;t Tieight of accur:n1lated pro_?ellants is converted to

~";aS, ancl the perfect gas la'\'is hold .

GALCil' Prc1e:!t l!"o. 1 Rc,7 . l·c. 21 ~ (Air. L.~".)

2 . T'r..e reactir,n cc::mrs inst<'..rtta'1eously, anC. the 1Jressill"o croatact

in the ;:;;ctcr c}u·~"llbor is the s :10 ns that wl.ich cculd be createcl

~he ~ressure create~ in ~ closn1 vessel is :

R ' w .,I.. : - 7(; )( -r M Vc

\7here W is the tctu.l 'l"'ei~! t of "".Ccm:n.1.lated "'ropellant and the

other quantities are d.efi11cd in Table I.

I:' the :per.fcr~1~nce chc...:-~teristics cf tLe prcpelln.nt are knci'l!l, the

follcv:ing GU.bstitatlcn c~n be ::~n.d.e (Cf defil1ition of c• in TablP. I):

so

R' -Tc M

the ::;-u.rc.;c :)re5su.re, P i , is not to be exceeded is :

is tl:e it·niHon ~i··e lnt; fror.1 th<? instrmt of injection.

Finally, tho e:ltl)rossion l'or CO..'l bo obtrtined:

Pi 5

GALCIT Pro.iect l:o . 1 Rep. No . 2h (Air. Lab .)

Cr::-T"'IDENTIAL Page 15

~ne equilibriu~ pressur e in the chamber with the initial rate of flo~,

It is apparent that W; : 1. ff W. 7i v;-

* from the definition of L in Table I. Considering a ,conventional liquid

propellant rocket motor and sub5titutine in the above equations, the

followi~~ r esults are obtained:

Pi • 300 psia

L* = 50 in

c* • 4500 ft/sec

... 3200 ft/sec c i -• 1 I r = .658

Vc = 132.8 in3

ft • 2. 65 in2

ti = 1/8 sec

The value of .pi chosen is the desi~ chanber pressure of the motor ,

~i the value of t1 is that indicated by a great number of cup tests. Then:

wi: ~· 9 _ 30o x .32.2" a = 7. 16. ~ c'*ar~·- 4.500;.~12 .OO '.3SS~C.X111 3

Wi' -­... .It

' C'O\,.. .6..., .. -'

.. . W"/ '; lh 1.. )( ~ =SO x.007.3S =. 368 4~c-. x 1n 2

. ~· ~ ....

GALCIT Pro.icct no . . 1 Rep. Uo. 2h ( Air. Lob.)

Pc· I.

Wi - ~~ X

P96e 16

C•,. = · 368 x ~200 = 36.8 psia

g :32.2

VI . . DISCUSSI ON OF !1.1-JSULTS

1 The preliminar;:v tests with propellants i n open beakers and with the

shielded and unshielded ·jets of the 50 lb TJctor injector were intended

to give o. rough _:1 icture of ·the ignition .henomenon . The conclusion

drawn froM this study, as stated in Pa1·t I , was that icLi t ion first

occurred. in a relat ively st~nn...~t body of mixed i!ropella.nt comnonents,

aml that i[0itiOJl Was de1a_ved \'ihen heat \7U.S abstracted. fr0r.1 the liquids

in an;-r nanner. T"nu.s, a :fine sp rny in t ho air vrould ~ot iGili te, and i r;-

ni tion woul9. not occur in propellant d.ro·)s which Tiere TJoving at hi,o;h

veloci t:• in contact with a :netal v1nll . lliese observations IJOinted to

the use of the nor.:~c.l .mu.lti- orifice in.jector with greatly reduced. rate

of floi7 as a oecns of startillb liquid 1)ropellant rocket motors .

The tests vii tlt motors of from 200 to 1000 lb thrust described in

Part IV confhnSJd L1a belie1' tl.at good start s cnulct be obtained by re-

stn.c t inc t l1e initial rate of flow i n motors of conventional design .

The ra..11ge of values of the startine flow given in Part ·I , VTi • . 002 to Vc

. Ouf5 , is nell \Yitbin the ranee where successful starts were made . A

conservative up:per licit is r;iven, s i nce a variety of factors could en-

ter to increa~e tne icnition time of :1. pro:?ellant ( iMpUre propellant

· com~onents, or i n,j ection in SU:ch a ma.nne:r.· that her:.t transfer is increased) ,

\

and thus the quo.nti t ·- of pro~ellant .:_Jresent rrhen ic;ni tion occurs mi,:;b.t

be increased. The loner limit is deterninacl by tae dele:; mlich CM bo

nn'D""'l':1Tn~'nnT .t.Y

I \ I

•

G.ALOIT Fro<1e~t l~o . 1 Rep. li0. 24 (.Air. Lab . )

CO'lFI DZllTI.A.L Pace 17

tolerated betneen o_!?erating the startin~ controls .:md the B.,!)pear~ce of

flame and by the lo 7 chexlber pressure of the resUJ. tin~ initial combustion.

The successful tests with large scale motors (Cf Ref 1) indicate that

the overall ig:ni tion ti~e in a motor ch..""Unber is decreased slie;htly and

reoains moro constant when large qunntities of propellant ~re present.

The theoretical calculations of the initial rates of f'lo'cr are in

close o,r~reement ui th the lini ts found des~ ruble by actual test stn.rts .

This acreement mS\Y' indice.te either of the follorlng:

1. ~ue assunptions r.ade in the the theoretical analysis are cor-

rect : that the iGnition, once sterted, proceeds at an extremely

ra~id rate anQ that iGnition tine cnn be satisfacto~ily esti-

natccl fran cup tests .

2. Larger accumulations of propellnnt thnn ~he mnxicrum indicated . .

b:'r tl:e previous calculations can ei ve sati sfnctorJ stn.rts if,

becCll.se of ?Oor r:aixinc or excessive heat tr.:msfer, part of the

1)ro:pellru1t is not involved in the initial rapid reaction.

The calculations ru1d recommended lini ts on startinr flow include a

"\7-i specification cf f- • This quantity is not direct~r conce1~od with_

t starting perfor~ance of motors of conventional proportions, but exerts

nn influence on the rn::;>idi ty with VThich tre.nsi tion to full thrust can be

r:aac1 e. TrL>.llsi tion involves an increa.se in the propellant rate of flow,

and the ea::.c m th which the additional propellant is ignited depends on

the intensity of the initi=:~ combustion in the chamber. The qtUmtity

ni ft dete~ines the initinl chamber pressure and, thUs, the intensity of

. wi tne initial reaction. Thus, 1'7hen ft is near the l0\7er limit, the

transition to full flov; r;ust be made -:raduall;;r over a period of at least

CONFIDEN'!'I.A.L

GJ'..L'JI~ Prc.i..;ct 1~0 . 1 F.e_) . He. 2'• (Air. LC'b.)

2 seconn:3 in order tc o.vcic a hich prnssuro sur.o;e. With

Par,c 1S5

near ita

Up!JOr li'C!it, a largo number of transitions havr been made Tiitl: no attempt

at throttlinr; and '171 th no indication of a hi •;h _:)res sure surGe , though

the abruptness o:: the tre.nsi tion may have pr"'Vflnted the ear fran detect-

inc n 9'..lrf'~" · So!'le throttlint; is cert r> inly d'3ni rnble 'uring the transi-

tion to full thru~t , and more is requirHd if the initial value of .ft

is low.

G·~\.LCI'.:' Project Fe . l :.c .. "' ~To . 2l :. (Air. L<;.o.)

CC'· ~I""~E" .I'IAL F:..,_;t:: 19 I

1. Powell, W. B. , "Tests of 2000 lb and 6000 lb Thrust Uncooled Liqnicl Pro;;ellant Jot !Jotors , 11 Air Corps Jet Pro 1Ulsion Research, GALCIT Pro,i act No. 1, Report No. 21, JE>.nua.r;.r 27, 194t~.

2. K.::Jllan, N., "InvestiGation of New Li qt. .... id ]Uels t'or Jet Propulsion, tt Air Corps Jet Prc_mlsion Reseerch, G.ALCI'l' Pro,ject 1::-o. 1, Progress Report No. 10 , November 16, 1943.

3. Sayv:a:rd, J. r.: . , Nary, D. E., and Sperr;r, R. L . , 11 I :nvesti.-;ation of Additives for :aa·::.- De:_)art~ent :&treau cf Aer·oneuti~J , ContrD..~t Uo. ITO a(s)- 870: InvestiGation of Se1f-I,::;nitin"" Liquic 1Hxturns, 11

Anerican ~Jnumid Co~?any Stamford Labora.toric3 , October 22, 1943.

4. ":FaciLties and Equipr.mnt of thE) Ai r Cor:..n Jflt Pro~1ulsion ::.~e3ee..rch Project," Air Corps .. Tet Proyulsion .Research, G.ALCIT Pro,iect He. 1, Progress Re1)ort He . 4, I.ley 28, 1943.

:; . Uruir_m, F. J., Poy;ell , Vi . B; , and Kaplan , N.; "An Invosti{;ation of a Number of Liq~i d Propellent s and a Sht~r 0f Sc~~e Effect en J et !;to tor PerforrJa.nce," Air Corps Jet Propulsion P.esearch, GALCI~ Project !To . 1, Re:oort No. 20, Oct ober 18, 1943.

CONFI DE: _ 'IA.L

I •

GALCIT PROJECT NO. I REPORT NO. 24 f A ll? L AB )

Cy

dc de ~ F=; XC

re ;f g .-I.e L~-K -;; ,Pc

Pe ;Oo

I? Wo r=-Uf'

PAGE 20

~ONFii'ENTIA:.

TABLE I.

NC?.lEN~ LA TuHE

Aspe c t r a tio of cylindr~cal c ombustion c~ber

Eff ective exhauet velocity

Charec ta r isti c veloc ity

Exhaust nozzl e dis charge cnetficient

Exhaust no zzle thrust coefficient

Specific he ~ t o f gns at c onsta nt pr es Gur e

Specific he fi t of 6 RB at c on et ~nt t emperatur e

) 1amete r of cylindrical c cmbus tlon chamber

Exhaus t ~o zzle exit dia me te r I

Exhaust no zzl e t hroat di ameter

rhrust

·Exhaus t no zz.le e xl t a rea

• Exhaus t no zzl e throat a r ea

Ac cel era tion due t o grav i ty

Le~gth o f cylindrical combustion chamber

~harEL: teri st1 c l en0 th o f combue't1on chamber

Gt.araber pr esst.re (a b s olute)

Exi t prees~re (a bcolute) ,.

External 1 ree 3ure ( abs olute )

Ga s cor.s ... a nt

Pr.•pel Jgnt ~1xture r~ tio ;-JD -114 CO~FIDENTIAL

--

G~LCIT PROJECT NO.I REPOR T NO. 24 (AIR CORP. )

t

Vc

v

v w

Wr Wa

WSp= ~ = !' 0(

c 7 ==- ::P-

Cy

A= j-(i -rcos ex) ~/-L ~- ;0

,P=t cr

PAGE 2 /

TABLE I . ( CONT.)

.J_r"T.J.aruic press ure of f l uid ,;. low

Heat flow per unit a r e a

Temper< ture of gases 1 .. cor.-.bustlon cha.nl:er

(absolute )

Time

Volume of combustion chamber, UD to the

throat of t he exhaust no zzle

Velocity

Speci!1.o volume

Rate of propellant oor. sum~tlon (to t al '

Rate of f uel coneum~tton

Rate of oxidizer c~ nsu~ptio~

Specific pro pel lant ~~nsumptl~n

Half angle of e xhaus t nozzle e:JCJ:·'lna.ing s ection

Half an~le o f e xr~us t no zzl e entrance . sP. 0t1~n . J ireotion of r e s ultant ruomestum s.fter

1Lj ertor jet i wpi ngement

lieig,ht a~naity

Exhaust nozz~e area r ntio ~ (r-rl) / 2 )E{9:.~ ff C?'?'-/7

Ratio of s pecific t.e a t e o f ga·a

Sxhau e..t noz zl e C.Lvergence a ngle func ti~n l

Kinema tic v iscosi ty

!>lase 5.ens 1ty

Spe:rl.. f .i o r: ravity relative to '.;ate r a t 4° c. and 1~. 7 ps 1.

Abs~lut~ v i s c os ity

OONfiDEN'I'IAL

. -;:- ..... ~ /1 T" t.• I c I ' " • : ,-.. -

'"- L---~"'"1'L 1 ~,, .•• ~ "u.t-· ··--·- · r~jr7~ CONFIDENTIAL -~ Ts-2~/2_j ----;;.:,;v ;.J-cu--- ____ J_

/J'(J/r't ... )J£/iL ;/i/45/l.f/?.5_- 4REt?

fill . \:!·L--;j ·)~s _AR~1 ~ _a_ . - - --,-~ --- - ~ -- _ j__-!1'/.) I'Y/1.51jif'_R -E/J~9

V-~,- r ,~ ~---~~~ .3NuTJ - .f-REt;_ ·- . - ....__ 8 ---

~ '

~ . . ·- .._ ._,, --:f /'t/YZ:-:1_- ~ !?~. ~~- ~-,_ ,-J _'_L,.. ':-~ VALY!"L'L. <fTE -~·~ ..e '@>: f=_,._, ~ ' '-- ... ,

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tlac> ~3t )4\ H e: -~ <H I ~ \' ,· ..... : .• ,.,.

wnh;t. ,h, ... 1een;,v :' ~ ~~· E.l ~ .1 ·~·· .'· ... -« \ -=-C.. .,_., • :tod -~ Ito ,·.,.~· ,; ... r~ t l ' :- ... ·.-r' J~· o• ll5 C'()n',·f.!' 1·1 , ~ ''', • w &II ~:1:1~ ~J ._,...

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NATE.IAL f'IH iliH I MEAT TREAT) ~AM I CHIICK8D I AP'P'ROV..O I &:NGINUR

GUGGENHEIM AMONo\UTICAL

LABORA1'011Y

CAUFORNIA IN8TITUT~ OF

TECHNOLOGY

.L/SJEAlBL .Y ::0 A1,_:r'~'; ,)l/ i t;//.

HAMil

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SCAL E I WE/6/'IT

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f~

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G.A.LC I T Pro "oct ro . 1 Rep . He . 2' (Air . L ::u .. . )

•

Pig 3 -- A:-:.pa.ratus for cup i c.ni tion test s

G.U.CI:' Pro.~ec-: _-c . 1 cc· "t'IJ:SNTI~ ~e "· ~-o . 2l~ (Air . Lao. )

Fi.:; 5 -- O~)en injector i .gni tiop. test setup

Pa.<je 26 Fic;5

G.AL:IT Project Uo . 1 Rep . No . 21! (A.:.r. Lab.)

C011FIDE.~1IAL

Fit; 6 -- O:,en injector ipd tion test, RF.l:IA and aniline at 300 ps~ feed pressure·

0

.047 sec

.313 sec

1 sec

G\LCIT Project -To . 1 Rep . r;o . 2h (Air. Lnb .)

Fig 7 -- Injector with o~en cylindrical chamber

GALCIT Project ITo . 1 Re:) . no. 21' (Air . Lab .)

I

•

Fig 8

GC:: 'I.JEI"'TI AL

I cni t ion t est wit h open chamber, 6-:--'% BFNA and aniline at 300 nsi feed nressure - ~

0

Page 29 Fi t; g

.047 sec

.313 sec

1 sec

VITf.tOGEN FEED

Plfl- SSURE REGl,L..:JTOR

N2 SUPPLY

Nl. VAL VE

MAIN PROPELLANT

OXIDIZER TANK

FUEL TA NK

DOWNSTREAM PROPELLANT VALVE

BY-PASS ORIFICE

PLATE RESTRICTION

BASIC CIRCUIT DI AGRAM WITH B Y- PASS STAHTI NG SYSTEM

:t. ~ ~h "tl,... \::)C) ::t~=; "'i

~~ ~ ~~ 1\)c:....

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REPORT NO. 2 4 (AIR. LAB.)

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,-

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/

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LET DATE CHANGE BY APP

JL - 72Sr NOZZ. ::· ASSEMB LV "'"---

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~ ~

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/ . '. . , ~-~ ..... / . ---

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.

-

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• gfJ .. L. '' A.L aAN ENE'•-K. a E.. CO ., H . '1".-JtRQ.U , • . I"A.T OPP'.

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.

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-- ---=---' A I" ART

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CO~"FI DTI"TIAL

Fig 12 -- Assembly of plue valves with electric motor drive

Par;e 33 Fi '"; 12