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Portions of this document are not fully legible due to the historical nature of some of the material. However, it is the best reproduction available from the original submission. , - -,' TT F-843"¡ ROCKET ENGINE SCT-1 /~,l ". ') LA Jc/fl.:':/,::>< /~ B~ Buitron by P. N71-71501~ '" o .<:¡ 2 ~ , --4l!fL- R) // _ ~ ~... ~(ÑASA '7 -:t (CODE) "" O ... ')- fj) ¡r' ---=:: CR 01< ~MX oRAD ~ NUMSER) Translation of !lEl Motor de Reaccion del Cohete SCT-1!1. Comunicaciones: y Transportes, No. 9, Vol. 2, pp. 49-56, Nov.-Dec. 1960 Mexico NATIONAL AERONAUTICS WASHINGTON TT F-8431 AND SPACE . ADMINISTRATION MAY 1963 I I I SCT -~ ROOKEI ENGniE The design fer the engine of the SOT-1 rocket is based on past experience with this type o~.engine, especial1y vdth the German V-2fS; we therefore decided to use 75% ethyl alcohol as fuel and liquid 02rJgen as eombustion agent. In .selecting this fuel, ,re based our study on 1're viere t:rying te achieve in the way of opt,i¡rn.rm what "fe kne,'/" and en w'.o.at results considering the high ex.haust speed of the combustion gases, ",¡hieh in turn of course ,vould give us IP..axi,'rlum altitudes. Ihe follo.",:L'1g of t;ypiea1 fue1 eharaoteristics ,,¡hieh have been usad table is an analysis L-rJ.experiments with recket engi..1'1es. TBEORETICAL EX.4UST SPEEJ) OF COMBUSTIOt'I GASES ll1 m.1sec FD15L HYDRC'GEN WITRIC ACID PEROXIDE H;vd.rogen Octane Carben Ethyl alchol r<!ethyl alcllol Aniline Yinyl et..J:\er B:yd.rate .oí:hydrazine OXYGEN 1630/3990 4570/4210 561(>52 ¡O 4190/3690 3810/3600 46lOj+150 't32Q/!;2'ÍS 1100/4200 'J2'6/3990 417ú/4J70 H"i5!"JJ2G 3860/3580 3540/3460 3980/3580 3700/3480 3900/3480 3640/3360 39&0/3640 37JO/3550 3990i3650 3740/3560 3760/3430 3960/3530 - 1 ~}280/397() '- i ""--~---"--'-- "C TBEOBETICAL EXAUST SPEED OF cmmUSTION GASES IN misec u -FUEL u 6S00/b300 4920/-\,,20 6095/5710 5090/4930 1790/1720 4840/4650 4640/4420 4765/4680 4890/4780 46J 0/4330 Octane Carbon Ethy1. alcohol Methyl alcohol Aniline Vinyl ether HYdrate of hydrazine 397;/3"40 47'10/.f62"1 -IÓ'5ii/H,';() ,1'17:JH90 '152;)/142(i 1610/)-\'10 u agents ~ Tht4'above table not yet inactual is calculated use. for the ,,-.. FLOURlDE OZONE Hyd.rQgen ""-u--u,,u,-- same fuels but with "'--"""""'-""""--'" " eombustion The highest theoretieal exhaust speed in our ehemieal formulas will be produced by the reaction oí pure ozone with pure beryllium at 7,,310 m/see. at sea level. By eomparison" our high explosives look rather slow; the theoretieal exhaust speeds here are: Nitroglycerine mjsee 3,880 Nitrocellulose 3,660 Dynamite 3,300 Double-base Picric powder 3,240 acid To burn hydrogen following values: 2,600 with oxygen, with an excess of hydrogen, 0.5 kgH2 1. O kgH2 1.5 kgH2 2.0 kgH2 2.5 kgH2 mjsee 5,170 5,030 4,890 4,770 4,680 4,570 1 kgH2- 8 kg02- 3.0 kgH2- 4,470 -1 kgH2 I . !t - 8 kg02 1 kgH2- 8 kg02 1l}gH2 -8 ~02 1 kga2 - 8 kg02 1 kgH2- 8 kg02 1 kgH2- 8 kg02 - --,2 - we get the 4 ~ These data were compute o. by Dr. Eugen Sanger, published in 1950. a German, anO. were Looking at the preceding table, we can see that the soliO. fue1s are those that have thé lowest ve10oity; as Dr. Sanger said, the values here are theoretioa1 anO. it is rea11y impossib1e to get these results in a rocket because of incomplete reaction in combustion, heat 10ss, and theoretical expansion ratios. However, we can say that there is a tendency to reach the va1ues shown in these tables. We can see that the se1eotion of a fue1 depends to a great extent on the conditions of the working medi'um anO. the achievement of target figures. CHARACTERISTICS Oxygen injection 75% alcohol Pressure pressure: infection in combustion Gas extraction Tempara ture OF V-2 ROCKET ENGINES 26 kg/sq cm (370 lbs/sq pressure: chamber: inch). 24.65 kgfsq cm (350 1bs/sq inch). 15.14 kgfsq inch). vel.ocity at sea 1evel: in combustion chamber: cm (215 1bs/sq 2,000 m/sec (6,560 ft/sec). 2,0000 C. Here are some more data derive o. from the dimensions of the rocket engine: 1 kg of 75% alcohol with LOX with the previous1y mentioned we need 6.8 cu dm/l kg Of 75% alcohol in the com.bustion To burn characteri'stics, chamber. DESIGN CONSIDERATIONS 1. Fuel consumption, 75% alcohol: LOXconsumption: 1 kg/sec. 1.3 kgfsec. 2. \ve a1so assumed that a11 materials would have to have high resistance anO.wou1.d have to take high temperatures without deformation. 3. Thé variolls injectors had to have ):~ha.:racteristics of high pulverization anO. hao. to be so constituted that combustion would takeplace in them éasi1y. 4. In order to be able te start en the construction right away, we specified th~t we were not going te use any material that was not readi1y available in Merico. - .---------- -3 ,--.-----------------------..------- -- ~ Our ca1culations refer to sea' 1eve1: that is to say, we took an sq mChJsicf) because in atmospheric pressure of 1.033 kg/sq cm (14.7 the beginningwe though that we would have to make OUTlaunoh tests at a site near the port of Campeohe, sinoe this was an area which offered gOQd environmenta1 oonditions. L In accordance with the mitia1 had the following value Vj impulse I " where: w = weight of combustion'gases/sec g = 9.$2 m/sec (32.2 v j = gas escape ve1ocity F the specific (1) =: l F oharaoteristics, -- 2.3 9.8 ft/sec2). -2,000 m/seo (such as we know it). x 2000 = 470 kg. Therefore, 1 = - 2.3 kg/sec 470 2.3 the specific impulse has the iollowing va1ue: = 204 see. , ?~!fJs OF .ca>1BU~T;¡:O~, ~B~ There is no doubt that the shape oi the combustion chamber oi the V-2 fS was so designed as to be able to withstand high pressure; that is to say, it had the shape oi a sphere which, as we 1mow, can take twice the inside pressure m eomparison to the cy1mdrica1 shape. 'Sffice tne'llse oi thespheriéa,lsha.pe tion me>re,cc>IÍJ;p1icated,we deÓidedto sett1e b,~q~-g;$e"{~l;fít$;tsÍ!IlP1icity:; he~e1¡ie'wer~a1so a.rrangement for the injedtionnozz1es. . _.Il1!;~AT.~O:N' OE.;,~~.ION>'O}i' w'ould rriake the entire construconthe cylindrica1 shape nere going to get an ad~qua.te ..Gq.IBTJ~TION GHAMBER The reasonable ve1ocity m,the .combustion chamber whicn was indicated here ¡,re.s 60-120 m/ sec (200-400 it/ sec), in case oi the cylilldrica1 shape; that is to day, these low velocities, with 1.iach numbers mucn lower than the -4 - j I ! 1 ¡ SCT.,.l CO~USTION 1 fJ5% CHAMBER alcohol betention wall o --- I 'f&' .1ox retenhon 141 :L - vTall ' t. nJec t~on "~ , ~ead li "" ---l. ~njection capsul e, 22 alcohol nozz ~~nd 22 LOX' ilozzles - <D "" M') 01 .... les O N - ( N I looling \ozzles I I [gnition capsule, \2 sec_i I I \lcohol l' COOr J.ng ;¡¡If) ;¡¡- jacket Iteel wire for ¡pening pressure valve Iiattery r I bpper wire for ~ition of 'capsule, ¡tartingcombustion ~_..JI"'-' I ~" I ""'- L . >""--"""""",,,,-," -~".,~I .,"';"' """ ,¡ '-- '1,-v.t" . 1 1 t,==~=~~'~'="=-~~=:='::::~:=-=~-"~r ! ~,1l dimensions in mm"-r-"'-"""'--'~"'T i ~tj¡ I,~,M I ~::='j'Lf-4 -_.~:..'~'11fJ¡ . 1-- L~J2:. '1$' 10 q ~ exhaust gas velocity, do not produce' strong, dynamic vibrations on the inside, which can become quite dangeI'ous; if we were to have to contend with this kind of vibration, we would need high-resistanee eombustion ehambeI's which in turn wou1.d have to be much heavier. In addition, these strong vibrations wou1.d lead to low temperatures in the wall.s of the combustion chambers, as we shall. see. For the ealeulation of this part, we used the fo1.1.owing formulas; -- which Sinee we must have continuity of combustion flow of gases -- we have the fo1.1.owing: henee: ~ A=- 'Vl (2) vl A = section of combustion V, = inside velocity . ehamber in sq m oI' sq ft. gases = 60 m/see of combustion w = weight of fuel and eombustion lb/sec = 2.3 kg/sec. volume in eu ft/lb From equation agent (200 ft/sec). (pI'opellants) in kg/sec or or cu m/kg. (1) w=Besides, a = w V, AVI Vl = specific produces Fg Vl we know that _ V _RT l the total volume is = - 1- - Pl But, as we know, the constant weight m, is constant; Substituting i.e., in V, Vl=- of the gases R, due to its molecular Rm.= R' = 847 in the metrie system. we have: RtTl 11 I . (3) m Pl - q -------- ~ ' '-' ' ' ~ Therefore, the circular have the fo110wing va1ue: A = FgR section of the combustion chamber wil1 t~T1 (4) V.mPlv1 J Here we have: F = 470 kg g = 9.$2 m/sec'¡.. R t = Rm = 847 = 1544 (US system) - F. TI = 20000 e + 273 = 22730 e = 4091.40 vj = 2000 m/ sec ¡..nÜchis the velocity of the exhaust gases at sea level and which was then checked out exper:i.mentally; further down, we wil1 ca1culate the coefficients with vlhich the nozzles are working in order to justify the va1ue m ..= 23.4 in lb/lb mol = 23.4 Pl of V j' when kg/kg = pressure in combustion chamber 215 lb/sq inch. Substituting we compute the nozzle opening. moJ.. 15.1 kg/sq cm = 15.1 X 10~ kg/sq m = ~ the va1ues, the section wi11 look like this: A = - 470 X 9.$22 X $47 X 2273 2000 X 23.4 X 15.1 X 10* X = 20$.$5 X lOA m'-= 20$.$5 60 cm""" ~ hence: d =. lb.32.cm. For the design, we took: d = 16.5 cm. In dimensions order to for the get the engine, figures we .need in order we set up the following: to find the right V.J = v2 P3 = P2 -7 --' '--'-'~ ""'- ..-~~ ~ ".--~~ ~ , .. r -'"f$ -1 ¡ I ..L , r ~ ~ h; f ~ v"" .-. 1.- v"' ¿., I "2 J '2 .. DETERt1D!ATION OF AREA OF THROAT The effective thrust is smaller than the one computed for an ideal rocket and "le can correct it by means of an empirical f9-ctor 6. F =""effective thrust Fi = ideal thrust. \fe now have: F = .~ Fi = 6GFPi Ag (5) [ Ag =.6,CF:1 I He~e P1 = p~ssu~ in combustion chamber I Ag = area ofthroat CF can = thrust coefficient. In order to avoid having to go through go to any text book on thermodynamics and value of CF: this find whole operation, l.¡e the fol1owing for the - - ~8 ---~--"--" ~-~--"--"' ~" ' '---'--'_~-~"--_'_'_'"--'--~' ' """ "-"--"---"'---"'."""""-' "'""- ~-\ ~~~~ ef = In this = pressure P, p 2. p I.~~) T t~ ~+I~ p + -p ~ p\ A (6) ---2::- A~ equation: in combustion chamber of nozzle. = pressure at nozzle out let. = atmospheric pressure. ,3 It is a good idea to set up P2 = P3 which means that the formula is reduced sole1y to the function of the square root; ttlat is to say: A - P2 ~ P3 X = O A g Pl e For dealing our here P1 ca1culations, with = 15.1 75% we alcohol take: with K = 1.22 = -E-- q .= The / 2. kg m . here 11 (6), we find that: 1.:35 correction factor the conditions of the nozzle. with are LOX. Sub§t:;itutihgth~ values in the formula F we X 10'+ kg/m'1- P2 = 1.033 X 10 ~C because e v is figured 6 varies from 0.92 to 1 and depends on Since in our case the force we are "\",rorking on an exper:irn.ental1y determined ve1ocity, 1Ira have: =1 - 9:;- ~ ,-""",-".,_.,---, ",--", ~,~~ ~~~, ~._-,..~-~ ~-~" '-' "--" '-.- - --.-...--- Therefore: =- F A g e FPl Substituting - A g - g we get: 470 kg. = 23.056 X 10,.,1/-m~ 1.35 X 15.1 X 101/- Therefore, d the values, the diam~~~~ of the throat = 5.4 g J cm2. be: cm. Checking all va1ues and taking the exact surface for the throat is: A will = 23.056 -- 22.$5 5.4 cm as our definite diameter, cm2.. We have thus determined the section of the throat L~ accordance with the thermodynamic characteristics and >'lenO1'lgo on to determine the other dimensions of the combustion chamber; here we start on the assumption that we are going to usethe characteristics ofthe combustion chamber of the engines for the V-2; as ¡'le said in the beginning, we need an average vo1ume of 6. $ cudm in order to burn 1 kg of 75% alcohol and the total ang1e in the cone of the combustion chamber must be 860. Through geometric calculations, we determined the form of the combustion chamber in accordance with the dimensions determined ear1ier; that is to say, a diameter of 16.5 cm and a cylindrical portion with a 1ength of 27.2 cm; the conical part, up to the throat, is 6.$ cm. We mustnote here that, if we are going to take into account the reConnnen.dations made in various books on the 1ength of the combustion qha.¡nbe~,we wouJ,d geta chamber that would be 50% smaller volume-wise thán the onedetermined.' with the data for the V-2; therefore we decided to ignore these reconnnendations and started designing the entire engine to fit the combustion chamber and nozzles, on the basis of the considerations arising out of the V-2 data. -10 - ESCAPE GAS NOZZLE The design of the V-2 rocket used an area ratio of 3.3 between the throat and the outlet and had an angle of 24 at the mouth. In accordance with the above data, we have: A2 = 3.3 X At = 3.3 X 28.85 = 95.205. d2 = 9.8 cm. We did not.provide for a cooling chamber in the first engine because we wanted to make an objective determination of the places where high temperatures 1¡J'e get due to the position of the injector capsules. In the second test'w6 'mad~ without coo~ing jacket on the engine at San Bartolome, we ¡"ere 9-ble to find the place s ¡"here we were going to have to put coo1ing nozz1es, regardless of ¡..hetherwe were a1so going to have cooling jackets there. 1 I As we can see in the drawing,we putin four injector capsules, each used for the injection of alcohol in a horizonal jet, as well as for the injection of LOX at an angle of 300 toward the center of the combustion chamber. The final design was very satisfactory becauS'e it was possible to get high-temperature flames in the central por.tion. The calculations for the alcohol.injectors as well as for the injectors -- keeping in mind the drop in the oxygen pressure from 26 / IDX to 15.1 kg/sq cm and in alcohol from 24.65 to 15.1 kg/sq cm, using the findings of various authors and figuring the costs -- later on led to completely unacceptable results because in 5afo of the cases the expenditures for 1 kg of 75% alcohol and 1.3 kg LOX came out smaller. ~ve are not going into the computations here because we do not really need to do so since the number of holes and diameters were determined expertmentally. The injection system adjustment was a ve~J delicate job because, in addition to handling the flow of 1iquid we also had to have perfect injection; we achieved this in the alcohol injection nozzles by means of radial guides and in the oX'".fgen nozzles by meaDS of some screws in the copduction pipes. In the chamber we had assemblies of miniature parts; thj,sgave uS aperfect mixture ineach capsule so that, after testing each motor, wé observéd practica1ly perfect combustion in each capsule. Number of injectors: for alcohol: 22 in each capsule; injector diameter 0.8 IDm; for oxygen: 22 in each capsule; injector di&üeter: 15 injectors with a diameter of 0.9 mm and seven injectors with a diameter of 0.85 ffiu. ... This means that we had a total of 88 alcohol and 88 oxygen injectors. - :11..-............ 1, ¡ In the cylindrical portion oI' the combustion chamber, at a distance of 15 mm, we had 24 cooling nozzles ,dth a diameter of 0.4 mm; our injection pressure was 24.65 kg/sq cm (350 lbs/sq inch); during the second test of the motor we had designed, ¡.¡e observed that the material in this part became red-hot. In the SCT-1 engine, the onl;}r cooling occurred when the rocket was 1aunched into space; in the other engines we noticed failures in some of the cooling nozz1es; this is why welater on provided the next rocket, that is the SCT-2, with a complete alcohol coo1ing ~system; that is to say, the alcohol had to pass through the cooling jackets before arriving at the injection nozzles. We decided to use inoxidable chrome-nickel steel because it was more resistant at higher temperatures and bécause we had this material available in the required size ando s.h~pe.Jn Merico. The strongest ness ~tress to which the design was subjected (a) in the cylinder of 1.6 illlll (1/16"): of the dXPl f=- combustion = 16.5 X 15.1 2 X 0.16 2 e thickness (b) in the outside of 1.6 mm: d ch X P cy1inder a - f = with 1$.4 X 24.65 a sheet thick- :: 778 kg/cm2.. of the coo1ing jacket, 2 e These stresses chamber, was as fol1ows: with a sheet = 1417 kg/ cm').. 2 X 0.16 at normal temperature kg/sq cm, where permanent deforrnation (304) N. i 18-20/C r $-10. begins ,.¡e re much lower than 3,500 in the chrome-nicke1 steel To make the combustion chamber stronger and more resistant to the high temperature!:! developed in it, .\Teprovided the SCT-l with reinforcing as wel1as the conical ri11gs (hoops) to strengthen the cy1inder interior part toward the top; but, as we said before, when it was decided to feed all the alcohol through the cooling jacket, engineer Walter C. Buchanan recommended that we insert four conduction screws into the cooling jacket so that each nozzle would have one out1et for each injection capsule. - J2- ~~ ~~ ,-- ~-~ ~~ ~ , '- ---- .... Combustion 4 conducting chamber of SCT-2 with screws for cooling. ~ Combustion chamber with upper cylinder of cooling jacket of SCT-2. ~ This system solved our problem here; we can thus say that the design oí the SCT-2 embodied all the lessons we had learned. The ignition capsule shovm in the illustration oí the engine \,ras designed íor a duration oí 12 seconds with an adequate mLxture of powder and phosphorus; íor our next rocket claculations we shall reduce the time since we no longer need this longer interval here. This capsule i5 ignited electrically and by remote control. As 1tfecan see in the illustrations, the principal alcohol conduit has a retaining wall, as does the outlet duct írom the LOX tank; when the pressure battery valve is opened; both of these conduits are closed off and injection into the combustion chamber begins. ... Further details of the design can be seen in the engine illustra- tions. We have thus shown how we designed the rocket engine. Ho\.¡ever,it was necessary to adapt its characteristic so tha.t we could work with a 1.4 ratio between LOX and alcohol; that is to say, the chamber worked with a combustion gas weight of 2/4 kg/sec. The dimensions of the tanks and the shape of the rocket were determined on the basis of the V-2 data; this is ,-.rhy thecombustion time was 44 seconds. -:¡4- - ----.---- ~