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LA Jc/fl.: :/,::><

SCT-1 ROCKET ENGINE

by B~ Buitron

N71-71501

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R)

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(CODE)

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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, ,rebased our study on

what fe kne,'/ nd en w'.o.at

're viere t:rying te achieve in the way of opt,i¡rn.rm

results considering the high ex.haust speed of the combustion gases,

 ,¡hieh

inturn of course ,vould give us IP..axi,'rlumltitudes.

Ihe

follo. ,:L'1g

table

is an analysis of t;ypiea1 fue1 eharaoteristics ,,¡hieh have been usad

L-rJ.experiments with recket engi..1'1es.

TBEORETICAL EX.4UST SPEEJ) OF COMBUSTIOt'I G SES ll m sec

FD15L

HYDRC'GEN

PEROXIDE

WITRIC ACID

OXYGEN

H;vd.rogen

Octane

Carben

Ethyl alchol

r< ethyl alcllol

Aniline

Yinyl

et..J:\er

B:yd.rate .oí:hydrazine

1630/3990

4190/3690

3860/3580

3980/3580

3900/3480

39&0/3640

 99 i 65

3960/3530

4570/4210

3810/3600

3540/3460

3700/3480

3640/3360

37JO/3550

3740/3560

3760/3430

561(>52 ¡O

46lOj+150

't32Q/ ;2'ÍS

1100/4200

'J2'6/3990

417ú/4J70

H i5 JJ2G

~}280/397()

- 1

'-

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TBEOBETICALEXAUSTSPEED OF cmmUSTION GASES IN m sec

 

C

  --~--- --'--

u u -u--u,,u,-- ,,-..

FLOURlDE

FUEL

Hyd rQgen

Octane

Carbon

Ethy1. alcohol

Methyl

alcohol

Aniline

Vinyl ether

HYdrate of hydrazine

OZONE

6095/5710

5090/4930

1790/1720

4840/4650

4640/4420

4765/4680

4890/4780

46J 0/4330

6S00/b300

4920/-\,,20

397;/3 40

47'10/.f62 1

-IÓ'5ii/H,';()

,1'17:JH90

u -- - --'

'152;)/142(i

1610/)-\'10

Tht4'above table is calculated for the same fuels but with eombustion

agents not yet inactual use.

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

Nitrocellulose

Dynamite

Double-base powder

Picric acid

mjsee

3,880

3,660

3,300

3,240

2,600

 t

To burn hydrogen with oxygen, with an excess

o

hydrogen, we get the

following values:

-1 kgH2 - 8 kg02

I .

1 kgH2- 8 kg02 - 0.5 kgH2

1l}gH2 -8 ~02 -   OkgH2

1 kga2 - 8 kg02 - 1.5 kgH2

1 kgH2- 8 kg02 - 2.0 kgH2

1 kgH2- 8 kg02 - 2.5 kgH2

1 kgH2- 8 kg02- 3.0 kgH2

-

--,2 -

mjsee

5,170

5,030

4,890

4,770

4,680

4,570

4,470

4

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~

These data were compute o. by Dr. Eugen Sanger, a German, anO.were

published in 1950.

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 OF V-2 ROCKET ENGINES

Oxygen injection pressure:

26 kg/sq cm (370 lbs/sq inch).

75 alcohol infection pressure:

24.65 kgfsq cm (350 1bs/sq inch).

15.14 kgfsq cm (215 1bs/sq inch).

ressure in combustion chamber:

Gas extraction vel.ocity at sea 1evel:

2,000 m/sec (6,560 ft/sec).

Temparature in combustion chamber:

2 C.

Here are some more data derive o. from the dimensions of the rocket

engine:

To burn 1 kg of 75 alcohol with

LOX

with the previous1y mentioned

characteri stics, we need 6.8 cu dm/l kg Of 75 alcohol in the com.bustion

chamber.

DESIGN CONSIDERATIONS

1. Fuel consumption, 75 alcohol:

LOXconsumption: 1.3 kgfsec.

2. \ve a1so assumed that a11 materials would have to have high resistance

anO.wou1.dhave to take high temperatures without deformation.

1 kg/sec.

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 -

,--.-----------------------..-------

--

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Our ca1culations refer to sea' 1eve1: that is to say, we took an

atmospheric pressure of 1.033 kg/sq cm (14.7

 

sq mChJsicf) because in

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.

In accordance with the mitia1 oharaoteristics, the specific impulse

had the following value

l F

=: Vj

I

(1)

where:

 

w =

weight of combustion'gases/sec   2.3 kg/sec

g = 9.$2 m/sec (32.2 ft/sec2).

v j = gas escape ve1ocity -2,000 m/seo (such as we know it).

F - 2.3

x 2000 = 470 kg.

9.8

Therefore, the specific impulse has the iollowing va1ue:

1 = 470 = 204 see.

2.3

, ?~ fJsOF .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 w'ould rriake the entire construc-

tion me>re,cc>IÍJ;p1icated,we deÓidedto sett1e onthe cylindrica1 shape nere

b,~q~-g;$e {~l;fít$;tsÍ IlP1icity:; he~e1¡ie'wer~a1so going to get an ad~qua.te

a.rrangement for the injedtionnozz1es.

._.Il1 ;~AT.~O:N' OE.;,~~.ION>'O}i'

. .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

 

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I

 

1

¡

1

fJ alcohol

betention wall

I

---

.1ox reten-

hon vTall

t.

t

 

n

 ~

,

nJec ~o

~ead ---l.

~njection capsul

l

e,

22 alcohol nozz es

~~nd22 LOX

ilozzles-

I

looling

\ozzles -

I

I

[gnition capsule,

\2 sec_-

i

I

I

l

 

\lcohol COOr J.ng

jacket

Iteel wire for

¡pening pressure

iattery valve

r

I

bpper wire for

~ition of capsule,

¡tartingcombustion

~_..JI - -~ .,~I i ~tj¡ I,~,M I

I ~

1

~::='j'Lf-4

  '-

.

., '; '

¡

'~'11fJ¡

'1,-v.t . -_.~:.. .

1

L

> -- ,,,,-, , '-- 1--

t,==~=~~'~'= =-~~=:='::::~:=-=~- ~r L~J2:.

  ~,1l dimensions in mm -r- '- '--'~ 'T

SCT l

CO~USTION

CHAMBER

j

(

;¡¡-

If)

;¡¡-

'1$'

10

q

o

'f&'

141

:L

li

 

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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 com-

bustion chambers, as we shall. see.

For the ealeulation of this part, we used the fo1.1.owing formulas;

Sinee we must have continuity of combustion -- which produces a

flow of gases -- we have the fo1.1.owing:

AVI

= w V,

henee:

~

 Vl

.

A=-

vl

  2)

A = section of combustion ehamber in sq m oI sq ft.

V, = inside velocity of combustion gases

= 60

m/see 200 ft/sec).

w = weight of fuel and eombustion agent pI opellants) in kg/sec or

lb/sec = 2.3 kg/sec.

Vl = specific volume in eu ft/lb or cu m/kg.

From equation 1)

Fg

w=-

Vl

Besides, we know that the total volume is =

_

V

_RT

l

- 1 - -

Pl

1

I

ut, as we know, the constant of the gases R, due to its molecular

weight m, is constant; i.e., Rm.= R = 847 in the metrie system.

Substituting in V,

we have:

.

RtTl

Vl=-

m Pl

  3)

- q

~

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~

Therefore, the circular section of the combustion chamber wil1

have the fo110wing va1ue:

FgR t~T1

A =

V.mPlv1

(4)

Here we have:

F = 470 kg

g = 9. 2 m/sec'¡..

R t = Rm = 847 = 1544 (US system)

-

TI =

20000 e + 273 = 22730 e = 4091.40

F.

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 of V j' when we compute the nozzle opening.

m ..= 23.4 in lb/lb mol = 23.4 kg/kg moJ..

Pl =

pressure in combustion chamber

~ 15.1 kg/sq cm = 15.1 X 10~ kg/sq m =

215 lb/sq inch.

Substituting the va1ues, the section wi11 look like this:

470 X 9. 22 X 47 X 2273

A =

-

2000 X 23.4 X 15.1 X 10* X 60

= 20 . 5 X

l

m'-=

20 . 5

cm ~

hence:

d =. lb.32.cm.

For the design,

we

took:

d = 16.5 cm.

In order to get the figures we .need in order to find the right

dimensions for the engine, we set up the following:

P3 = P2

V. = v2

-7

'--'-'~ '-

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

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r

-1

..L

r

~

h;

I

, I

~

-' f$

f

~ v .-. v '

1.- ¿.,

¡

 2 J '2

.

DETERt1D ATION OF AREA OF THROAT

The effective thrust is smaller than the one computed for an ideal

rocket and lecan correct it by means of an empirical f9-ctor 6.

F = effective thrust

Fi = ideal thrust.

 f

now have:

F = .~ Fi = 6GFPi Ag

[ Ag =.6,CF:1

 5)

He~e P1 = p~ssu~ in combustion chamber

Ag = area ofthroat

CF

= thrust coefficient.

In order to avoid having to go through this whole operation, l.¡e

can go to any text book on thermodynamics and find the fol1owing for the

value of CF: -

-

~8

-

~-~-- -- ' ~ ' '---'--'_~-~ --_'_'_' --'--~' ' - -- --- '--- '. -' ' -

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e =

f

~~~~

t~ ~+I~

~-\

I.~~) T

+

p

-

p

~

p\

A

---2::-

A~

(6)

In this equation:

P,

= pressure in combustion chamber of nozzle.

p = pressure at nozzle outlet.

2.

p = 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; ttlatis to say:

A

P2

- P3 X ~

=

O

A

Pl g

For our ca1culations, we take:

e

K = 1.22 =

-E-- because we are

e

 

dealing here with 75 alcohol with LOX.

P = 15.1 X 10'+ kg/m 1-

1

q

/

2.

P2

= 1.033 X 10 kg m .

Sub§t:;itutihgth~ values in the formula (6), we find that:

~C

.=

1.:35

F

The correction factor

6 varies from 0.92 to 1 and depends on

the conditions of the nozzle. Since in our case the force we are \ ,rorking

with here is figured on an exper:irn.ental1y determined ve1ocity,

1Ira have:

11

=1

- 9:;-

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

--.-...---

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Therefore

F

A =-

g e FPl

Substituting the values, we get:

A - 470 kg.

= 23.056 X

10 . 1/-

m~ = 23.056 cm2.

g - 1.35 X 15.1 X 101/-

Therefore, the diam~~~~ of the throat will be:

d = 5.4 cm.

g

Checking all va1ues and taking 5.4 cm as our definite diameter,

the exact surface for the throat is:

A - 22. 5 cm2..

g -

We have thus determined the section of the throat

 ~

accordance

with the thermodynamic characteristics and

> le

nO1'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.

Wemustnote 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 -

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ESCAPE GASNOZZLE

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

  ¡J e get

high temperatures 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-bleto 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.

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 IDX

injectors   keeping in mind the drop in the oxygen pressure from 26 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   fo of the cases the expendi-

tures 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' .fgenozzles

 

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-

 

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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 ~tress to which the design was subjected was as fol1ows:

(a) in the cylinder of the combustion chamber, with a sheet thick-

ness of 1.6 illlll (1/16 ):

dXPl

f=-

2 e

=

16.5 X 15.1

2 X 0.16

:: 778 kg/cm2..

(b) in the outside cy1inder of the coo1ing jacket, with a sheet

thickness of 1.6 mm:

d X P

ch a

2 e

- 1$.4 X 24.65

2 X 0.16

= 1417 kg/ cm').

=

These stresses at normal temperature   ¡ee much lower than 3,500

kg/sq cm, where permanent deforrnation begins in the chrome-nicke1 steel

(304) N. i 18-20/C r $-10.

To make the combustion chamber stronger and more resistant to the

high temperature : developed in it,  \Teprovided the SCT-l with reinforcing

ri11gs (hoops) to strengthen the cy1inder interior

as

wel1as the conical

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-

~ ~~ ,--

~-~ ~~ ~ , '-

----

....

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Combustion chamber of SCT with

4 conducting screws for cooling

Combustion chamber with upper

cylinder of cooling jacket of

SCT 2

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~

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  tfcan 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.twe 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

.

~