I '9~¿ eJe
/1 .~-~[,/'..!/j'
/ .. ----
A Service of:
National Aeronautics and
SpaceAdministration
$CII!NTlFIC &.
T!!CHNJQ\L INFoitMAnoN
'"
,
"
Svcs:
r-
---
IP
ENVELOPE
9368
TRCK:
5581
0246
ORIGIN ID: BWIAt.jfIJ1} b21-03!:1
.
are: 18APR08
EDNA PAULSON
NASA CENTER FORAEROSPACE INFO
7115 STANDARDDRIVE
ActWgt:
System#:
Account:
HANOVER, MD 210762304
UNITED STATES, US
1.0 LB MAN
272451/CAFE2332
S 122516687
EIN/VAT:
(555) 560-7387
TOLUIS GERARDO SAUCEDa ZARATE
INST MEXICANO DEL ESPACIO EXTERIOR
COLINA DE LAS ORTIGAS
126
NAUCALPAN
ESTADO
DE
53140
,
~
/
lp
FedEx
Express
(MX)
~
~a
I
RWB
X7 NLUA
9368
TRK#
5581
Form
0430
0246
1/1
IP ENVELOPE
REF: 150377/D-04859
DESC1: TECHNICAL REPORTS
DESC2:
DESC3:
DESC4:
EEI:
NO
EEI
30.37(a)
SIGN: Edna Paulson
TIC: S 122516687
D/T: S 122516687
~OUNTRY
MFG: US
~RRIAGE
VALUE: 0.00 USD
GUSTOM VALUE:
1.00 USD
~
;!'hese
commoc:UUes,
technology.
or
soTtware
were
eMported
.rom
the
tUed
States
1n accordance
wHh the eMport adm1nlstraUon
regulaUons.
D1verslon
contrary
to US law 1s proh1bHed.
~e Montreal
or Warsaw ConvenUon
may apply and wll I govern and 1n most
~ses
I1mH the IlabllHy
o. Federal
EMpress .or loss or delay o. or
;¡¡'amage to your sh1pmant.
SubJact
to the cond1 Uons o. the
~ntract
~IGIN
n.
on the
ravarse.
RWB COPY
-
PLERSE
PLRCE
- -
IN POUCH
---
::b ifJc!a;,mel'
(jenel'af
Oneor more ofthe FoDowlngStatements may affect tbIs Docmnent
~
.
.
.
This document has been reproduced from the best copy furnished by the
sponsoting agency. It is being.released in the interest ofmaking available
as much information as possible.
This document may contain data, which exceeds the sheet parameters. It
was furnished in this condition by the sponsoring agency and is the best
copy available.
.
.
This document may contain tone-on-tone or color graphs, charts andlor
pic~res, which have béen reproduced in black and white.
. This document is paginatedas submitted by the original source.
.
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-
- ----.----
~