Development of Russian rocket engine technology

NASA TECHNICAL TRANSLATION

NASA TT F-15,408

DEVELOPMENT OF RUSSIAN ROCKET
ENGINE TECHNOLOGY

Ye. K. Moshkin

Associates) 2° 8 F

H74-22**11

Onclas
G3/2B 3803U

Translation of Razvitiye Otechestvennogo

Raketnogo Dvigatelestroyeniya , Moscow?
Mashinostroyeniye Press, 1973, 256 page

pages.

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION
WASHINGTON, D. C. 20546 MARCH 1974

TABLE OF CONTENTS

From the Author 1

Chapter 1. The Period of Theoretical Foundation of the

Capabilities and Areas of Application of LRE . 3

1.1. At the Wellsprings of Soviet Rocket Design "*

1.2. The Works of N. Ye. Zhukovskiy and I. V.

Meshcherskiy 6

1.3. K. E. Tsiolkovskiy, the Founder of Astronautics ... 10
The Works of K. E. Tsiolkovskiy on the Creation of the

Theory of Reaction Engines 18

The Formula of K. E. Tsiolkovskiy 22

Suggestions for LRE Fuels 28

Recommendations for the Design of Combustion Chambers . 31

Uevelopment of Feed Systems 36

1.4. One of the Pioneers of Rocket Technology,

Yu. V. Kondratyuk 40

The Works of Yu. V. Kondratyuk on Rocket Engines .... 44

Suggestions for LRE Fuels 44

Recommendations for the Design of the Combustion

Chamber 45

Development of Feed Systems 46

1.5. The Scientist and Inventor F. A. Tsander 47

The Works of F. A. Tsander on Rocket Engines 53

Investigation of Fuels 55

Study of Processes Within the Chamber and Cooling

Conditions 57

Increasing Specific Impulse and Efficiency 58

The OR-1 Reaction Engine 59

The OR- 2 Rocket Enf.ine 61

Plans of Rocket Engines 65

Chapter 2. The Firs-, Rocket Scientific Research and

Experimental Design Organizations in the USSR . 67

2.1. The Initial Period of Development of GDI. -- the

N. I. Tikhomirov Laboratory 68

2.2. The Gas Dynamics Laboratory 71

2.3. Liquid and Electrical Rocket Engines and Rockets

of GDL 76

Experimental Electric Rocket Engine 77

Selection of Fuel for LRE 80

Engines with Annula. Combustion Chambers 82

Engines with Radially Placed Nozzles 85

Engines with Internal Protective Coatings 88

Engines with External Cooling 91

GDL Engines for Flight Vehicles 84

Fuel Feed Systems and Stands 99

The Rocket of GDL . 101

2.4. The Moscow Group for the Study of Reaction Motion,

CS Osoaviakhim USSR (MosGIRD) 106

PRBCEDING PAGE BLANK NOT FILMED
iii

2.5. Liquid-Fueled Rocket Engines and Rockets of GIRD . . 115

The 02 Engine 115

The 10 Engine 118

The 09 Engine 121

The 03 Engine 123

The GIRD-09 Rocket 124

The GIRD-X Rocket 125

The GIRD-07 Rocket 127

Air Breathing Reaction Engines . 128

2.6. The Leningrad Group for the Study of Reaction

Motion (LenGIRD) 131

The Powder Rockets of LenGIRD 134

Liquid-Fueled Engines 135

2.7. The Work of the Society 137

Chapter 3. The Reaction Scientific Research Insti-
tute (RNII) 146

3.1. Creation of the Institute 146

3.2. The Activity of the Institute 148

Powder Rocket Weapons 148

Liquid-Fueled Rocket Engines 150

Air- Breathing Reaction Engines 153

Flight Vehicles 153

3.3. Nitric Acid LRE 154

The ORM-53 - ORM-63 Engines 154

The ORM-64 - ORM-70 Engines 155

The ORM-101 - ORM-102 Engines 162

The GG-1 and GG-2 Gas Generators 162

The RDA- 1-150 Engine 164

The RDA- 300 Engine 167

3.4. Oxygen LRE 169

The 12K Engines a69

The 205 Engines 170

The RDK-1-150 Engine 172

The Engine of P. I. Shatilov 173

3.5. Developments by Design Bureau No. 7 (KB-7) 174

Chapter 4. Liquid-Fueled Rovket Engines for Aviation . . 177

4.1. The Liquid-Fueled Rocket Engines of OKB-NKAP .... 177

The RD-1 Engine 179

The RD-lKhZ Engine 183

The RD-2 Engine 187

The RD-3 Engine 188

The RD-4 Engine 190

4.2. The Liquid-Fueled Engines of RNII and the NKAP

Design Bureau 190

Conclusions 195

IV

FROM THE AUTHOR

The achievements in the mastery of space have attracted the [V
attention of mankind to various problems of astronautics,
including problems of the history of its development. In the
USSR in recent years we have seen increasing interest in the
study and analysis of the documents from the archives describing
the development of domestic rocket technology. It is gratifying
to note that the teams of scientific workers in this area are
growing -- the history of astronautics is now being studied not
only be veterans of rocket technology, but also by young
specialists as well.

Various periods are studied -- before the October Revolu-
tion, before the beginning of the Great Patriotic War, before
the launch of the world's first artificial satellite (4 October
1957) , before and after the first flight of man in space (12
April 1961) . Additional periods have been determined by new
achievements in the mastery and study of space --by successful
flights to Venus and Mars, soft landings on the moon and Venus,
launching of automatic interplanetary stations, unique experi-
ments vith canned spacecraft, etc. Widely varied aspects of the
history of astronautics are being studied deeply: the develop-
ment of the design of rockets and engines of various types, the
work of experimental-design, scientific research, administrative,
party and social organizations, the activity of individual
persons. Therefore, it is impossible at present to speak of the
history of rocket and space technology as a single theme, to
attempt to write an all-encompassing book on the development of
astronautics, or to pretend completeness of presentation.

This book is dedicated to the history of the creation and
development of Soviet liquid-fueled rocket engines [LRE] (as we
know, LRE are the most important engines in modern astronautics).

The author has attempted to describe the contribution made /4
by our countrymen K. E. Tsiolkovskiy , Yu. V. Kondratyuk, F. A.
Tsander, V. P. Glushko, S. P. Korolev, M. K. Tikhonravov and
others to the science of rockets and rocket engines burning
liquid fuels, as well as the successes achieved during the Great
Patriotic War in the preparation of the fundamental basis for
the further development of rocket engine construction, and most
importantly to show the basic role of the gas dynamic laboratory

*
Numbers in the margin indicate pagination in the foreign text.

(GDL), Group for the Study of Reaction Engines (GIRD) nd the
world's first Reaction Scientific Research Institute (KNII).

The author found it impossible to analyze all of the
engines designed and produced in the USSR. However, in order
that the reader might gain a more complete concept of the inter-
relationship of the widely varied and highly complex problems
solved in the creation of engines, some LRE designs are des-
cribed rather completely.

The book was written utilizing materials from the archives
of the Academy of Sciences USSR, the GDL Experimental and
Design Bureau and many other organizations. Many comrades
kindly provided the results of their own historical studies
and made useful recommendations during the preparation of the
manuscript. The author is truly grateful to all those com-
rades who took part in the creation of this book, and particu-
larly to N. V. Ivanov, V. M. Komarov and D. A. Shushko.

First must come thought, imagina-
tion and dreams. They are
followed by scientific calculation.
Then, finally, the thought is
brought to life.

K. E. Tsiolkovskiy

Chapter 1. The Period of Theoretical Foundation of the /5

Capabilities and Areas of Application of LRE

1.1. At the Wellsprings of Soviet Rocket Design

The development of rocket technology before the 17th
century has been very little studied. The first reliable
information on the use of rockets in Russia relates to the last
half of the 17th century. In 1680, an "institution" was
created in Moscow, where firework rockets were manufactured.
The production of powder rockets in Russia expanded continually
after that time, but these rockets were quite primitive, even
during the latter half of the 18th century.

After the use of military rockets by the English army in
the seige of Bulon and Copenhagen in 1805-1807, a military
scientific committee began to study military ro< cets in Russia.
After a number of unsuccessful experiments, a member of the
military scientific committee named Kartmazov made two types of
military rockets in 1814 -- incendiary and explosive. In 1815,
the famous artillery scientist A. D. Zasyadko (1779-1837) began
to perform experiments with military rockets. In 1832, all the
"rocket institutions" in Russia were combined into the Peterburg
Rocket Institution, which served as a center for the creation
and manufacture of domestic military rockets. Until the mid-
1840 's, rocket building in Russia developed slowly, producing
low quality rockets due to the primitive state of the technology
of their production. Then, due to the wide use of rockets
during military actions in the Caucasus, the attitude toward
problems of improvement and production of rockets in Russia /6
changed sharply. At this time, the greatest Russian artillery
scientist, Konstantin Ivanovich Konstantinov (1818-1871) began
to work on the development of rockets. By 1845, 1000 two-inch
military rockets were delivered to the Caucasus. The quality of
the military rockets produced by the Peterburg Rocket Institu-
tion was significantly improved. By the mid-1850's, military
rockets were widely used and proved their utility. As a result
of this, military rockets were made a part of the armament of
the Russian army and navy.

In the 1850's and 60's, K. I. Konstantinov published
several works on problems of the production and use of rockets.
K. I. Konstantinov first noted that the eccentricity of the
reaction force was one of the main reasons for scattering of
rocket impacts. Discussing the principle of motion of rockets,
he noted that as the powder burned, the impulse imparted to the
rocket was equal to the impulse of the exhaust gasses. Thus,
K. I. Konstantinov first formulated the basic law of motion of
rockets, although the mathematical interpretation and production
of a formula for determination of the flight velocity of
rockets were not developed in the works of K. I. Konstantinov.

The possibility of the application of rocket motors for
human flight attracted the attention of many of our engineers,
inventors, designers and scientists. For example, in 1849 I. I.
Tretesskiy (1821-1895) developed plans For rocket powered flight
vehicles j to be powered by steam. In 1866, N. M. Sokovnin
(1811-1894), in his work Vozdushnyy Korabl' [The Airship],
described an aerostat design to be driven by reaction force. In
1867, N. A. Teleshev was awarded a patent for a jet airplane.
In 1880, the talented scientist and inventor S. S. Nezhdanovskiy,
based on theoretical studies, calculations and computations,
concluded the possibility of construction of a reactively
powered flight vehicle. Between 1882 and 1884, he studied the
problems of the energetics of reaction motors, analyzing the
possibility of using liquid two-component fuels for rockets*.
In 1887, F. R. Geshvend in his brochure "General Basis of the jj_
Design of a Steamship for Air Travel," suggested a plan for a
vehicle with a steam reaction engine. In 1896, A. P. Fedorov
in his brochure "A New Principle of Air Flight" abandoned the
atmosphere as a supporting medium and presented a description
of a reaction motor in which gas was to flow from a central tube
(cylinder) .

This hardly exhausts the list of Russian researchers and
inventions dedicated to problems of reaction flight. Among
these efforts we should particularly note the work of N. I.
Kibal'chich (1853-1881).

Nikolay Ivanovich Kibal'chich, the author of the world's
first plan for a rocket flight vehicle 2 , was born 19 October
1853 3 in the city of Korop. In his sixth year in school, he

The works of S. S. Nezhdanovskiy were published only in 1964.
The notebooks and drawings of S. S. Nezhdanovskiy are stored in
the N. Ye. Zhukovskiy Museum of Scientific Memorabilia in Moscow.
2

Kibal'chich, N. I., A Plan for an Aircraft , Byloye, 1918,

No. 10-11.
3
Here and throughout the book, all dates from the prerevolu-

tionary period are given in the old style.

mrt to ipated act ivi • he creat n • - •••.,:. i-

raininu ■.:■-;: \ put Hi ■ ions -f )>. 1. Pis rev, N. G, Chcrn; • ev-
skiy , ■ ; • • • - ■•• S. P. >), ir and edi red f h :.'•,. n i-

script journal , wr i c iag fci i it rt i* >. .. tepan lazl ,
. . >,c 1 ' i tn '•;-.,.••- md tl '■ in h revi its it. In 1871, N. I.
•.,';'./•,,./•;.,'- the .' . ' rhu r i' raj • • . ition Ins • ut e , and

,- ; . ' :• the < ■'■' 'a I -'.;..- 1 1 Ai adt -y .

In .. 878, N. I . Kibal - hi< h rtt over
to an i 1 legal po: t ton - - : ■ u lea
an u i TgJ >und , ilo: • ■ \ ibo a .-ry ,
se •■ ;.' th* .'■ i h mmi t - >r the

Pe ■ ■ s "; i '< ; it Ik - 3 < t iir , he
st u i i he K • . I ies oi : ' • e of
po d< o ! ighi hie if md ci r ici zed
scient i 51 ttc-mpti: to o the j *ob-

le • • hi! ■'. r ig! • ■• , ■ -'■''., :hc ' i Ight
of hi rds,

N. I . Kii :hi« >.• arrest ••• in

co, icct ion with he urde i ■ = v
A I -j is Jer 1 -•> ch U81 i 2 5

Ma eh 1 88 J ittsi before entc -• : prison,

lie pi - .'.••'••". ' n for a • I igh

veil ici bas ,. . • ' ..■• princi pi< ' " reac-
tion i o " i on ,

/8

- kol • • • ii« ■ ..h At the begi ntng ,- ,; his >lan,

Kibal * ehich aut! ijavc his cason for s lect i

w< kt tig laid nd •• do • - - of one: \
. > • -i .. • --.•--- ti cs n ■ • rate cnci ins

] ••.•;• [uant i - > in • •. • hi : periods i ' : in v ; iplosi
(•'or t!?< •• •• ; 1 uf . he : I it >d the ideal o ' . • '•■ : hi lit

snj blast -'.-...■ nd th< eed t< assure ; -ogrammei
of hum in oi the powder. The plan studied "i I feed rig
:•.-•• <■'■■ d -.,',-' • ,- • . ' * problem • , r, too. He

. . . c cd that powd • harges b . - to the o >us t ion ch
; ;-• lutoi ti< c • ='.-•• k uinisms . N. I , Ki hai ' el sals

v i ; • . ' he p rob J , . • • • ,' ■ I f light u a ted

flight cou J '• r =" '. i i both hv n >pci i v c it of

•*, : • / . . >, fhe • > ilso dis •

p oh I i \ '. blowing, th ■ hi • ' u\ on descent. At the end
h is expi i ■""••, tl uthoi . ' forth th< op : m th

ion of th problem J< en : ,, , he •• v lect i
the re* ti * u| ; rwec tie j nad nass , } • ,..•,• , -• of
- • •• i it miet ': oi tl e »mhus j n • •• nl er -- the mai
t ion o ' ! he ipp rat us .

the

ng his

. ' He

• h

res.

y of

mi de

a id

amber

o

hat

rtass

ed the

of

at

on o f

powder

n por-

Th • .-■ ; •• 1 , . ■ ch el
: •: ' ' • ■ • ■' . - ' . o f a

L s a v e i , - '• . , ■ ,
f 1 igh - hide, •■ ib <

to

to perform experiments, the author developed his idea on the
basis of guesses and scientific calculations.

We have presented a brief biography of N. I. Kibal'chich,
since the life of this remarkable son of the Ukrainian nation
has been described repeatedly in the popular and special
literature. However, even now certain facts remain unclear.
Some historians, for example, consider the question of Kibal*
chich's place of residence during the last days of his life
still unanswered.

On 20 January 1960, a memorial museum was opened in the
home where Mikola (Nikolay) Kibal'chich spent his childhood.

The name of Kibal'chich has been given to a crater on the
far side of the moon.

1.2. The Works of N. Ye. Zhukovskiy and I. V. Meshcherskiy /£

During the second half of the last century, 1830-1890, the
first works of two outstanding Russian scientists appeared --
Nikolay Yegorovich Zhukovskiy and Ivan Vsevolodovich
Meshcherskiy. These studies were dedicated to problems of
reaction-powered motion.

The founder of modern aeromechanics and hydromechanics,
Nikolay Yegorovich Zhukovskiy, was born on 5 January 1847. His
childhood was spent in the village of Orekhovo, in the Vladi-
mirskaya region. N. Ye. Zhukovskiy received his secondary edu-
cation at the Fourth Moscow Gymnasium. After completion of the
gymnasium, he entered Moscow University, where he participated
from his very first year in the work of the club which later
became the Moscow Mathematical Society. Graduating from the
University in 1868, N. Ye. Zhukovskiy, who always dreamed of
becoming an engineer, entered the Peterburg Institute of Rail-
roads 1.

Beginning in 1870, N. fe. Zhukovskiy was an instructor of
physics at the Second Moscow Women's Gymnasium, until in 1872
he transferred to the Imperial Moscow Technical School (now the
Moscow Higher Technical School imeni Bauman) . At first, N. Ye.
Zhukovskiy taught mathematics, then for 47 years -- mechanics.
It was at this school that Nikolay Yegorovich began to study one
of the most complex and interesting sections of theoretical,
physics -- hydromechanics. The results of his first studies
were published by N. Ye. Zhukovskiy in his dissertation "The

"Astronautics," Moscow, Sovetskaya Entsiklopediya [Soviet
Encyclopedia], 1970.

Kinematics of a Liquid Body." After an outstanding defense in
1877, Nikolay Yegorovich was awarded the degree of Master of
Science. In 1879, N. Ye. Zhukovskiy was selected as a super-
numerary professor of analytic mechanics by Moscow University.
In 1882, he published his original work "On the Reaction of
Inflowing and Outflowing Fluids," in which he first produced
the formulas for determination of the reaction force of a stream
of fluid flowing from a moving vessel. His monograph "The
Strength of Motion," written in 1887, won N. Ye. Zhukovskiy
the degree of Doctor of Applied Mechanics.

Ne. Ye. Zhukovskiy was given great latitude for comprehen-
sive scientific activity, both in the technical school cad in
the university where later, in 1891, N. Ye. Zhukovskiy was made
an ordinary professor.

By the end of his life, N. Ye. Zhukovskiy had become the
organized leader of the domestic school of hydroaeromechanics.
Constantly developing the theoretical principles of the
mechanics of an incompressible fluid, N. Ye. Zhukovskiy pub-
lished works between 1890 and 1907 which laid the foundation
for a new science -- the dynamics of the flight of aircraft.
In 1902, under the leadership of N. Ye. Zhukovskiy, one of the
world's first wind tunnels was created, in 1904 -- the first
aerodynamics institute in Europe, and in 1910 -- the aerody-
namics laboratory of IMTU. In 1908, Zhukovskiy published his
work "On the Theory of Vessels Powered by the Reaction Force of
a Stream of Water."

The Great October Socialist Revolution opened a new stage
in the development of domestic aviation science and technology.

In 1918, the Central Aerohydrodynamic Scientific Research /10
Institute (TsAGI) was organized, headed by N. Ye. Zhukovskiy.
The theoretical courses of MVTU served as a basis for the crea-
tion of the Aviation Technical School, converted in 1921 to the
Institute of the Red Airforce [IKVF]. In 1922, based on this
institute, the Military Air Academy imeni N. Ye. Zhukovskiy,
now the Military Air Engineering Academy imeni N. Ye. Zhukovskiy,
was created.

V. I. Lenin, beginning in the very first days of Soviet
power, constantly followed the work of N. Ye. Zhukovskiy and his
scientists and gave them comprehensive aid. N. Ye. Zhukovskiy
was called by Vladimir Il'ich Lenin the "father of Russian
Aviation."

The works of Nikolay Yegorovich in the area of aerodynamics
and flight served as the theoretical basis of modern aviation
science.

N. Ye. Zhukovski) >poko out publicly
on ' . probl > tct ion mol n for air-

crafi 3 j e firs tiros on , Move ribe-
188 ! at a mee t in >i 1 he Pol) I .• hn*
Soci ty ol tin Mos ■ - i ser Tecl tl

School [M\ • •■ ; in coo ic • . m < i th ,e
pub I i cat i oi 1 he ! . hi -e ' On \i ro-
5 tits" by V. Merchinskiy. N, Ye Zhu-
kovsl • • ' the i fa sis tei
dev:-' desc ibed in thi brochure, '. • -.ed
on th tse of the react ioi '.', i stream
of m< cui . . tn$ roni I o control

the verti' •; tnotic - oi i aerostat.

I • o v - . ; * o I H 8 i 1 1 j me c t i > o f
th Pin al Sciences : - * • . * *he

Soc = et> fc :ura 5c j do - • hi lasts

in !; »cow, M. Ye, Ihuko fski) rea* ; 'port
"On • tea ' ' it f ] ng an J it flow-

ing luii ' ^rl< epr< rtted s ig-

ni f t m ' : n ; . r • fc n c .' thcoi •- o [

reaction ot i< -J. In it, N. Ye, Zhukov-
skiy reported the results of his own studio's on the determination
of the reac io force ting on a ves: uhmerg< i a f 3 u id

b> Cor ing oul or : :l ■ in a flui ' thi "*• ; ■ - ube.
He showed ••:• • i overs e react i >n ' -cac.t ion upon suck i - in of
a f 1 • > ' very si ight in comp; . :o the direct rem on

c :• - .. J - •• pud $ on of ... fit

'. I c la; Vegoi »vich
Zhukovskiy

Thus , N. Ye, Zhuko -skiy
in and <\ Hi ng • a ■ r , a ves
• - , - • - p p o s i t e 1 ; c xp

rme . tits

• t to tl - . '- »c nt . On
tJ < Mathematical "ty, X.

to th • • hi m n b • •; r< • ort "
; • . • ' f ; lui ," pub J i shed in
uko ;ki m ••.'■ thai he re a
eg I i gib le u «u ! r how i t
laine< *. '.-,.•: for
ph> ;al st an ' * ' the fa
i n t ; - • • f i >m a J
speeds, whereas upon expulsio
a sp - * vh , ' is null pi ied b
the • : • . nfoi c rodu ed.

. •■•,: d thai ty ceessi vely »ucki \\%
scl c i .•• ••■ i I m< ; i t the
ulsic )f i • - . • fhe cport

of expci •• nt; , hi. ch .<•• -. : . iison-

17 Deci rtl : i 85 a i: nc , i t of
Ye. Zl , • •-• ono nor« ;tus ed
The {■•, t ion c Inf 1 < -■ •■ • ai d - ut-

18 ''; h I thi- report, N\ Ye.
ct ion of the in ft > i . •; fluid is
is dra «n i n i o I he vet scl. He

ce ol . ' di reel react ion ron the
ct that the exi srn; . li |uid miss
1 di j „ ■ on; il * . iti nu > . ; ; * • rying
n a dirci ed i earn is fo ied with
y the i lav r; : e r - ; e ond I o -,i ve

■] 1

jhun I t ■.:*_ i_z Kb m. /'•. ; i, >.o. 4, Nov, 1882, pp. 470-4 -.
- ,. ; -. Shorn, e- ; , Vol. XII, Mo. 4, pp. 78 7%.

In his work, N. Ye. Zhukovskiy theoretically predicted a
number of possible flight trajectories of an aircraft, in par-
ticular the "dead loop " In 1904, he discovered the law deter-
mining the lift of an ai -craft wing and published the results
of his investigations on this problem in 1906.

His final work on the theory of reaction engines was the
article "The Theory of Vessels Driven by the Reaction Force of
a Stream of Water," printed in 1908. It presents an objective
analysis of the problem of the reaction force for vessels of any
shape, submerged in a fluid and moving at arbitrary speed, with
fluid flowing in and out of the vessel. In this report, N. Ye.
Zhukovskiy avoided the error of certain scientists: he noted
that the phenomenon of reaction must be studied together with /12
the factors influencing the resistance to the motion of the
vessel, and analyzed the change of this resistance as a function
of the point where the liquid was drawn into the vessel.

Ivan Vsevolodovich Meshcherskiy made a significant contri-
bution to the theory of reaction motion.

Ivan Vsevolodovich Meshcherskiy was born on 29 July 1959 in
the city of Arkhangelsk. After secondary school, he entered
the University of Peterburg in the Physics and Mathematics
Department in 1878. Here Ivan Vsevolodovich showed great inter-
est in scientific research work and, after his graduation in
1882, he remained at the University. In 1890, he began his
teaching activity as a teaching assistant at Peterburg Univ
sity. In 1891, I. V. Meshcherskiy was selected as the Hea
the Department of Mechanics of the Peterburg Higher Courses igr
Women, and in 1902 he headed the Department of Theoretical
Mechanics of Peterburg Polytechnical Institute, where he worked
through the rest of his life.

The name of Meshcherskiy has been given to one of the
craters on the far side of the moon.

The most important works of I. V. Meshcherskiy were dedi-
cated to a new section of theoretical mechanics -- the mechanics
of bodies of variable mass, the basis of rocket dynamics. The
significance of this science results from the fact that it
allows precise calculation of the motion of a rocket and deter-
mination of conditions under which rockets will reach given
orbits or trajectories with the minimum expenditure of energy, /13
and allows many problems related to the creation of rocket
engines to be solved, leading directly to success in the pene-
tration of space.

Ivan Vscvolodovich
Meshcherskiy

reactive force is equa
speed of the particles
forces will act on the
separation of particle
attachment. The final
by I . V. Meshchersl iy
Variable Masses" (1918
syr-tem of poi nts with

The first studies of 1. V. Meshcher-
skiy on the theory of the motion of bodies
of variable mass became known in 1893
when he read a report to the Mathematical
Society of Peterburg on the theme: "One
Particular Case of the Theorem of Gulden."
The principles of this theory were set
forth in his master's dissertaion "The
Dynamics of a Point of Variable Mass,"
which he defended in 1897. In this work
for t e first time an equation was pro-
duced for the motion of a point of vari-
able mass for the case of separation or
attachment of particles, in particular
the vertical motion of a rocket. In 190 4,
I. V. Meshcherskiy completed his work
"The Hquationsof Motion of a Point of
Variable Mass in the General Case," pre-
senting a general theory of motion for
the case of attachment and separation of
particles. It was shown in these works
that when particles with zero relative
velocity are attached or separated, the

1 to zero. However, if the relative
is not equal to zero, supplementary
body: a reactive force in the case of

s and a resistance force in the case of
work in this direction was an article

entitled "A Problem from the Dynamics of

), in which he studied the motion of a

variable masses.

I. V. Meshcherskiy was not only a remarkable scientist, hut
->lso an outstanding teacher. lie fundamentally changed the
t. -aching of the course on theoretical mechanics, bringing it
closer to the needs of practice. lie trained engineers, many of
whom later became great scientists, including specialists in
me area of rocket and space technology.

1.3. K. H. Tsiolkovskiy , the Founder of Astronautics

The idea of flight with rockets, i.e
moved by the effect
tion of the mass of
scicnti fie Has i s i.r

flight vehicles
of the reaction force, arising when a por-
tne vehicle is expelled, was given a deep
the works of the outstanding Russian

scientist, Konstantin hduardovich !'•• i olkovsk iy (1857-1935)
K. I;. Tsiolkovskiy suggested a rocket with ,i liquid fueled
motor, theoretically studied some of the specifics of the ©por-
tion of individual units and of the engine as a whole during

/14

in

flight

of LRli
by K.

v a r i a b

theory
Is 10 Ik
origin
voyage
space ,

in space, created the principles of the theory and design
One of the fundamental problems developed and studied
li. Tsiolkovskiy was the movement of rockets as bodies of
le mass; these studied laid the foundation for the

of rocket flight. The scientific activity of K. li .
ovskiy was mult i faceted and unique. To his pen belong
al works on aerodynamics, the theory of interplanetary
s, work on the problem of life on artificial islands in

on biology, geophysics and philosophy.

Konstantin i-duardovich Tsiolkovskiy
was born on 17 September 185/ in the
village cf 1-hcvskiy, Spassky district,
Ryaianskiy region in the family of a
forester. His childhood years were marred
by serious illness; at the age of 9, he
almost lost his hearing. This made it
impossible for Tsiolkovskiy to enter
school. The mothei of Konstantin
Hduardcvich -- Mariya I vanovno Tsiolkov-
skaya-Yumasheva [1830-1870) -- taught her
son herself, teaching him to love work.
At age 14, Konstantin Hduardo i ch began
to study independently, using the books
of his father fduard Igant ' yevtch Tsi ol-
kovskiy (1820-1880), who noticed the
capability of his son and his love for
scientific experiments. His father helped
him to make balloons, to construct models
of various machines, inspiring young
Tsiolkovskiy with a love for technology,
nature and experiments, teaching him to
analyze the results even of the simplest
experiments .

Konstantin fduarclo-
vich Tsiolkovskiy
(1899 ph to)

T< i olkovsk iy was
Moscow. During these
an interest in space flight

16 years old when his father sent him to /15
years, Konstantin F.duardov t eh developed

In 1879, Tsiolkovskiy passed the examination and was named
a Teacher of the People's School and in 1880, he began teaching
arithmetic and geometry at the Borovskiy School in Kaluga
region. Here Konstantin Hduardov i ch Tsiolkovskiy performed his
first studies on the subject of interplanetary voyages.

The dream of traveling in space did not leave the scientist
throughout his life. His studies on the theory of reaction
motion are broad in scope and show a surprising combination of
strict mathematical analysis with brav^ flights of fantasy.
K. h. Tsiolkovskiy believed in the immortality of developing

11

mankind, and all his creative activity, in the final analysis,
was dedicated to seeking out means for the improvement of the
living conditions of future generations.

In order to solve the primary problem -- the overcoming of
the Earth's gravity -- the scientist had to solve many problems,
widely varied in content and complexity.

In 1881, K. E. Tsiolkovskiy worked on problems of the
kinetic theory of gasses. For his work entitled "The Mechanics
of the Animal Organism," he was selected as a member of the
Physical-Chemical Society.

Beginning in 1883, Konstantin Eduardovich dedicated his
time primarily to problems of flight in the air and in space.
On 20 February of that year, Konstantin Eduardovich completed
the manuscript to "Free Space," in which he described the
properties of the medium and the conditions of movement in
space. Here he analyzed the design of a "shell for voyages in
free space."

His works on the design of an all-metal controlled
dirigible became widely known. He set himself the task of creat-
ing a metal controlled aerostat, turning his attention to the
essential shortcoming of dirigibles with balloons made of
rubberized materials: these envelopes wore rapidly, represented
a danger of fire, were low in strength, and the gas which filled
them diffused through the fabric and was rapidly lost.

The progressive, for its time, dirigible plan was not
supported; the author was not even given a subsidy to construct
a small model. In order to test a number of his own calculated
data and prove the possibility of constructing his dirigible,
K. E. Tsiolkovskiy made a model at his own expense. /16

In 1897, K. E. Tsiolkovskiy constructed the first wind
tunnel in Russia, developed his model testing methodology and
a chived interesting results. In 1900, K. E. Tsiolkovskiy tested
several models which he had made in the wind tunnel and deter-
mined the drag factors of bodies of various shapes.

In 1895, his science fiction story "On the Moon" and the
work "Dreams of the Earth and the Sky" were published. The
first work, in particular, describes how people who found them-
selves on the moon would feel, while the second work, in addi-
tion to presenting many original thoughts, sets forth the idea
of the creation of a "falling laboratory" and describes various
phenomena occurring in weightlessness.

12

This idea of reproduction of the conditions of weightless-
ness is based on the fact that if a man is placed in a flight
vehicle which moves toward the Earth at an acceleration equal
to the acceleration of the force of gravity, the force of the
interaction of the man with his support (the wall of the cabin
of the flight vehicle) will be zero, i.e. , the acceleration will
be equal to zero, and the man will be under conditions of
weightlessness. A state near weightlessness is experienced by
a pilot at the peak of a climb. "Falling laboratories" are
presently used for training of astronauts and to study phenomena
occurring under conditions of weightlessness.

The style of the work of K. E. Tsiolkovskiy is distinctive
and unique. His persistence in seeking out the most convincing
and simplest (and consequently most possible) solution, his
tendency to produce a clear picture both from the physical and
mathematical standpoints -- these are the characteristic fea-
tures of the style of K. E. Tsiolkovskiy which have made his
works understandable, readable and convincing.

Konstantin Eduardovich wrote, "I have been studying reac-
tion devices since 1895. Only now, after 34 years of work, have
I come to a very simple conclusion concerning the proper system. "1
And further, "In 1896, I purchased a book by A. P. Fedorov
entitled 'A New Principle for Air Travel' (Peterburg, 1896).
This book seemed to me to be unclear (since no calculations
were made). In such cases, I perform my calculations indepen-
dently, from the very beginning. This was the beginning of my
theoretical studies on the possibility of using reaction devices
for space voyages." 2

In 1892, Konstantin Eduardovich moved to Kaluga. Years /17
filled with productive, creative labor passed in this city.
K. E. Tsiolkovskiy produced his formulas for rocket travel,
allowing him to solve the problem of the most realistic method
of mastery (study) of space in the theoretical plane.

Konstantin Eduardovich Tsiolkovskiy lived 29 years in the
house on the corner of what is now K. E. Tsiolkovskiy Street and
Sovkhoznoy Street. Konstantin Eduardovich bought this one-story
house in 1904. In 1908, the house was expanded, adding a
second story --a sun room and veranda. In the fall of 1933,
the family of the scientist moved to a large, well-built home
given to K. E. Tsiolkovskiy by the Kaluga City Council of
Worker's Deputies.

^Tsiolkovskiy, K. E. , Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 296.

2 Ibid. , p. 179.

13

The Home

of K. H

bicycle

**hich Ts

room was

his off

eh.ti rs .

A iv i re

kerosene

lantern

workbenc

i. Here

specimen

> of cor

On 19 Sept em -
b e r 1936, a memo rial

museum was opened in
the old home. Most
visitors to the
museum begin on the
second floor. There,
in a small entrance
hal 1 , is a port ra it
of the scientist at
age 75, A small
showcase contains
personal belongings,
photographs of the
family, portraits
of scientists who
were friends of
K. V.. Tsiclkovskiy.
Here also is the
iolkovskiy used far into his old age. The sun
ice and bedroom. By his desk are two comfortable
is strong across the room, and from it hangs a
On the veranda is a homemade lathe and small
also is a bending machine with wooden rolls and
rugated sheet metal for a di r iglhle ,

Ts iolkovskiy in Kaluga

On the first floor, in what was a storage room, the works
of tin- scientist on dirigible building and aerodynara : --s are
displayed. The displav stands show models of the all-metal
dirigible, equipment for manufacturing its shell. One small
room on the first floor contains the model wind tunnel made by
K. E. Tsioikovski v .

/IS

A larger room contains the published works of K» li, Tsiol
kovskiy, models of his rockets, drawings, plans and literature
on his works. In the garden outside the museum house is a
bust of K. b. Ts iolkovskiy.

Considering the great interest in proble
technology in Kaluga, construction was undert
building for the museum, which wa-> dedicated
Today, the K. II, f> iolkovskiy State Museum of
Astronautics stands on Academician Korolev St
of the trees of an old park. The Museum Tstan
the Oka and Yachenka Rivers, with a view to t.
limitless spaces of Russia merge here with th
space, to the glory of man and his inexhausti
t'he Museum of the History of Astronautics • •
tells the story of the life and activity of K
vich, the development and practical appiicati

ms of rock

et

aken of a

new

on IS June

1961.

the History of

reet in the shade

ds on the

banks of

he horizon

. The

e i n f i n i. t y

of

ble imagin

at ion.

"the Space

['a lace" -

onstant in

Hduardo-*

on of his

remark able

14

; - i'h cases ; J (iisji ••• ic lp the v t s if 01 to fami 1 i art :
h i •..• J wit! i U ' wo o! th< - Lcnti • »

r i •

o f

Kal

K. ..

the Hi
. :a

Ts i o

story

of As

of using atomic ener

; i , io , in his

work "1 lives 1 • ; on of
Space lb i i - React ion
Ik ic ; ," 1 . nd in in my
other wo e* h ! . h< U

in ; . > i- ; i: )1 : i ml
19 lis f as we I ; is ' : ipace
Ro ■ •• : I .-- ,,-.■ ,' pub -
lisJlt I l 1 > », K. !:.
F> oil ski ■ :• iy
an,' . tel ; . , the

fat . hi •' . oo - ' the
the i i •:■■ ■ ••>. k ". 1 1 ;ive I »
Jo >.:"■' • ', he p rim pies
of th i ,ii i)l" roikcf
and • • k* t ■• -,' -, s
bin - - i -v ,":<"-. fue I ,
He stu - : •. •• • . on
me ] Ic •" '>. ; i o ■ ue ' : for
rocket e-e. i rtes . K. 1- .
Ts : •"•; koi *• ! > • y I - j sot
exc <•' de 1 1 e pos: ibi i i ty
ity in . ocki t t < mo ' o ; .

i y 5- ife "•': ; ■
t r o n t ■ ■ in

Th • , i ; ..-,'.• ki a iso stud icd p rob len elated to
h mas to oi hit rpianetar 1 space ll< pro ed tha front
ten netan ■ " ' *. ; :>. i • ;* '. to the pi ai • . to I . j 1

c t. Id .,.:: - , . , , •, mei in .■ ■ to the us

-. no r i ici- : ' ... ;c foi I h • nci J - I m on - »d .

K. V. . . • ••.;,, igi to J that - • * . -. m tor i a 1

to he • - • ; i e! 1 1 • ; should In h st< •. - • i Iso

• : : • I • , - • ■• : I : ' as ' , ,' movin I t h< in in elliptical

orbits, mostly I itcv •■ - , • ; . orbit? • , Mar: ml Jupiter.

The total ma the sterol . cunt I shoo •-.'•• of

- • • of the a ' ' ie • -,„..•■ u I e to ids m a fev* ! u • ed

'-■ f k t lomt - rs in di mi ! • ; ' orae sf the i I planet? lur in

ho pn .'« of f i .•,-..-■ cork at the s . . to the liar th.

I ie anal 1 Line tear i i I -,,..,- > ■ • . • , ,. i i 1 1 j on

. - or * • ■ ' th* hi) - 1 - '*•.., »'hi le llei nt oe - ' . '. •, . 'Ji

tad tanee of aboi ' ., in h'.U (th ; a.ua om

the Ihu li h tin h m is *'--_;,•• km) .

f!9

" ! ; i rs? r intet 1 1 th • imal moil so) Obozren i , - o, 5,
:. M)3.

BcJ'oi :• tin
I iolki • . iy ai

in t! e : ,'iet
tion, "Under

te Ore • ■
Hid his

$ t v t e w
the Sot

wrote, "I eou : i i -. c
red ■ • : .
•011 .

i Lmos 1 i 110 ic
In 1 U9, he wa
fommunist Wad .*. S
From teaching in the
years and poor 1 ?alth
5 tat ' Jt ree hub I ic
cc ti runt J v.

and
in >rk ..."

c tube i So ial • ■ I* .- •*• -n, K. ti,
ivorts iirrc no proper Jo a* know J Iged, Only

. '•■ , . • • • . •'- concern a I at ten -

let ,"

'.;•■ ->.' 1 :' id it i ■• t ■■

re, ire ork; nov att j :ic ted a I nt ion. "
i ik ' . • s m< ml er of th ? S n. y ; i r, I tci
inee in 1920, K. F.. '! iolkovskn ret re •

ndar> ;hoo due ti hi sgt )1 3
, he wa s ! '-> • J .' ■ .• a ' •■ • " on by a

orcein!

nations ' .; i K. \\. T: iolkovskiy

h> thoi -■: o , !:. Tsiol t I iy, .set )rt] n hi books,
still - ;■ ,. . hi : .:.'• 1 oj hie , hi> math ma i si sti ic 1 c- . and
the a- i u i ac ■ u 1 hi : - s ion .

Betu . : '• •• nJ l! , four t im< men lii s art ic les ,

hrocl . • , ook • > pul died

r . . than to v !u nt i re >re-

revol t Lunar ,- : . s - en e f'roi , ! 5 to 1935 alone,

ahout 60 worl in- Li. 1 dkovskl; *en publish I on ph; ;ies,
astro-- ut ics , a : ro i n> , -,-. ■,-•.■ and id. osophy .

In 1952, th- , tire count i celc-
brai • li i - • . u liday and I s
Si) ih yea r o! :rcat i vi c i i I , : . ic
ac v'ity. he USSR \- i -. my of
Se em • •;- he Ida ;ol< inn iee i rig in
ceJ bi t i •• -. >H .".• and I -

ko-. s'kiy it ti nded .

i ted the »f> '-nee

more, when M I Kalinin awarded him
th Red kmnei Laboi ,-•• - : !*oi ?5

cr iv< ~ . * i 3 h i country.
Ac- -',-.'••• h aw; , K. V. . I's iol kov-
sk c rati , "I in tha £ eovern-
iiie s - c i h ,- : h ij»h .- ••: : .- " y

further work..."

P tr ini I he . :i •■ yea ■: : o the 1 i fe
of L I ; . Tr i o I kovsk iy , h ;- ■ o? '- . d
to- - -. . u .-•.-. '.. s

ideas , f re<|uet5t \) ec tu in pp :ar-

• te. 1 i a ■: i ■■. : i> - •.•.■!>-• -in.i : r.i J i n>.
report ^ be f« i sol • . , worker re tent i :;1 ami fa r«a i - .

Konst ! - • i'.duai '-'or ich
Tsio i kov ■: '■ i v

Tin vork: * f K, I; , i'r io 1 kovsk i h; eon * i bed by

a sent i st s as V. P, •'.- .-.•.,.--., N. A. Ryni ■• - i . I. i'erel ;ian,

V. V. Ryusiiin and oth« m esse • iiui sopul sri; r< oi •; ; - .-.-as
-.'«.,-*,:. i - - S, P. kovolcv, i f ikhonr ov,

/20

A. A. Kosmodem'yanskiy. Their books, brochures and articles
have carried the ideas of the scientist to the masses, stimulat-
ing interest in problems of the study of space and multiplying
the ranks of rocket technology enthusiasts.

Beginning with the first weeks of existence of the Reaction
Scientific Research Institute [RNII], which was created in
Moscow in September of 1933, scientific contact and fruitful
correspondence were held between K. E. Tsiolkovskiy and the
Institute. For example, in February of 1934 he composed a "Pro-
gram for the Work of the RNII," and in March of that same year
he wrote his article "The Energy of Chemical Compounds and the
Selection of Component Parts for an Explosion," etc.

In August of 1935, K. E. Tsiolkovskiy *s health began to /21
deteriorate. On 13 September 1935, the scientist sent a letter
to the Central Committee of the Party. "All my life I have
dreamed that my works might move mankind forward, at least a
little. Before the revolution, my dream was impossible. Only
October brought recognition to my works... I have felt a love
for the people, which has given me strength to continue my work,
even in my illness. However, my health will not allow me to
finish the work I have begun. All my labors on aviation, rocket
flight and interplanetary voyages I bequeath to the Bolshevik
party and the Soviet government -- the true leaders of the pro-
gress of human culture. I am sure that they will successfully
complete my labors. "1

Exceptionally valuable and progressive works of K. E.
Tsiolkovskiy are his works on reaction motion, which preceded
the development of science in this area by many decades. K. E.
Tsiolkovskiy first developed the laws of motion of a rocket as
a body of variable mass, indicating efficient paths for the
development of astronautics and rocket building. He found a
number of important engineering solutions to problems of rocket
design, he analyzed and recommended fuels for use for rocket
engines. K. E. Tsiolkovskiy laid the foundations of the theory
of LRE.

A number of the technical ideas of K. E. Tsiolkovskiy were
applied in the creation of modern rocket engines, rockets and
spacecraft.

Since 1966, on 17 September each year in Kaluga readings
are held dedicated to the development of the scientific
heritage of K. E. Tsiolkovskiy.

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 20.

17

The readings are conducted by the State Museum of the His-
tory of Astronautics, the Commission for the Development of the
Scientific Heritage of K. E. Tsiolkovskiy of the Academy of
Sciences USSR, the Astronautics Committee of DOSAAF USSR, the
Institute of the History of Science and Technology, Academy of
Sciences USSR and the Institute for Medical and Biological
Problems of the USSR Public Health Ministry. Soviet scientists
discuss the most pressing problems of missile and space techno-
logy and rocket engine construction at these readings.

The Works of K. E. Tsiolkovskiy on the Creation of
the Theory of Reaction Engines

One of the achievements of K. E. Tsiolkovskiy is the deter-
mination of the expediency of the use of LRE as spacecraft
engines. He suggested a plan for a rocket equipped with an
LRE, determined the areas of application of such engines,
selected and evaluated various types of rocket fuels, i.e., Ill
substances or combinations of substances to sarve as the source
of energy and working fluid for a rocket engine, studied some
of the design peculiarities of individual units and of the
engine and its operating conditions, and noted the main paths
to be followed in the creation of powerful liquid-fueled rocket
engines.

The power of a rocket engine is equal to the kinetic
energy of the mass of gases in the reaction stream flowing from
the reaction engine per second, or half the product of the
thrust times the effective exhaust velocity. The power of the
engines of modern booster rockets reaches tens of millions of
kilowatts.

K. E. Tsiolkovskiy pointed out many scientific and techni-
cal problems which had to be solved during the course of further
development and improvement of rockets and their engines.

The selection of a plan for a rocket engine is a difficult
task. In order to properly select a plan, one must consider
the values of the fixed parameters of the engine, the purpose
of the rocket, its range of flight, the level of technology
in the country and available experience. If the problem is to
be solved today, one plan will be suggested; for rockets of
the future -- another. K. E. Tsiolkovskiy suggested a plan for
a rocket and a rocket engine for the future, considering the
possible progress of science and technology. He believed that
the time had come to begin such development. Later events have
confirmed the correctness of his views. Modern engines and
rockets do not differ in principle from those he suggested: a
two-component liquid fuel, pumped fuel feed, acceleration of
the gas jet in a nozzle, etc.

18

In his work "Investigation of Space with Reaction Devices'*
(1903), K. E. Tsiolkovskiy described the plan and operating
principle of an LRK using liquified gasses as components in the
following words. "The chamber 1 contains a great reserve of
substances which, when mixed, immediately form an explosive mass.
These substances, fully and evenly exploding in the area set
aside for this purpose, then flow as hot gases through tubes
which expand at the end like a horn or other musical instru-
ment." 2

The combustion chamber of a rocket engine is the most
important part of the rocket engine, which creates the reaction
force due to the flow of the working fluid. A modern rocket
engine consists of a combustion chamber and nozzle. The nozzle /23
is that portion of the rocket engine in which the thermal
energy of the compressed working fluid -- the combustion pro-
ducts -- is transformed to kinetic energy, i.e., the gas jet is
accelerated to the exhaust velocity.

Further, K. E. Tsiolkovskiy wrote, "In one, narrow end of
the tube, the explosive substances are mixed: thence, flaming
gases are produced here. In the other, expanding end, these
gases, greatly rarefied and cooled, burst outward through the
aperture with tremendous velocity.. The two fluid gases are
separated by a barrier. "^

In 1922, K. E. Tsiolkovskiy wrote an article entitled "Star
Flight" for the magazine "Znaniye-Sila," in which he described a
rocket with an LRE designed to be used as a jet aircraft. For
this purpose, the rocket was equipped with wings.

In 1927 in Kaluga, K. E. Tsiolkovskiy published his work
Kosmicheskaya Raketa. Opytnaya Podgotovka [Space Rocket. Exper-
imental Preparation]. This work presents a still more detailed
description of an LRE; it is pointed out that the fuel components
must be fed to the combustion chamber by "...two pumps, driven
by a single engine. The first pumps the oxygen compounds to the
combustion chamber, the other pumps the hydrogen compounds."^
Here also we find the idea of maintaining a certain ratio of
fuel components during the operation of the LRE: "regulation is
important: if there is more oxygen than needed, the combustion

Having in mind the interior of the rocket.
2
Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of

Sciences USSR Press, Moscow, 1954, p. 73.

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 261.

4 Ibid. , p. 75.

19

chamber itself might burn, if there is less
expended uselessly."

- the fuel will be

In this same work, K. E. Tsiolkovskiy describes the opera-
tion of en LRE and studies the conditions needed to ensure
safety. If a rocket were made according to these plans, pub-
lished by Tsiolkovskiy early in the 20th century, the unused
volume of the rocket would be very slight, since every free
space, not occupied by structural elements, is filled with fuel;
the LRE is submerged in the fuel components. This arrangement
provides the minimum mass and size of rocket.

In his work "A Semireaction Stratoplane," first published
in the magazine "Khochu Vse Znat ,M in 1932, K. E. Tsiolkovskiy
wrote, "In the lower layers of the atmosphere, an aircraft
cannot reach a high velocity. ...my ideas and calculations have
led me at present to the following, most possible type of high-
altitude aircraft. "1 Further, K. E. Tsiolkovskiy presents a
description of a jet engine^ driving a propeller.

The design of
the "semireaction
stratoplane" devel-
oped by K. E. Tsiol-
kovskiy was as
follows. As the
device moves, air
enters the internal
portion of the body
through adjustable
inlet aperture 1.
The gas stream is
accelerated by pro-
peller 2, driven by
gasoline engine 3.
The spent gases move
through tubes 5 and
flow out of their
exhaust sections.
The air and spent
gases exhaust through
adjustable nozzle 9.
Air compressor 8 is

Rudders

- czr i-

/24

/25

Plan from "Star Flight" by K. E. Tsiol-
kovskiy

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of

Sciences USSR Press, Moscow, 1954, p. 389.
2
A jet engine refers to a reaction engine which utilizes the air

around the vehicle to burn the fuel.

20

Liquid, freely
evaporating
oxygen at very
low temperature

Men and
equipment
for
breathing

Liquid
hydrocarbon

Plan of Liquid- Fueled Space Rocket of
K. E. Tsiolkovskiy

Plan of "Semireaction Stratoplane" of
K. E. Tsiolkovskiy

Shaft of propeller

and other devices

For mixing
Air compressor

Mixing 6
combustion

Plan of the "Steam- Gas Turbine Engine
for a Dirigible" of K. E. Tsiolkovskiy

mounted on a common
shaft with engine 3.
Air from cavity 4
partially enters
cavity 10, then
cavity 6 and, through
annular space 7, it
moves past tubes 5.
Washing over these
tubes , the air is
cooled and enters
the compressor. The
compressed air
flows through tubes
11 to the gasoline
engine.

Finally, in his
work "A Steam- Gas
Turbine Engine,"
published in 1934,
Konstantin Eduardo-
vich suggested a
unique turbocompres-
sor engine, which he
suggested be used for
dirigibles. This
engine is a prototype
of one version of
modern jet engines.
In this engine, the
incident air stream
is sent by means of
compressor 7 and
diffusor 1 into
generator 2 under
pressure, where the
oil fed into the
generator by a pump
(not shown on the
drawing) is burned.
The combustion pro-
ducts spin multiple-
stage turbine 3. The
rotation of the tur-
bine is transmitted

through system 4-5 to a propeller, which drives the dirigible.
Furthermore, the rotation is transmitted by system 6-7 to the
compressor, and by system 8-9-10 to agitators, which contin-
ually mix the oil in order to equalize its temperature in tank
12. The generator is cooled by the water filling space 11.

/26

21

Thus, K. E. Tsiolkovskiy suggested a plan for a liquid-
fueled rocket and plans for jet engines as well. All of the
plans which he suggested were later utilized in principle in
practice.

The Formula of K. E. Tsiolkovskiy

The creation of the most efficient engine design continues
to be one of the most important problems of rocket engine con-
struction. The rockets suggested by Konstantin Eduardovich,
naturally, were not developed by him to the stage of a complete
plan. They were more like reports of new ideas, inventions,
discoveries, but reports based on scientific and technical cal-
culation.

The development of the theory of rocket engines and rockets
in the works of K. E. Tsiolkovskiy and in the works of other
authors are based largely on the formula which is known by the
name of its author -- K. E. Tsiolkovskiy.

This is the basic formula for the motion of a rocket,

defining its maximum velocity V, equal to the product of the

absolute value of exhaust velocity W„ of the combustion

a

products from the reaction nozzle times the natural logarithm of

the ratio of the initial launch mass of the rocket M„ to its

final mass M, (considering payload) , remaining after fuel mass 111

M t is expended in flight:

In calculating the motion of a rocket equipped with a
modern LRE, if the difference p - p„ is other than 0, W in the

Tsiolkovskiy formula must be replaced by the effective velocity,
which is

W eff=^+^- (*»-/>/,).

where F is the area of the nozzle exit plane;

22

G is the mass flow rate of fuel per second, equal to the
flow rate of the combustion product;

p is the gas pressure at the nozzle exit plane;

p„ is the pressure of the surrounding medium at the flight

altitude H.

The ratio M t /Mj c is called the Tsiolkovskiy number and is
represented by the letter Ts.

This formula is developed in the work "Investigation of
Space with Reaction Devices" (1903) . Using the Tsiolkovskiy
formula (in his 1903 work, K. E. Tsiolkovskiy ca 11 this
formula the "relationship of masses in the rock' , we can
calculate also the velocity increment of the in lual stages
of multistage rockets.

Tsiolkovskiy 's formula was refined by him to consider the
influence of the resistance of the surrounding medium and the
forces of gravity on the final flight velocity of a rocket.
This formula was the first step made in the development of the
requirements for LRE; during the initial period of development
of rocket technology, it allowed scientists to determine the
primary paths for improvement of the design of an engine. It
is understandable that, when modern engines are produced, all
of the accumulated experience of rocket construction and
engine construction, the achievements in neighboring areas of
science and technology are used, attempting to satisfy the
continuously growing demands on the design of rocket engines.

It follows from the formula of Tsiolkovskiy that in order
to increase the flight velocity of a rocket, one must increase
the Tsiolkovskiy number Ts and the effective exhaust velocity
of the gases W -^.

The exhaust velocity of the gases from the nozzle /28

Vw

U -v

IV? •
in

where Q is the quantity of heat liberated upon combustion of a
unit of mass of fuel;

n t is the thermal efficiency;

W. is the velocity of entry of the fuel components into

the combustion chamber;

<j> is the proportionality factor.

23

The higher the heating capacity of the fuel, the more heat
is liberated upon its combustion. However, the same fuel
components, depending on conditions, liberate different
quantities of heat which, in particular, depends on the ratio
of the components

k i g~»

where G is the mass flow rate of oxidizer per second;
o

G f is the mass flow rate of fuel per second.

The optimal relationship, for which the exhaust velocity
reaches Its maximum, depends for a given pressure in the com-
bustion chamber on the type of fuel, degree of expansion of
gases in the nozzle and a number of other factors.

In order to increase the completeness of combustion in
the smallest possible chamber volume, the quality of spraying
and mixing of the components must be improved. "The problem
is that the force of the explosion in a given tube* depends on
the completeness of mixing of the combustion elements. "2

The more heat which is liberated during the combustion of
a unit mass of fuel, the higher the energy characteristics of
the products of combustion -- heat conduct and the product of
the gas constant of the products of combustion R = 848/y times
their temperature T..

With a given heat liberation, as the mean molecular mass
of the combustion products p decreases, the gas temperature
decreases, simplifying the solution of one of the most complex
problems of rocket engine construction -- the problem of
effective cooling of combustion chamber walls.

The thermal efficiency /29

Having in mind t\ „• combustion chamber of the rocket engine.
2
Tsiolkovskiy, K. E., Collected Works, Vol. 2, Academy of

Sciences USSR Press, Moscow, 1954, p. 201.

24

characterizes the conversion of heat to Kinetic energy of the
combustion products flowing from the nozzle.

In order to select the best value of gas pressure in the
nozzle exit plane p , we use the thrust formula

P - GW eff .

We recall that in this formula G represents the mass flow of
fuel per second, equal in the stable mode to the mass flow of
combustion products per second. Analysis shows that in order
to roduce the maximum thrust, pressure p should be equal to

the pressure of the surrounding medium p„. If the pressure of

the surrounding medium changes during the flight of a rocket,
the equation p = p H can be maintained by changing the para-
meters of the combustion chamber or the critical cross -sectional
area, or the nozzle exit plane area.

However, adjustable nozzles have not yet been created for
LRE, forcing us to utilize a certain mean value of p , selected

during the process of ballistic planning of a rocket to provide
the maximum flight velocity at the end of the powered stage of
flight with a fixed payload mass and the selected value of
Tsiolkovskiy number.

If an engine must operate at very high altitudes or in
space, where the pressure of the surrounding medium is very
low, in order to increase the thermal efficiency, the lowest
possible pressure should be maintained at the nozzle exit plane.
If this pressure is fixed, the thermal efficiency can be
increased by increasing the pressure in the combustion chamber,
which also helps to improve the combustion conditions, decrease
the size and mass of the combustion chamber.

In analyzing the operation of a combustion chamber, K. E.

Tsiolkovskiy based his calculations on pressure p. = 100 atm.

1 i
This pressure could not be achieved by the first LRE. For

example, in engine 10 of the GIRD-Kh rocket (1933), the pressure

in the combustion chamber was only 8 to 10 atm, while the ORM-50

and ORM-52 engines (GDL, 1933) achieved 20-25 atm, the RD-107

engine (GDL-OKB, 1954-1957) produced 60 atm, the RD-119 engine,

Here and in the following, the units of measurement are pre-
sented as in the arch 3 materials.

25

developed in 1953-1962 (GDL-OKB) produced 80 atm, and later /30
engines have produced still higher pressures. Thus, the
pressure intuitively assigned by K. E. Tsiolkovskiy for the
chamber was approximately equal to the pressures achieved by
modern engines.

Comprehensive improvement of engines has increased their
economy. 1

For example, in engine number 10, the specific impulse
achieved in test stand operation (1933) was 162-175 s, in the
ORM-52 engine (1933) -- 210 s, while the specific impulse of
the combustion chamber of the RD-119 in a vacuum reaches 358 s
(1958-1962).

In order to increase the Tsiolkovskiy number

M t

Ts = sr

one should use fuel of the highest possible density p . This

maintains the requirement mentioned above for a high value of
the quantity of heat Q liberated in the combustion chamber in
each second of operation. In order to decrease M. , the parts

of the rocket should be made of structural materials for which
the ratio of strength (or yield point) to density is as high
as possible.

The Tsiolkovskiy number can be increased during planning
of a rocket by successful selection of -■« plan of motor, rocket
in general and individual rocket units and by assuring opera-
tion of the units as near as possible to their optimal oper-
ating modes. If a pressure-expulsion fuel- feed system is used,
the fuel tanks must be made with thick walls, but if a pump-
feed system is used, tanks are maintained at low pressure and
their walls can be made thin. Therefore, the Tsiolkovskiy
number for large rockets is higher with a pump- feed system than
with a pressure-expulsion system.

As we have stated, in determining the Tsiolkovskiy number
for a rocket, mass M, refers to the mass of the structure of the

rocket and its systems, including the engine, the residual

The economy of a rocket engine is defined by the specific
impulse, the ratio of the thrust of the engine to the fuel
consumption per second.

26

liquids and gases at the end of the powered portion of flight
and the payload mass M (nose portion with instruments or cabin

plus astronauts, etc.)- With a given value of Tsiolkovskiy
number, as the payload is increased, the mass of all the other
elements of the rocket must be decreased, which is achieved by
comprehensive improvement and lightening of the design, or the
launch mass of the rocket (or each stage) must be increased
by increasing fuel mass M..

The Tsiolkovskiy formula allows us to judge the effec- /31
tiveness of utilization of the fuel energy of a rocket. K. E.
Tsiolkovskiy defined the work performed by a rocket

L P ■ i V 2 ;

the work of the exhausted gases

1 2

L — -sr M.W ,

a 2 t a'

and calculated the efficiency of a rocket as the ratio of L

P
to the sum of L ♦ L .
P a

The power of engines and frequency of launches have become
so great that, considering the prospects for the development of
rocket and space technology, the determination of means for
increasing the total efficiency and its current values, cal-
culated for various moments of operation of an engine, have
become a very pressing problem. The great consumption of fuel
expected in the near future has placed the problem of the
creation of rockets with external power supply on the agenda
for the day.

The analysis of the formulas presented here led K. E.
Tsiolkovskiy to the idea of space trains. Various versions
of connection of rockets were studied: sequential, parallel
and combined; the so-called "second type" of compound rocket
of Tsiolkovskiy called for parallel connection of rockets in
groups. We know that all modern spacecraft booster rockets
are multistage rockets, with both sequentially and simultan-
eously operating motors considered tht most favorable combin-
ation.

27

Suggestions for LRE Fuels

Analyzing the properties of fuels, K. ; : .. Tsiolkovskiy
wrote, "They should perform the maximum work per unit of i ass
during combustion." And further, "For a reaction apparatus,
the greatest possible portion of the thermal or chemical
energy of the particles must be converted to coordinate! motion
of the particles. "1

In his work, "Investigation of Space with Rocket Devics,"
K. E. Tsiolkovskiy in 1903 suggested liquid oxygen and hydrogen
as fuel components for LRE. "At the present time, the conver-
sion of hydrogen and oxygen to liquids represents no great
difficulty. Hydrogen could be replaced by liquid or condeiiL^J /32
hydrocarbons, such as acetylene or petroleum. "2

In this same work, the scientist studies certain inorganic
compounds as possible fuels. "For example, silicon, burning in
oxygen (Si ♦ 2 = SiO ? ) , liberates a tremendous quantity of

heat, 3654 cal per unit of mass of product produced (Si0 2 ) , but

unfortunately forms substances which volatilize with great
difficulty. "^ K. E. Tsiolkovskiy gave great attention to the
study of the fuel consisting of liquid oxygen and hydrogen.
"Accepting liquid oxygen and hydrogen as the material most
suitable for explosion..." he wrote in the work just mentioned.
However, the scientist was bothered by the low density of
hydrogen, requiring large containers, which would require an
increase in the volume and mass of the rocket. In 1927, in
the work "A Space Rocket. Experimental Preparation," he noted,
"Liquid hydrogen is generally unsuitable, particularly for the
first time. Reasons: high cost, low temperature, heat of
evaporation, difficulty of storage." 4

In 1903, he wrote, "...the quantity of energy per unit
mass of the products of a compound depends on the atomic
weights of the simple substances combined: the lower the atomic
weight of these elements, the greater the heat liberated as
they are combined. "5

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 79.

2 Ibid ., p. 81.

3 Ibid.

4 Ibid. , p. 270.

5 Ibid. , p. 81.

28

In 1914, K. E. Tsiolkovskiy suggested that ozone and other
components be used as oxidizers in engines. "We must find
compounds of hydrogen with carbon which contain the greatest
possible quantities of hydrogen, which are formed as tht>
are produced of elements with absorption of heat, for example
acetylene, which, unfortunately contains little hydrogen. In
this latter respect, turpentine is more suitable, and methane
or swamp gas is still more suitable; this last substance is
unfavorable in that it is difficult to liquify. "1

In his work "The Investigation of Space with Reaction
Devices," 1926 edition, ■<.. E. Tsiolkovskiy compares hydrogen
with hydrocarbons: "It is difficult to liquify and store, since
unless particular precautions are taken it evaporate* rapidly.
Most preferable are liquid or easily liruefied hydrocarbons. /33
The more volatile they are, the more hydrogen they contain and
the more suitable they are for the business at hand. Oxygen
is tolerable in liquid form, particularly since it can serve
as a source of cooling..." Further, the scientist notes, "But
it is most suitable to work as follows: store most of the
reserve of oxygen on-board in the form of one of its endogenic
compounds, i.e., those which are synthesized (made up) with
absorption of the material." 2

In this same work, in 1926, methane, benzene and oil are
recommended as fuels. In 1927, liquid air was recommended as
an oxidizer: "Initially, liquid air can be used. The nitrogen
present will weaken the explosion and decrease the maximum
temperature. "3 The idea of using high-boiling oxygen-containing
compounds was set forth by K. E. Tsiolkovskiy repeatedly. He
also noted the expediency of using hydrocarbon compounds as the
fuel. He considered the use of such fuels in his work "A
Space Rocket. Experimental Preparation," of 1927. In his
work, "Reaching the Stratosphere. A Fuel for a Rocket, "4 He
presents and analysis of the influence of the quality of
fuel on the exhaust velocity of gases from the nozzle and the
flight velocity of a rocket. Here, in particular, Konstantin
Eduardovich wrote, "It is most suitable to replace oxygen with
NCu. This is a brown, chemically stable liquid, denser than

water. "^

tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 145.
Ibid . , p. 24 li .

3 Ibid .

Manuscript received Osoviakhim Central Council in 1934.

5 Tsiolkovskiy, K. E. , Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 373.

29

Konstantin Eduardovich did not limit himself to the study
of the possibilities for the use of liquid fuel components alone.
In his work, "A Space Rocket. Experimental Preparation," he
spoke of the possibility of using solid substances as fuels
and suggested, in particular, carbon powder.* Although this
type of fuel is not used in LRE at the moment, the idea of
the use of powdered products and components in various states
has been applied to some extent.

Konstantin Eduardovich was not fully satisfied by the
energy qualities of chemical fuels. He presents a number of /34
considerations concerning the possibility of using nuclear
fuel.

In 1912, he wrote, "Therefore, if it were possible to
accelerate the decomposition of radium or other radioactive
substances sufficiently, this could provide, with otherwise
equivalent conditions, sufficient velocity to a rocket that it
could reach the furthest sun (star) in ten to forty years. "2
And again, "If radium, giving up its energy a million times
more rapidly than occurs presently, could be used, inter-
planetary flights would be possible."-* Later, in 1926, the
scientist wrote, "The splitting of atoms is a source of tre-
mendous power... This energy is 400,000 times greater than
the most powerful chemical energy. "^

However, at that time it was impossible to plan on the use
of artificial radioactive isotopes and the use of fission or
synthesis reactions.

In his work "The Investigation of Space with Reaction
Devices" (1926), K. E. Tsiolkovskiy convincingly showed the
undesirability of using artificial radioactive isotopes as a
source of power. However, in this same work he wrote, "But
we cannot be sure than inexpensive, rapidly fissioning sources
of energy will not be found in time. "5

Now, when artificial radioactive isotopes are produced
easily, when spacecraft carry reaction engines which produce
energy by the decomposition of artificial isotopes, the

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 262.
2 Ibid. , p. 136.

5 Ibid ., p. 143

4 Ibid . , p. 189.

5 Ibid.

30

scientific forethought of K. E. Tsiolkovskiy on the possibility
of acceleration of the splitting of isotopes is receiving its
deserved attention.

Artificial radioactivity is the radioactivity of artifi-
cially produced atomic nuclei. Some artificial isotopes have
short half lives, which allows significant power to be produced
with these substances.

Current experimental models of radioisotope rocket
engines utilize the energy of the decomposition of artificial
radioactive isotopes, such chemical examples as polonium-210,
strontium-90, plutonium 238, etc. The possibility cannot be
excluded of the production and realization of the energy of
extremely short lived isotopes directly on-board a spacecraft.

K. E. Tsiolkovskiy stated in 1912 the idea of the possi- /35
bility of creation of electric rocket engines: "Possibly elec-
tricity might in time be used to attain a tremendous velocity
in the particles ejected from a reaction device. "1 At the
present time, electric rocket engines of various types are in
use. Modern radioisotope and electric rocket engines develop
low thrust and are designed for installation on spacecraft.

Konstantin Eduardovich studied a large group of chemical
oxidizers and fuels for LRE, noted the possibility of using
radioactive isotopes and electric power. In his works, he
laid the foundations of the science of fuels for rocket engines.

Recommendations for the Design of Combustion Chambers

During the years when K. E. Tsiolkovskiy worked on problems
of the theory of rocket- and engines, it was difficult to
imagine the design of a combustion chamber and produce any
sort of precise id«a of the processes occurring within it.
General machine building did not have a single device in any
was similar in its operating mode or magnitude of thermal and
dynamic loads to an I, RE combustion chamber. The design of this
new thermal engine had to be developed, determining the nature
and mode of its operation, analyzing the peculiarities of the
design of the individual elements and selecting the structureal
materials to be used.

From one question, K. E. Tsiolkovskiy went over to another,
then, after achieving a solution, he returned to earlier prob-
lems, continuing deeper studies, considering the results
produced earlier.

tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 36.

31

Let us see how Konstantin Eduardovich imagined the design
of a combustion chamber, which he called the "explosive" chamber.
In his work "Investigation of Space with Reaction Devices" (1903),
K. E. Tsiolkovskiy , speaking of the rocket, noted that it
"...has a great reserve of substances which, when mixed, imme-
diately form an explosive mass. These substances, regularly
and evenly exploding in the place set aside for it, flow in the /36
form of hot gases through tubes expanding toward their ends..." 1

K. E. Tsiolkovskiy described the burning of tne fuel as
follows, "In essence there is no sharp difference between the
process of explosion of a substance and simple combustion.
Actually, both amount to more or less rapid chemical combina-
tion. Combustion is slower combination, explosion is rapid
combustion. "2

K. E. Tsiolkovskiy wrote of the possibility of controlling
the motion of a rocket by changing the thrust vector as follows:
"We see a rudder serving to control the motion of the rocket."-*
This suggestion of Tsiolkovskiy was practically realized in the
form of gas rudders, as used presently to control the flight of
a number of Soviet geophysical and other rockets. K. E.
Tsiolkovskiy also suggested another means of controlling flight.
He wrote: "Finally, by rotating the end of the tube, we could
also keep our vehicle moving in the proper direction. "^ These
These methods were studied by designers. Some modern rockets
control the thrust vector by rotation of the primary combustion
chamber or with control engines as, for example, on the booster
rocket of the Vostok spacecraft.

Thermal and thermodynamic calculations, i.e., calculations
of the thermal processes of conversion of the working fluid in
the combustion chamber and in the nozzle of the reaction engine,
performed by K. E. Tsiolkovskiy, noted the necessity of cooling
the walls of the combustion chamber. As one version of cooling,
he suggested a circulating system: "...the circulation of a
metallic liquid in the air surrounding the tube is necessary for
another purpose: to maintain an even, low temperature of the
tube, i.e. , to retain its strength." 5 To assure reliable pro-
tection of the chamber, Konstantin Eduardovich recommended that
refractory insulating coverings be used: "... the inner portion

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 73.
Ibid . , p. 368.

3 Ibid., p. 74.

4 Ibid ., p. 75.

5 Ibid ., p. 79.

32

of the tube will be covered with some sort of special refractory /37
material: carbon, tungsten... Some metals are made stronger by
cooling; these are the sort of metals which must be used,
for example copper. "1 In 1911, in the work "Investigation of
Space with Reaction Devices. The Reaction Rocket of K. E.
Tsiolkovskiy," he discussed the need to cool the combustion
chamber, the "explosion tube," with liquid hydrogen and oxygen.

The scientist imagined a system of internal and external
cooling with both components as follows: "Furthermore, the
tube is continually cooled on both the outside and inside.
Actually, a continuous stream of two very cold liquids is
sprayed into the initial section of the tube: liquid oxygen
and oil cooled by the liquid oxygen. The outer walls of the
tube are cooled by the cold oil, which itself is cooled by the
liquid oxygen which surrounds it. "2

K. E. Tsiolkovskiy emphasized that iron could not be used
to make the nozzle. He stated that more refractory materials
were required, for example tungsten: "It does not seem impossible
to find materials which could withstand this temperature. Here
are a few of the melting points of materials known to me:
nickel -- 1500, iron -- 1700, indium -- 1760, paladium -- 1800,
platinum -- 2100, iridium -- 2200, osmium -- 2500, ..tungsten --
3200, while carbon does not melt even at 3500° C." 5

The recommendations of Konstantin Eduardovich for the
design of the combustion chamber and selection of materials
to assure a normal thermal operating mode of the walls are
interesting. "The explosion tube should be made of a material
which is strong (even at high temperatures) , refractory and non-
flammable; it would also be good for it to be a good heat
conductor. It seems most favorable to make the tube of two
envelopes: the first -- inner envelope -- of a less refractory
but strong, good conducting material.'"* And again, "It would
be useful to cover the steel tube with a layer of a metal which
conducts heat well, for example cuprite, aluminum and others
(for better cooling of the tube)."5 Copper-based alloys have /38
been very widely used in domestic rocket engine construction.

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 79.

2 Ibid. , p. 271.

5 Ibid ., p. 133.
4
Ibid . , p. 263.

5 Ibid. , p. 272.

33

Other works of K. E. Tsiolkovskiy are known, dedicated to
the problems of design and reliable operation of the combustion
chamber of a rocket engine, as well as the selection of materials
to assure normal thermal mode of the walls.

Many of the ideas of K. E. Tsiolkovskiy have been used in
the design of modern LRE; in particular, they almost all have
external cooling with the oxidizer or fuel, and internal cool-
ing is also used. For example, the engines of the V2A, VSV
geophysical rockets, the RD-107, RD-119 and other engines have
inner walls cooled by enrichment of the combustion products with
fuels in the layers near the walls and by the use of natural
flow-through cooling. The heat liberated from the walls is
returned to the combustion chamber. This method of cooling is
called regenerative; it was also suggested by K. E. Tsiolkovskiy.
Materials with high heat conductivity, highly refractory and with
good strength characteristics are currently used to manufacture
combustion chambers.

Thus, K. E. Tsiolkovskiy, in order to assure a reliable
thermal mode of the combustion chamber wall, suggested that
high strength, thermally stable materials be used, that the
steel wall be clad with copper, that copper be used as a
structural material, that the chamber be equipped with a heat
insulating refractory liner and that the outside be cooled by
flowing fuel components or a circulating system with liquid
metal, that the heat flux from the gases to the wall be reduced
by means of internal cooling.

Modern methods of investigation of LRE cooling systems
have also led to recommendations quite similar to those of
K. E. Tsiolkovskiy.

Let us take for example a chamber with external flowing
coolant. Let us assume that the inner surface of the wall is
heated by convection. In order to increase the permissible
temperature of the inner surface of the chamber wall, high-
strength, thermally stable materials should be used. To
decrease the wall temperature, the heat conductivity of the
wall material should be increased, which is possible if copper
or copper alloys are used, if the external cooling is inten-
sified by increasing the heat transfer coefficient from the /39
wall to the fluid. This is achieved by selecting a liquid with
optimal cooling properties, by increasing the flow rate of
cooling fluid per second, which is possible if a circulating
cooling system is used. The wall temperature can also be reduced
by decreasing the heat transfer factor from the gases to the wall
(using the principle of internal cooling). The temperature
of the gases in the layer next to the wall is decreased in this
case, also leading to a decrease in the temperature of the wall.

34

Tsiolkovskiy published the first results of thermochemical
calculations in 1903, presenting data on the thermal effect of
combustion of hydrogen and oxygen. In 1914, in the work "Inves-
tigation of Space with Rocket Devices," he spoke oi" the deter-
mination of the temperature of the combustion products consider-
ing dissociation. Consideration of dissociation allows more
precise determination of the value of the thermodynamic para-
meters, the most proper approach to analysis of structural
elements. Based on nonrelationships , he calculated the ins-an-
taneous values of the temperature of the expanding gas stream.
In 1926, his calculations were continued to the point of deter-
mination of the parameters of the gas and the efficiency of the
engine depending on the degree of expansion of the gases in the
nozzle.

Analyzing the operating conditions of a combustion chamber,
K. E. Tsiolkovskiy concluded that its weight and volume would
be low. In his work "A Space Ship," written in 1924, the com-
bustion chamber is described as follows: "Only this chamber and
its continuation -- the explosion tube, into which the products
of the explosion will flow, gradually expanding and cooling due
to the conversion of disordered thermal energy into kinetic
energy -- will experience the pressure of the gases... The tubes
and explosion chamber are very low in volume,"!

In 1926, in his work "Investigation of Space with Reac-
tion Devices," comparing possible modes of operation of chambers,
he wrote "the pressure of the explosive substances can be varied
from 5,000 atm to a desirable lower value." And again, "The
mixing may be so complete, so close, that the explosion will be
almost instantaneous or, conversely, it can be as slow as com- /40
bustion. . ."^

Studying the operation of a combustion chamber in its
interaction with the fuel feed system, the scientist comes to
the conclusion of the necessity to limit the pressure in the
chamber: "We can now indicate the required minimum pressure."
And the conclusion, "In any case, we can limit ourselves to 100
atm." 3

In the work just cited, Tsiolkovskiy describes the
process of conversion of heat, liberated on combustion of the

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 164.

2 Ibid. , p. 201.

3 Ibid . , p. 202.

35

fuel, to kinetic energy in the gas stream and presents the
results of his calculations.

In this same work, the peculiarities of the design of some
parts of the rocket and its engine are noted. As concerns the
nozzle, the following statements were made: "However, the
greater its angle, the greater the loss of energy, since the
motion of the gases is deflected to the side. Still, with an
angle of 10°, the losses are almost unnoticeable." 1 However,
in 1927, he recommends that the optimal value of the angle of
expansion of the nozzle be determined by experimentation.

In 192 7, in the work "A Space Rocket. Experimental
Preparation," the method of injection of fuel to the chamber
and its preparation for combustion is described as follows:
"...gratings with slanted holes for better mixing of the
hydrocarbon with the oxygen mixture. The beginning of the
explosion tube is divided by a channel. Along one half flows
the oxygen mixture, along the other half -- the hydrocarbon. "2

Development of Feed Systems

The plan of the system which feeds fuel to the chamber
of the rocket engine was developed by K. E. Tsiolkovskiy in
1903. In his work "The Investigation of Space with Reaction
Devices," K. E. Tsiolkovskiy suggested and himself described
a system of fuel feed with unloaded tanks, i.e., tanks in which
the fuel is stored under low pressure.

At first, considering that there would be very high pres-
sures in the combustion chamber, K. E. Tsiolkovskiy concluded
that it would be necessary to use a pulsed fuel flow mode.
In 1914, he wrote, "Ordinary types of pumping should not be
used. It would be simplest of all to place a certain charge /41
in the tube and allow it to burn and fly out. Then, when the
pressure in the tube had dropped, another charge would be
injected, etc."-* Here also he stated the idea of the possi-
bility of using a gas-jet ejector: "Theie should be a oranch
at the very mouth of the tube, through which the gases would
be returned once more to the mouth and, due to their velocity,
entrain and force the explosive material in a continuous stream

T

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 247.

Ibid . , p. 263.

3 Ibid. , p. 147.

36

into the very mouth of the explosion tube."

Analysis of the weight qualities of the feed system led the
scientist to the idea of the need to reduce the pressure in
the chamber, to select its optimal value. "At high pressure,"
Konstantin Eduardcvich wrote in 1926, "the use of the energy
is great, but impossibly great work is required to force the
masses into the explosion tube. Therefore, the maximum pressure
in the tube should be reduced as greatly as possible, without
losing efficiency. "^

The idea of pulsating feed was formulated by him as
follows: "...it could be made so that the pressure at the
beginning of the tube varied periodically, for example, from
200 atm to and from back to 200 atm. The variation would
occur in waves. "3

K. E. Tsiolkovskiy believed that the walls of the tank
should also form the shell of the rocket. "The main shell of
the rocket," Konstantin Eduardovich wrote in 1926, "should
withstand without danger a .pressure of at least 0.2 atm, if
filled with liquid oxygen." "Then, to store them (i.e., the
fuel components) ordinary tanks or even the rocket itself could
be used. "^ These rockets have been widely used. Thus, K. E.
Tsiolkovskiy suggested to so-called "load-bearing tanks,"
i.e., fuel tanks, the side surface of which is at the saire
time the outer shell of the fuel section, receiving external
longitudinal forces and Landing moments acting on the section.
The use of load-bearing tanks allows the mass of a rocket to be
greatly reduced in many cases.

In 1927, the scientist suggested that a pumping unit be /42
installed between the tanks and the chamber: "... -- two pumps,
driven by a common motor. The first pumps the oxygen compounds 6
into the explosion tube, the second -- the hydrogen compounds."
Calculations have shown that the fuel consumption for the
pump drive would be insignificant: "...the motor would use
several hundred times less fuel than the explosion tube. "7

Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Tress, Moscow, 1954, pp. 147-148.

'Ibid. ,

P-

201.

3 Ibid. ,

P.

202.

4 Ibid.,

P-

243.

5 Ibid.,

P-

246.

6 Ibid.,

P.

261.

7 Ibid.,

P-

265.

37

As a result of his studies of the peculiarities and
conditions of operation of individual units and systems of
the rocket, in 1927, in his work "A Space Rocket. Experimental
Preparation," K. E. Tsiolkovskiy presented a description of
the launch and operation of the motor in flight.

Konstantin Eduardovich Tsiolkovskiy was an outstanding
researcher, whose scientific activity was unusually broad.
He made many discoveries in the area of rocket dynamics,
aerodynamics, the theory of aviation, the theory of inter-
planetary voyages, the theory of engines, etc. The work on
rockets performed by K. E. Tsiolkovskiy, did not amount to
a completed technical plan. We can gain an idea of his
design only by looking at his calculations and descriptions.

The most important thing in the works of Tsiolkovskiy was
the proof of the possibility of constructing a large ^pace
rocket with an LRE, confirmed by calculations. K. E.
Tsiolkovskiy pointed the way into space. Soviet and foreign
scientists recognize the priority of Tsiolkovskiy as the founder
of theoretical astronautics. The name of Tsiolkovskiy has been
given to a crater on the far side of the moon. Konstantin
Eduardovich Tsiolkovskiy is rerognized as the 'iead of a new
trend in science and technology -- astronautics and rocket
building.

In connection with the launch of th ••world's first art .
ficial satellites, a gold "Tsiolkov v > "al" was founded,
awarded by the Academy of Sciences USSR to^ outstanding work
in the area of interplanetary voyages. In 1958, the first
medal was awarded to the Chief Designer for Rockets and Space-
craft, Academician Sergey Pavlovich Korolev, while the second
medal was awarded to the Chief Designer of Rocket Engines.

1.4. One of the Pioneers of Rocket Technology, /43

Yu. V. Kondratyuk

The attempts of historians to write a detailed biography
of Yu. V. Kondratyuk, a talented and gifted man, a remarkable
scientist, mechanic and inventor, and vhe author of the well-
known works "Mastern of Interplanetary Space" and "Th^se Who
Will Read in Order to Build," have not yet been fully success-
ful. Too few documents have been retained in the archives.

Yu. Vasil'yevich Kondratyuk was born in 1897 in the
Ukraine, in the city of Poltava. The unfortunate conditions of
his life did not allow him to complete his education: Yu. V.
Kondratyuk worked as a hired laborer, chopped firewood, and
worked as a lubricator and mechanic at mills. He studied

40 PRECEDING PAGES;

mathematics, physics and chemistry independently. In his youth,
Yu. V. Kondratyuk became interested in the theory oi" inter-
planetary voyages. In 1918, looking over some old magazines,
he came upon one of the articles of K. H. Tsiolkovskiy on his
stratoplane, while he read the other works of Tsiolkovskiy,
particularly his article "The Investigation of Space with
Reaction Devices" (which was written in 1911) only in 1925.

The
tary fli

bas
ghts

ic problems and physical princi

were set forth bv Yu. V. Kondr

"Those' Who Will Read

.__ fie work on this man

1916 and completed i
of Yuriy Vasil'yevic
in 1964. Based on h
his familiarity with
of K. li. Tsiolkovski
reworked this articl
He performed careful
of rocket and space
new solutions and pe
tions.

Yuriy Yasil'yevich
Kondratyuk

pxes of interplane-
atyuk in his work

in Order to " .ild,"
uscript was ! gun in
n 1919. This work

was first published
is own studies and

some of the works
y, Yu. V. Kondratyuk
e several times.

studies of a number
problems, presents 1
rformed many calcula-

Yu. V. Kondratyuk produced the basic
equation of motion of a rocket by his
own original method independently of
K. E. Tsiolkovskiy, with the works of
whom he became familiar only later.

In 1925, the manuscript of
"Mastery of Interplanetary Space" was sent
to Professor V. P. Vctchinkin (1885-1950)
who made a positive response.

/44

Encouraged by
studies and in 1929
edited by Professor
book, Vladimir Petr
"The book by Yu. V.
most complete study
Russian and foreign
studies were pcrfor
book discusses with
terei in other work
a number of new pro
other authors. The

success, Yu.

V. Kondratyu

published the manuscript

V. P. Vctchinkin. In th

ovich wrote

the following

Kondratyuk

which you hoi

of interplanetary voyage

1 iterature

up to the pre

med by the author quite i

exhaustive

completeness

s and, furthermore, prese

hi ems of primary import an

se problems

include :

k continued his

in Novos ibirsk ,
e foreword to the

on 4 December 192 7:
d is doubtless the
s wri . en in the
sent time. All the
ndependent ly . This
all problems encoun-
nts the solution to
ce , not mentioned by

''The suggestion that solid fuels (lithium, boron, aluminum,
magnesium, silicon) be used in addition to gaseous fuels, both
to increase the heat of combustion, and in order to use

41

combustible tanks which, after they are emptied of liquid fuel,
are themselves processed and sent to the furnace. This same
suggestion was made by engineer F. A. Tsander in a report at the
Theoretical Section of the Moscow Society of Astronomy Enthusi-
asts in December of 1923, but this suggestion was included in the
manuscript of Yu. V. Kondratyuk before the report of Tsander.

"He first presented a formula considering the influence of /4S
the weight of the tanks for fuel and oxygen (proportional
passive to use the terminology of the author) on the total
weight of the rocket, and proved that a rocket which did not
jettison or burn its tanks during flight could not escape the
bonds of the Earth's gravity.

"He also first made the suggestion to make a rocket with
wings and fly it in the air like an airplane. This suggestion
has not appeared in the foreign literature at all (it being
rather suggested that parachutes be used to return the rocket
to the Earth), while Russian works have seen this suggestion,
stated by F. A. Tsander at the same meeting and later printed
by K. E. Tsiolkovskiy, but only after it appeared in the manu-
script of the author. However, the studies of Yu. V. Kondratyuk
go further, since he not only indicates the need for the use of
wings, but also presents a rather detailed study as to the
accelerations at which wings will be useful, the trajectory
angles of the rocket to the horizon for the use of wings, and
gives the most favorable force of reaction of the rocket during
flight in the air; it is found to be on the same order as the
initial weight of the rocket.

"Generally, the dynamics of the takeoff of the rocket repre-
~ent the most difficult portion of the problem, and Yu. V.
Kondratyuk has solved it more completely than any other author.

"Here also is presented a study of the heating of the foreward
portion of the rocket by the air considering both adiabatic com-
pression of the air, and radiation of the surface of the rocket
and of the heated air itself. This problem was also studied by
no one.

"All numbers were given by Yu. V. Kondratyuk, although rather
roughly (which he himself mentions in the foreword) , but always
with his error in the direction unfavorable to the designer.

"This book can serve as a desk reference book for all those
involved in problems of rocket flight."

In the early 1930's, Yu. V. Kondratyuk, without interrupting
his work on rocket technology, began studying high power wing
installations. Supported by the People's Committee for Heavy
Industry and TsAGI , he headed the planning of a wind power plant

42

at the Ukrainian Scientific search Institute for Industrial
Power Engineering (Khar'kov). The plan was approved by the /46
Academy of Sciences USSR. To bring it to life, "Teploenergostroy"
Trust (Moscow) was directed to construct a wind power plant with
a capacity of 12,000 kw in the Crimea, under the leadership of
G. K. Ordzhonikidze. In 1938, Kondratyuk was named Chief of the
Technical Section of "Teploenergostroy" Trust, then Chief of the
Planning Section of the "Planning and Experimental Office for
Electric Power Plants." In later years, Yu. V. Kondratyuk
studied the construction of powerful wind power plants, as
before without interrupting his studies on interplanetary voy-
ages.

In 1947, the book of Yu. V. Kondratyuk "Mastery of Inter-
planetary Space" was reissued. Some of the conclusions of
Yu. V. Kondratyuk agreed to some extent with those made by
K. E. Tsiolkovskiy. However, Kondratyuk* s book contained a
great deal of new and original material. The young scientist
was the first to develop: the energetically most favorable
trajectories for space flights, problems of the theory of multi-
stage rockets, designs for intermediate filling stations on the
artificial satellites of the planets, particularly the moon,
the conditions for economical landing of rockets on the Earth
using atmospheric braking, approximate methods of calculation
of the heating of a rocket as it moves through the atmosphere.
He recommended that a number of types of oxidizers be used,
particularly ozone, while recommending metals, metalloids and
their hydrogen compounds such as boron hydrides as fuels.
After suggesting that winged rockets be used, Yu. V. Kondratyuk
indicated the areas of their application and performed studies
on the selection of the most suitable aerodynamic character-
istics.

Our attention is drawn to the idea of Yu. V. Kondratyuk

of the utilization of solar energy: solar heat is converted by

electricity, then thrust is created by expulsion of elementary
particles.

On 7 June 1941, Yu. V. Kondratyuk enlisted in the People's
Volunteer Core. Leaving for the front, he gave his friends a
suitcase and portfolio with his manuscripts for safekeeping.

Yu. V. Kondratyuk was a soldier in the Communications
Company of the 2nd Regiment of the People's Militia Division of
the Kiev Region of Moscow. He took part in battles with the
German Fascist invaders and died at the front in 1942.

The name of Kondratyuk has been given to a crater on the far
side of the moon.

43

The Works of Yu. V. Kondratyuk on 'Rocket Engines

Like K. E. Tsiolkovskiy, Yu. V. Kondratyuk came to the con- /47
elusion that rockets should be driven by LRE and should have
more than one stage. In his book "Mastery of Interplanetary
Space," he wrote that the reserve of energy to be used to impart
speed to a flight vehicle can be carried on board in quite
varied forms, but that only the chemical energy of the compounds
of certain substances would be sufficient to allow flight in
practice. Planning on the use of a multistage rocket, Yu. V.
Kondratyuk objectively studied its design, flight conditions and
provided a foundation for the selection of fuels, suggesting an
arrangement of the combustion chamber and nozzle and indicating
the need to use a turbine pump unit.

Suggestions for LRE Fuels

In selecting a fuel, Yu. V. Kondratyuk first turned his
attention to its efficiency. Furthermore, he believed it neces-
sary to consider all of the variety of properties of a fuel, as
well as the design of the rocket and the specifics of the con-
ditions of its use. If the rocket is composite, i.e., a multi-
stage rocket, a greater quantity of fuel is required for the
operation of the first stages than for the latter stages.

In selection of fuel, Yu. V. Kondratyuk noted, one must
also turn attention to its cost. According to Kondratyuk, the
use of the least expensive fuels would be expedient for the
first stages of the rocket, with more efficient and costly
fuels to be used in later stages. Kondratyuk suggested a formula
considering the cost of the fuel, its mass and thermal efficiency
to estimate the "cost of reaction."

He considered liquid air, oxygen and ozone to be the most
effective oxidizers, with petroleum products, liquid acetylene,
methane-based fuel, hydrogen and its compounds, as well as
products containing aluminum, magnesium, silicon and boron to
be the best fuels. The last fuel could be used as an amorphous
powder, pulverized in the combustion chamber by a stream of
hydrogen or methane or added to oil before it was fed into the
combustion chamber. Yu. V. Kondratyuk suggested boron hydride /48
as a fuel.

Yu. V. Kondratyuk studied several groups of fuels: first,
liquid air-petroleum or liquid oxygen-petroleum, then liquid
acetylene, then liquid hydrogen. He then studied the possibility
of using several metalloids and metals. Kondratyuk calculated
the thermal effect of a variety of fuels, as well as their
combustion product exhaust velocity from the nozzle and other

44

parameters. Kondratyuk stated his doubts concerning the
expediency of using liquid hydrogen, due to its low density.

Recommendations for the Design of the Combustion Chamber

Kondratyuk turned a good deal of attention to problems of
the organization of combustion in LRE combustion chambers.
As early as 1918-1919, studying the combustion of hydrogen and
oxygen, he wrote that the combustion of the fuel could be
organized by three methods -- either a prepared mixture could
be ignited, or the gases need not be mixed until the actual
moment of ignition, or they would be only partially mixed,
with the best method to be determined by experience.

To assure completeness of combustion, Kondratyuk suggested
a checkerboard placement of the fuel component sprayers in the
spray head of the combustion chamber. He also suggested a
"stratified" version. In this case, the sprayers would be
placed along walls of the chamber in belts, alternating with
each other. In his thermodynamic calculations, Yu. V. Kondra-
tyuk considered the dissociation of the combustion prcducts;
he believed that the process in the chamber is nearly isothermal,
while adiabatic expansion of the gases occurs in the nozzle.

According to Yu. V. Kondratyuk, the "combustion chamber"
and "expulsion tube," i.e., the nozzle, should be made in one
piece, and he believed that the surfaces exposed to gases at
temperatures higher than those which could be withstood by the
refractory material applied to the walls of the chamber, should
be made of metal -- copper or one of the refractory metals
(chromium or vanadium) , and that the walls should be intensively /49
cooled on the inside by liquid gases fed into the combustion
chamber.

Studying the design of the nozzle, Yu. V. Kondratyuk wrote
that the most favorable nozzle shape approximates a paraboloid
of rotation, but not a quadratic paraboloid, rather one of
higher order; toward the nozzle exit plane, it should be con-
verted to a cylinder. The flow of combustion products leaving
the nozzle would then by one dimensional, not diverging, in
order to achieve the greatest possible efficiency and, conse-
quently, thrust. Yu. V. Kondratyuk pointed out that the finish
of the inner surface of the nozzle should be such as to provide
the minimum loss due to friction of the combustion products
against the wall, and that the profiling of the nozzle and cal-
culation of cross sections should be based on the condition of
conservation of constant flow rate (continuity of flow) of the
combustion products.

45

Studying the influence of external conditions on the opera-
tion of an LRE, Yu. V. Kondratyuk recommended that, in order
to avoid decreasing the efficiency as the engine operated at
low altitudes, the cross-sectional area be decreased, i.e.,
the nozzles should be equipped with an additional device in the
form of a constricting cone at the exit plane of the nozzle, to
be used in the lower layers of the atmosphere then jettisoned as
the altitude increased. As another version of thrust regulation
with altitude, he suggested that the combustion chamber be
equipped with a dual nozzle -- the first to provide optimal
parameters for operation at low altitude, the second to be used
at high altitude and to begin operation after the first nozzle
is jettisoned.

Many of the suggestions of Yu. V. Kondratyuk concerning the
design of combustion chambers have been realized in practice.

Development of Feed Systems

In his work "Those Who Read in Order to Build," he noted
that a rocket engine with a chemical source of energy should
consist of vessels, tanks, the combustion chamber tube and
devices to feed the fuel components from the tanks to the combus-
tion chamber of the rocket engine. Yu. V. Kondratyuk suggested
that pump systems be used to feed both single-component and two-
component fuels. At first, he planned on the use of piston
pumps. Later, he wrote that pumps could also be made pistonless. /50
Kondratyuk' s pumps were to be single- cycle pumps, and each com-
ponent was to have its own pump. The liquefied gases fed by the
pumps were to be used primarily for combustion, partially to
pressurize the tanks carrying the fuel components.

To assure normal operation of the engine, Kondratyuk
suggested a fuel- feed regulation system. The sensing element
used wr.s a device similar to an aneroid barometer, reacting to
the pressure difference inside and outside the tanks.

The actuating element regulating the tank pressurizing
system is a choke valve installed before the inlet for gas
products into the tank. Yu. V. Kondratyuk also suggested that
a mixture quality regulator be installed before the inlet to
the combustion chamber, although the introduction of the regula-
tion system complicates the design of the engine.

Furthermore, Yu. V. Kondratyuk turned particular attention
to the need for preliminary development and experimental check-
ing of the elements of the engine. Thus, Yu. V. Kondratyuk
suggested methods of assuring the required operating mode of the
engine by its adjustment to a fixed mode and regulation during
operation, now used in practice.

46

K»- " . • •• mmc « ' t ha u • rial - mi is t ion

or turbine >•*"•' ' ►•the main om oni its of the fu< ., -<ed

as pum Irivc; the use of oxyh .'•• •,,--•.. . - : lied.

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the ,•• : ; - fl jght,* fie j I so

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l.S, hi icntist i ml Inveni >i F. A, Tsander

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«... piont : '■ ' ■ i ket on: i ■■ ' ... tnd an enth ;iast for

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as nainci an '" rij r > i = i

fr idr i kh rtui ' I •■ ram ter-

■'.•-' ; techno] . . •-*,• is
student yc irs In 1 < I; •. .' i c \ rote ,

"when I was 2i, I *eg«n ' . keep a
spec ' not '<•-. ok . or the des igri >f
eshtj t It hong] l km el

\ut - "' oj t i ;'••- • :' ," un ..*"•; i hi ves ,

little, under the influence of my calculations I had already
begun to hope for the possibility of flights in space."!
In 1909, F. A. Tsander was an initiator in the creation in
the Institute of the "Second Riga Student's Society for Air
Travel and Flight Technology," and in that same year he con-
structed a glider with his comrades.

F. A. Tsander advanced from the idea of reaching great
altitudes by means of an airplane and propeller motor to the
idea of the possibility of interplanetary space flight with a
rocket engine. In order to attempt to realize his plans, F. A.
Tsander began work at Moscow Aviation Plant No. 4, "Motor"
in February of 1919 as the head of the technical bureau. Late
in 1921, F. A. Tsander presented to the Moscow Governor's
Conference of Inventors a plan for an engine for an interplane-
tary airplane-spaceship. From June 1922 through July 1923,
Tsander, on temporary leave from the plant, worked at home.
He constantly felt the support of the workers, who gave him
significant material assistance. F. A. Tsander valued this
relationship, and reported to the workers. For example, in
April of 1923, at a plantwide meeting of workers of "Motor"
Plant, he reported his hope to be able to give his plan to the
plant for construction.

In 1924, in the journal "Tekhnika i Zhizn'", No. 13, the
first printed work of F. A. Tsander appeared -- the article
"Flights to Other Planets." In this article, he presented his
basic idea -- the combination of a rocket with an airplane,
with subsequent burning of the metal parts of the airplane.
In 1924, F. A. Tsander wrote the article "Description of the
Interplanetary Spaceship-Airplane System of Tsander," which was
sent to the Committee for Inventions of the All-Russian Council
of the National Economy 8 July 1924. This article was pub-
lished in the collection " Rake t nay a Tekhnika" [Rocket Technology]
in 1937. F. A. Tsander believed that an airplane with a
piston engine could achieve an altitude of about 28 km and a
speed of 350 to 450 m/sec. After this, the ship is switched
from the piston engine to a rocket engine. No longer needed,
the airplane is pulled piece by piece (wings, tail, chassis,
piston engine, etc.) into a special device, where it is
melted and used as an additive to the liquid fuel. At the end
of the acceleration run, at an altitude of 85 km, only the
rocket with small rudders and wings as needed for a gliding
descent would be left.

Attempting to get his works published, F. A. Tsander sent
some of them, particularly "The Utility of Acceleration of the
Flight of a Rocket at Moments when the Flight Velocity of the

Autobiography of F. A. Tsander , family archives.

48

Rocket is Great," "Flights to Other Planets," and "Calculation of
the Flight of an Interplanetary Ship in the Atmosphere" to the
Scientific Council of the People's Commissarirt for Education,
RSFSR, Professor V. P. Vetchinkin. In his review of 8 February
1927, which was sent to the Scientific Department of the Main
Administration for Science, V. P. Vetchinkin, noting the value
of the ideas and works of F. A. Tsander, considered it quite
necessary to help F. A. Tsander to prepare and publish his works,
some chapters of which had already been presented to the Admin-
istration for Science, as rapidly as possible. Actually, due to
the fact that the publication of scientific works was not given
its proper significance, in those years we lost priority even in
those cases when it factually and undisputably belonged to our
country. For example, in 1925 the work of engineer Gochman was
published abroad, in which he suggested flight on wings and
gliding descent. The ideas developed by Yu. V. Kondratyuk and
F. A. Tsander were published in this work.

A few days after he received a reply from V. P. Vetchinkin,
F. A. Tsander sent the Scientific Division of the Main Adminis-
tration for Science an announcement, in which he requested to
be allowed to work at the Central Institute for Aerodynamics and
Hydrodynamics (TsAGI) or the Aviation Trust exclusively in the
ar«a of interplanetary voyages, and permission to prepare for
printing a book on interplanetary voyages. In July of 1927, the
Administration sent a message that the request of F. A. Tsander
was not approved.

In order to make his employment more closely related to the
development of space flight, F. A. Tsander had earlier, in
October of 1926, transferred to work at Aviation Plant No. 4 in
the Central Design Bureau of the Aviation Trust as a Senior
Engineer. F. A. Tsander reported the results of his works on
problems of the theory of rocket engines in a report "Preliminary
Work on the Construction of a Reaction Apparatus," which he read
on 30 November 1928 at the ISth Session ot the Commission on
Scientific Air Travel of the Moscow Aerological Observatory. In
1929-1930, F. A. Tsander, at the request of the Aviation Trust,
prepared a report on the basis of his studies entitled "Problems
of Superaviation and Immediate Problems on the Preparation for
Interplanetary Voyages" for the Fifth International Congress on
Air Travel, which was planned for September of 1930 at the
Hague. After a number of revisions of the material which formed /S4
the basis of this report, F. A. Tsander prepared his book "The
Problem of Flight with Reaction Apparatus," which was published
in 1932.

In December of 1930, F. A. Tsander began to work at the
Central Institute of Aviation Motor Building (TsIAM) , where in
1931 he began the construction of the OR- 1 aviation reaction
engine, followed by the OR- 2 LRE. The OR-1 engine operated on

49

compressed air (supplied from cylinders or by a compressor) and
gasoline; the OR- 2 LRE at first (in 1933) was tested with liquid
oxygen and gasoline. We can trace the sequence of work in
this direction in the diary of F. A. Tsander.

On 15 September 1931 in his diary he comments on his work
on the airplane with the reaction engine; on 1 October, he
discussed with Yu. A. Pobedonostsev "installation of the
reaction engine on the airplane," and on 2 October he wrote
in his diary "about the oil-oxygen rocket for the airplane"*;
on 7 October he noted the conduct of the 32nd test of the OR-1,
held in the presence of S. P. Korolev and other specialists,
while on 19 October we see the first mention of the OR- 2 engine;
on 18 November 1931, F. A. Tsander concluded a Socialist contract
with the Aviation Technology Bureau of the Scientific Research
Sector of the Osoaviakhim CC for the planninf of a reaction
engine, including its installation on an aircraft. 2

F. A. Tsander agreed to plan and produce working drawings
for the OR- 2 reaction engine for the RP-1 jet aircraft in the
following periods of time: combustion chamber with nozzle,
tanks for fuel with safety valve, tank for gasoline -- by 25
November 1931; compensator for cooling of nozzle and heating of
oxygen -- by 3 December 1931.' The times for completion of
calculation of the temperature in the combustion chamber,
exhaust velocity and axial pressure of the jet in the nozzle
at various pressures in space, weights of the parts, flight
duration of the RP-1 reaction aircraft with various oxygen
contents, calculation of the heating and cooling system, approx-
imate calculation of the temperature of the walls of the com-
bustion chamber -- all corresponded to the time for completion
of the drawings.

Manufacture and testing of the nozzle and combustion
chamber were planned for 2 December 1931; the fuel tanks for
liquid oxygen and gasoline -- by 1 January 1932; installation
of the OR- 2 on the RP-1 aircraft and flight testing were planned
for the end of 1932.

An addendum to the agreement noted that if the planned
improved nozzle included a direct and reverse cone, calculations
and drawings were to be completed by 15 January 1932. This
extiemely short period of time for completion of a complex prob-
lem of large volume, including theoretical study, calculation,,

F. A. Tsander family archives.
2 Archives of Academy of Sciences USSR, F 573, d 269, p. 10.

50

planning, manufacture and testing, characterized both the enthu-
siasm and optimism of the contractor, and the underestimation
of the difficulties which would arise in completing the obliga-
tions undertaken. This was a result of the lack of experience
in development of LRE, as well as the mismatch between the com-
plex technology of manufacture of the engine and the relatively
low production capacities which could be found at the time.

In 1931, Osoaviakhim allocated F. A. Tsander 1000 rubles
for the study of reaction motion, on 25 February 1932 another
13,000 rubles tor the testing of rocket aircraft, followed by
80,000 rubles in March of the same year.

It soon became clear that the preparation of detailed work-
ing drawings and the completion of full calculations of a reac-
tion engine with a complex control system were simply too much
for F. A. Tsander alone. The need thus arose to concentrate
the efforts of scientists and engineers working in the area of
reaction technology.

F. A. Tsander also believed that for practical development
of rocket technology, the largest possible number of engineering
and technical workers, particularly talented young people, would
be needed. We will discuss in detail the creation and develop-
ment of the creative team headed by F. A. Tsander.

In 1932, Tsander' s work "The Problem of Flight Using
Reaction Apparatus" was published as a separate book. Here,
in addition to the presentation in the theory of the flight
of rockets and airplanes, we find methods of selection of fuel
and design of various reaction engines.

In 1932, F. A. Tsander began working on the creation of
his first LRE, called the OR-2. The engine was tested for the
first time in 1933, burning liquid oxygen and gasoline. Later,
at the RNII, the design of the engine was significantly changed
in order to improve its efficiency, and in version 02 it used
liquid oxygen and highly concentrated ethyl alcohol.

Liquid oxygen (like liquid fluorine, liquid hydrogen) is
a cryogenic rocket fuel component. It is a compressed gas,
cooled to a low (cryogenic) temperature. Cryogenic fuel must
be used when this is justified by the increased specific impulse
which it provides, for example in the boosters of spacecraft.
Cryogenic fuel is not suitable for long-term storage, due to the
evaporation losses.

At GIRD from the very first days of organization of this
group and formation of the first team, Fridrikh Arturovich
worked on other problems as well. He turned his attention to
the construction of a rocket, later called the GIRD-X. Work was

51

begun on this rocket in January of 1933, and it was launched on /56
25 November of the same year, but without Tsander. The creation
of the GIRD-X rocket was preceded by many calculations, rough
plans and experiments, performed and conducted by Fridrikh
Arturovich.

F. A. Tsander spent most of his day in calculation, while
also working on production, helping the mechanics who encoun-
tered slight difficulties in the manufacture and tuning of
apparatus new for the time. The engineers and designers, with
F. A. Tsander as their chief, worked together in a small room.
They worked morning, noon and night, whenever needed, as long
as they had strength.

In addition to his plan tasks, F. A. Tsander calculated and
thought about the design of individual units of the rocket, which
he called a "spaceship."

"Forward to Mars!" "Faster to Mars" -- these words symbol-
ized the goal of his life. He frequently shared his thoughts
with his coworkers on the first team, tossing off drawings of
individual parts of the spacecraft. Gradually, the form of the
future rocket developed, the rocket which Tsander dreamed would
fly to our neighbor planet.

During the last months of his life, Fridrikh Arturovich
worked especially hard. As a result of overfatigue, systematic
lack of sleep, poor and irregular feeding, F. A. Tsander began
to lose his ability to work. On the insistance of his coworkers,
Fridrikh Arturovich traveled to Kislovodsk for treatment. On
the way, he contracted typhus and died on 28 March 1933.

In 1947 a collection of the works of F. A. Tsander was
published under the title "The Problem of Flight Using Rocket
Apparatus." The collection was reissued in 1961, expanded to
include many works published for the first time.

The archives contain many more unpublished works of the
scientist. Most of the remaining manuscripts require long and
tedius work to translate Tsander 1 s shorthand to ordinary text.
The difficulty of decoding is explained by the fact that F. A.
Tsander used a long- forgotten type of shorthand, which he him-
self altered somewhat, writing on specific problems of the
theory of engines and rockets in German. Thus, the work with
the manuscripts of F. A. Tsander requires specialists familiar
with his system of writing, fluent in German and familiar with
rocket technology.

The first deciphering of the works, of F. A. Tsander was
performed at RNII. In 1934, a group of stenographers under the
direction of Ye. K. Moshkin decoded several notebooks filled

52

with writings recorded by F. A. Tsander during the early
period of his activity. Up to I960, the study of the heritage
of F. A. Tsander was conducted with no strict plan, unsystem-
atically. The appearance of earlier unpublished ideas of F. A.
Tsander in print and the organization of a number of meetings,
jubilees and conferences dedicated to the memory of F. A.
Tsander are largely due to the efforts of Astra Fridrikhovna /57
Tsander, the daughter of the scientist, who also prepared i
collection of the works of F. A. Tsander "From his Scientific
Heritage" for printing (Nauka Press, 196/ J. The documents in
this collection, from the archives of Tsander, are interesting
in many respects. In particular, it is noted here that Tsander
began planning the OR-1 engine in 1928. A method of the cal-
culation of "Hydrogen-Oxygen Rockets" is presented (April, 19?.?),
in which the thermodynamic calculation of LRE is accurately
explained.

Since 1965, the deciphering and study of the works of
F. A. Tsander have been included in the plan of the Institute of
the History of Natural Science and Technology of the Academy of
Sciences, USSR.

In May of 1970, the first "readings" dedicated to the study
and realization of the scientific heritage of F. A. Tsander were
held in Riga, and adopted a resolution to hold "Tsander readings"
systematically. The second "readings" were held in May of 1972
in Leningrad.

The name of Tsander has been given to a crater on the far
side of the moon.

The Works of F. A. Tsander on Rocket Engines

F. A. Tsander, a great scientist in the area of the devel-
opment of a broad range of problems on the theory of space
flight, dedicated a significant portion of his scientific and
technical activity to theoretical studies of the possibility of
constructing highly efficient reaction flight vehicles, as well
as theoretical and practical work on the mastery of liquid-
fueled rocket engines during the initial period of their devel-
opment in the USSR.

Many of the theoretical and experimental works of F. A.
Tsander are dedicated to the finding of means for achievement
of his basic idea, that the combustion of the metal parts of
the rocket along with the liquid fuel after the parts were no
longer needed could increase the exhaust velocity of the combus-
tion products and also increase the ratio of the mass of fuel
burned during the process of a flight to the final mass of the
rocket.

S3

This idea attracted F. A. Tsander as early as the 1920 's,
but was most completely presented by him in his work "The
Problem of Flight Using Reaction Apparatus." In this work he
presents a description of two flight vehicles: an airplane wit.
a rocket engine, the wings of the airplane an some other parts
being drawn into the vehicle and melted to be used as fuel, and
rockets surrounded by a set of containers filled wi+h fuel com-
ponents, with the containers drawn into the central rocket after
their fuel content was exhausted, then melted and used as fuel.

m

F. A. Tsander believed that only the designs which he sug-
gested could achieve interplanetary speeds.

A

i
i i

i

i

'mi
ii 1 1,
ii 1 1
it ;j
irH
' > .'

1 1

i i

'i 1
/ ' i
i ' *
/ i >

3

\S?

m

F. A. Tsander' s Plan for the Interplanetary

Spacecraft

Floats

Liquid

Metal

Tube

takeoff

Plan of Central Rocket Surrounded by Side
Rockets and Fuel Tanks After F. A. Tsander

54

His total confidence in the correctness of the scientific
and technical direction he had selected also determined the
nature of his theoretical and practical developments. He
turned his attention to theoretical study of possible means of
increasing the sepcific impulse of his engine and the efficiency
of its individual unit; theoretical study and experimental
development of possible application of metals as additives to
fuel; and theoretical study and experimental development of LRE.

In the early 30*s, the level of technology and available
structural materials did not allow a rocket with a high ratio of
launch weight to final weight to be constructed (for example,
in the first Soviet rocket with LRE, the GIRD-X, this ratio was
approximately 1.4), so that the idea of F. A. Tsander was pro-
mising, but was found to be practically impossible.

In the best modern rockets, thanks to the use of the latest
structural materials, optimal design of all rocket units and
operation in the most suitable modes, very high ratios of
launch weight to final weight have been achieved.

Investigation of Fuels

F. A. Tsander was a proponent of the use of fuels with low-
boiling oxidizers. He based this opinion on the fact that
this type of fuel has exceptionally great capabilities as con-
cerns further increases in specific impulse. As an oxidizer,
he believed it desirable to use liquid oxygen, with liquid
hydrogen, gasoline or alcohol as the fuel. Gasoline, in par-
ticular, drew Tsander' s attention not only by its high heat
content, but also che possibility of its use in the aviation
and rocket engines which he planned for interplanetary flight
vehicles.

As we have stated, F. A. Tsander performed investigations
on the possibility and expediency of using metals as additives
to liquid fuel. As we know, when some metals burn, more heat is /60
liberated than when liquid fuels are burned, even such liquids
as gasoline; therefore, the addition of metal to a liquid fuel
under certain conditions might cause an increase in the specific
impulse.

For example, the heating ability of coi/ustion products in
oxygen, per kg of fuel, according to F. A. Tsander, are as
follows: for gasoline -- 2350 kcal, for aluminum -- 3730 kcal,
for lithium as high as 4710 kcal.

However, at the temperatures characteristic for LRE, solid
oxide particles are generally formed. It is therefore impossible
to calculate the exhaust velocity, thrust and specific impulse

55

by the formulas designed on the assumption of gas flow alone.
F. A. Tsander studied the conditions of motion of products con-
sisting of a mixture of gases and solid oxides. For example,
in his article, "The Use of Metal Fuels in Rocket Engines," 1 he
presents an approximate determination of the reaction force
produced by an engine expelling particles from its nozzle at two
significantly different velocities. "It is possible," F. A.
Tsander wrote, "to burn metal with liquid fuels in proportions
such that no decrease in thrust is observed. "2

In order to check his calculations and the practical
possibility of utilizing the burning of metal in the chambers
of reaction engines, F. A. Tsander first performed a number of
simple laboratory experiments on the ignition and combustion of
metals. Then, the combustion of metals was studicl using the
OR-1. Later, the program of experiments was expanded.

F. A. Tsander suggested that the metal fuel be made of
those parts and units which had performed their functions and
were no longer needed for flight or landing of the airplane-
rocket or central rocket with many side rockets and liquid fuel
and oxygen tanks which he designed.

For this reason, F. A. Tsander attempted to determine the
possibility of processing individual structural elements into
powdered or melted metal (magnesium, aluminum) and developed
plans for engines allowing this idea to be realized. F. A.
Tsander came to the conclusion that it was expedient to use /61
lithium as not only an additive to the fuel, but also as the
structural material of a spacecraft.

In his article, "Problems of the Design of a Rocket Using
Metal Fuel," published in 1937, the requirements are set forth
for metals of which the structural elements later to be burned
in the combustion chamber were to be made. They are as
follows: the metal should be sufficiently strong at normal tem-
peratures, the light and heat of melting should not be too
great, the heat generating capacity should be as great as pos-
sible, the melting point -- low. His work presents a method for
determination and selection of the optimal dependence between
the mass of a metal on the one hand, and of the liquid oxidizer
and fuel on the other; between the masses of all structures and
of the metal burned, between the solid and gaseous combustion
products.

Raketnaya Tekhnika , 1936, No. 1.
2
Tsander, F. A. , Problema Poleta pri Promoshchi Reaktivnykh [The

Problem of Flight Using Reaction Apparatus -- Collection of

Works], Moscow, Oborongiz Press, 1961, p. 241.

56

In his work "The Problem of Flight Using Reaction Appara-
tus," the use of boron or liquid boron hydride as an additive to
the fuel, suggested earlier by Yu. V. Kondratyuk, was studied.
"However, boron will probably be used only as a powder for insu-
lation (amorphous boron) or in the form of rods subject to com-
pression (crystalline boron). Liquid boron hydride could also be
taken if kept very cold." 1 When boron burns in oxygen, the min-
imum quantity of solid product is produced with very high heat
liberation, which Tsander calculated at 3900 kcal/kg.

This work also suggests that solid nonmetallic materials
such as celluloid, etc. be used as additives to liquid fuel.
"Experiments could also be conducted to find pressed masses,
used in almost all areas of chemical technology and possible
for our purposes as well." And further — "We can imagine
masses containing naphthalene or other fuels in mixture with
materials which, when heated, would melt and then be fed from
the melting vessel into the motor pumps of the rocket as liquid
fuel." 2

Study of Processes Within the Chamber and Cooling Conditions /62

F. A. Tsander developed methods for thermal and thermo-
dynamic calculation of a reaction engine, presented in two
articles under the title "Thermal Calculation of a Liquid-
Fueled Rocket Engine," first published in 1936-1937.

In these articles, the author presented examples of cal-
culations for a fuel consisting of air enriched in oxygen and
gasoline. There also he analyzed the influence of the adiabatic
index, gas constant, gas temperature and degree of expansion on
the ideal exhaust velocity of gases from the nozzle; he presented
a method for determination of the area of the critical and exit-
plane cross sections of the nozzle; he studied the flow of
actual gases considering loses and considering the influence of
gas friction on the wall on the characteristics of the nozzle.

F. A. Tsander determined the combustion temperature con-
sidering dissociation of gases and constructed graphs character-
izing the thermal parameters of an engine as a function of the
oxygen content in the oxidizer.

In these articles, F. A. Tsander performed his calculations
not only analytically, but also using entropy diagrams.

Tsander, F. A., Problema Poleta pri Promoshchi Reaktivnykh [The
Problem of Flight Using Reaction Apparatus -- Collection of
Works], Moscow, Oborongiz Press, 1961, p. 119.

2 Ibid. , p. 117.

57

A number of the works of F. A. Tsander have been dedicated
to determination of the heating and cooling of the walls of a
rocket engine combustion chamber. These works analyze the
peculiarities of heat transfer from the gases to the walls of
LRE. The most detailed thermal calculation of the cooling
system for a rocket engine is presented in the article "Thermal
Calculation of a Rocket Engine Designed for Liquid Fuel."*
In it, the heat transfer coefficient is calculated on the basis
of the formula of Nusselt, after which formulas for determina-
tion of the external, internal and average temperature of the
engine chamber wall are given, methods are studied for deter-
mination of the physical parameters of cooling media and the
required flow rates.

Calculation of the cooling system helped F. A. Tsander to
determine the limiting possible pressure in the chamber for
each specific fuel composition. Using the results of calcula-
tions, he determined the thermal efficiency, thrust of the
engine, exhaust velocity and selected the volume of the com-
bustion chamber. Thus we see that the planning and construe- /63
tion of the OR-1 and OR- 2 engines were preceded by calculation.

The scientist set himself the problem of transfering rocket
technology from the area of theory to the area of engineering
practice. "I am primarily a mathematician," Fridrikh Arturovich
said of himself. However, analyzing the results of his activity,
we can bravely state that F. A. Tsander was a great scientist,
inventor, engineer, designer and experimenter. He created a
number of experimental installations, a flame stand, created
and experimentally developed the OR-1 reaction engine, devel-
oped the OR- 2 rocket engine and the initial version of the
GIRD-X rocket with the type 10 rocket engine.

Increasing Specific Impulse and Efficiency

F. A. Tsander studied various means for increasing specific
impulse. Giving this problem prime importance, he based his
studies on the use of a fuel with high specific heat content,
consisting of liquid oxygen and gasoline, and considered the
use of oxygen most promising. In addition to the use of metal
as an additive to fuel in order to increase its heat-producing
capability, he also suggested that specific impulse be increased
by acting directly on the gases leaving the chamber of the
rocket engine by installation of restricting fittings at the end
of the expanding nozzle, the so-called reverse cone.

Raketnaya Tekhnika , No. 1, 1936.

58

As heat is transferred from the supersonic stream to the
constricting fitting, the gas velocity should increase.

The increment in work of the cycle is obtained by the
additional adiabat-^ expansion (line CE) and subsequent iso-
thermal compression ^line EF) . F. A. Tsander called this
cycle the improved working cycle. The pressure at the outlet
of the nozzle p remains at the design level, and optimal

thrust is achieved; the temperature of the gases leaving the

nozzle T is decreased, consequently increasing the heat drop
a

generated in the combustion chamber. However, F. A. Tsander
did not consider the presence of compression jumps in the super-
sonic stream and did not study the possibility of producing
effective cooling of the entire mass of exhaust gases.

At the present time, thermal efficiency is increased by
increasing the degree of expansion of the gas by increasing the
pressure in the chamber with the optimal gas pressure at the
exit plane of the nozzle.

/64

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Results of Thermodynamic Calcula-
tions of the 0R-1 Reaction Engine
After F. A. Tsander

in the USSR by V. P. Glushko, A. P.
and other scientists.

In 1930, F. A. Tsander
developed an approximate
method for calculation »f a
reaction engine. He gave
particular attention to the
calculation of thermodynamic
processes in the combustion
chamber, allowing him to
determine the basic para-
meters of the LRE with the
necessary accuracy as they
were planned. During these
years, the approximate
method of design of a rocket
engine was developed and
successfully used by V. P.
Glushko at GDL. Later,
methods of thermodynamic
calculation considering
dissociation were improved
Vanichev, A. I. Polyarnyy

The OR-1 Reaction Engine

The first experimental reaction engine (OR-1) used com-
pressed air and gasoline. The planning and construction of the
OR-1 were preceded by laboratory experiments and creful calcu-
lations performed by F. A. Tsander. In 1917, he performed

59

Indicator Diagram of Theoretically
Improved Engine of F. A. Tsander

experiments on the burning of metal; beginning in 1922, he
selected and systematized the calculation dependences, without
which it was impossible to create the method of calculation of
LRE, developed plans and drawings for an experimental reaction
engine, and on 30 November 1928, F. A. Tsander read a report /65
at Moscow University entitled "Preliminary Work on the Construc-
tion of a Rocket Apparatus," in which he presented the results
of preliminary calculations and a plan which served as the
basis for development of the OR-1.

In October of 1929,
F. A. Tsander began
detailed design calcu-
lation of the OR-1
engine.

The initial calcula-
tion data were: gasoline
consumption 3S0-400 g/hr,
theoretical air consump-
tion per kg of gasoline --
14.2 kg. Thus, the fuel
consumption was approxi-
mately 1.67 g/sec.
Thermodynamic calculations
determined the composition
of the combustion products, gas temperature in the combustion
chamber, approximately 2440 K, thermal efficiency, or. the order
of 0.105-0.125, exhaust velocity, about 840 m/sec, and design
thrust -- 0.145 kg.

F. A. Tsander began assembling his engine immediately after
completion of the calculations and manufacture of the parts. 1

On 30 September 1929, he wrote: "Due to the funding problem,
I suddenly got the idea to redesign the torch for the first
reaction engine... I redesigned the fitting and surrounded it
with a sleeve, into which air was blown under pressure. Inside
the sleeve was a special tube forming the space for combustion.
At the end of this tube was an interchangeable conical fitting
to produce exhaust velocities greater than the speed of sound.

"The copper tube for liquid gasoline was replaced . ith a
longer one, which was wound around the conical fitting to preheat
the gasoline. Furthermore, the tank was equipped with a mano-
meter to measure the gasoline feed pressure and a nipple to let /66

The descriptions of the OR-1 presented by a number of authors
include many individual errors; we therefore considered it
expedient to present the description of the OR-1 given by Tsander
himself.

60

out air. A thermometer was attached to the tank to measure the
tank cover temperature. A special valve was fitted to regulate
the consumption of fuel.

"The compressed air for combustion and cooling of the com-
bustion chamber was fed into the cooling line through a nipple
attached to the sleeve in front of the nozzle. The mixture was
ignited by a spark plug soldered into the head." 1

Inlet valve I S P ark P lu «

The First Reaction Engine - the OR-1

The first tests
of the OR- 1 were
conducted by F. A.
Tsander in the
laboratory for air
aviation engines of
the screw-motor sec-
tion of TsAGI in 1930.
The engine was sus-
pended so that the
gases exiting from
the nozzle were
directed toward a
small metal disc con-
nected to a balance.
The indications of
the balance were used
to determine the pres-
sure of the gases on
the disc.

In 1931, the OR-1 engine was finally developed and begin-
ning in 1932 it was used to study the influence of the addition
of metals to the liquid fuel on thrust. The low accuracy of the
measurements characteristic of the time and the insufficient
thrust developed by the OR-1 did not allow any influence of these
additives on the operating mode of the engine to be measured, but
calculations still indicated the expediency of the use of metal
additives to the fuel. Therefore, studies were continued in
later years.

/67

The OR- 2 Rocket Engine

The OR-2 engine was developed by F. A. Tsander. The plan-
ning of the engine was begun in September of 1931, but preliminary
calculation of the units and of the engine as a whole had been
conducted by Tsander even earlier. The engine was designed for

,

Tsander, F. A. , Problems Poleta pri Promoshchi Reaktivnykh [The
Problem of Flight Using Reaction Apparatus -- Collection of
Works], Moscow, Oborongiz Press, 1961, p. 47.

61

installation on a piloted vehicle -- the RP-1 "flying wing"
glider designed by V. I. Cheranovskiy. This glider was manufac-
tured by Osoaviakhim activists.

Thus, the OR- 2 is the first domestic LRE designed for a
piloted vehicle. Liquid oxygen and gasoline were selected as
the fuel. The operating time of the engine was designed to be
30 sec, with a thrust of 50 kg and a chamber pressure of 6 to 8
atm. However, the OR- 2 was never installed on the RP-1 glider,
since the engine was never successfully developed. Later, not
at GIRD but rather at RNII, a modification of the engine (02) was
developed, differing from the OR- 2 in design and fuel used.

The combustion chamber of the OR- 2 had an elongated cylin-
drical shape, the nozzle was conical and supersonic. The mixing
head carried sprayers and an inlet valve for the fuel. This same
valve allowed thrust to be varied by gradually changing the fuel
consumption. Ignition was by an electric spark plug. The
cylindrical portion of the chamber of the engine was cooled
externally by the liquid oxygen, which entered the chamber in
gaseous form, and the nozzle was cooled with water.

The extractive feed system included pear-shaped fuel tanks,
which were to be suspended in the internal sections of the glider.
The fuel components were fed to the chamber under pressure
created by gaseous nitrogen. This was achieved by the use of
a "nitrogen compensator" --a separate tank containing liquid
nitrogen. The water cooling system for the nozzle included two /69
liquid oxygen evaporators, the nitrogen compensator heat
exchanger, a water tank and pump. The water, heated in the
nozzle cooling cavity, passed through the pump and tank into the
nitrogen compensator and evaporators. Heat exchange between the
water and the liquid nitrogen caused the latter to evaporate.
Additional cooling of the water occurred in the oxygen evapor-
ators. The oxygen gas was used to pressurize the oxygen tanks.
The cold water was returned to the nozzle cooling cavity. All
parts of the OR- 2 were placed in the glider. Assembly of the
OR- 2 engine was completed in December of 1932.

By early March 1933, the engine was installed on a test
stand at the Nakhabinsk range and prepared for testing. Since
Fridrikh Arturovich was then in Kislovodsk for treatment, the /70
flame tests were performed by his working colleagues.

The first test of the OR- 2 was held on 18 March 1933. The
feed pressure was first held rather low -- from 3 to 4 atm. The
fuel in the chamber ignited, but the combustion was unstable and
rough and in a few seconds the engine had to be shut off. During
the second test on 21 March 1933, one oxygen evaporator operated.
During the seventh sec^ad, the motor burst in the region of the
head. The third test was conducted on 26 March. The feed

62

/68

Chamber

77

oooo
Gas 6 line
Symbols:

— Fuel

— Water

-- Nitrogen

Diagram of the OR- 2 Engine and External View
of Its Chambers: 1, Gasoline Tank; 2, Safety
Valve; 3, 20, Oxygen Tanks; 4, Evaporator; 5,
Combustion Chamber; 6, Valve; 7, Pump; 8, Water
Tank; 9, Additional Heating; 10, Roller; 11,
Line; 12, Nitrogen Under Pressure; 13, Evapor-
ator; 14, Control Panel; IS, Manometers; 16,
Thermometers; 17, Values; 18, Magnito; 19,
Valves; 21, Cylinder of Hot Water; 22, Nitrogen
Compensator

63

system was operated with two evaporators, but the combustion of
the fuel in the chamber was rough, and in a few seconds the
chamber burst on a welded seam. The cooling jacket burned
through. During the fourth test, on 28 April 1933, the pressure
in the chamber changed suddenly, but at times briefly stabilized
and held almost constant at 8 atm; the feed system operated with
two evaporators. In danger of damage due to the great dynamic
loads developed, the engine was shut down in the 35th second.

During the first
tests of the OR- 2, the
members of the team
held to the working
style of F. A. Tsander
and followed his
instructions to test
the entire motor at
once , i.e., the com-
bustion chambers
together with the
fuel feed system and
supplementary appa-
ratus. This method
of testing is more
complex than stage -
by- stage development
of units but, as
Fridrikh Arturovich
believed, it allowed
more complete consider-
ation and clearer determination of the interrelationship of all
processes occurring in the engine.

It is hard to decide what plan of further testing F. A.
Tsander would have suggested after analysis of the results of
the first flame tests. We know that he did not deny the possi-
bility of using oxygen- alcohol fuel; therefore, after processing
of the experimental data, gasoline was replaced by ethyl alcohol
in further experiments .

The combustion chamber was simplified and equipped with a
refractory heat insulating lining consisting of aluminum oxide
and magnesium oxide; an extractive fuel component feed system
was installed, consisting of the fuel tanks and a gas accumula-
tor --a high pressure cylinder. A valve an reducer were
installed between the cylinder and tanks in order to reduce the
pressure. This new version of the engine was called the 02.
Subsequently, only the combustion chamber, rather than the
entire engine with all of its units was developed. The descrip-
tion of the 02 engine is presented below.

Diagram of Placement of OR- 2 Engine
Components on the RP-4 Glider

ai

64

(J

Diagram of Second Version of ST Rocket
Engine After F. A. Tsander: 1, Liquid
Fuel; 2, Oxidizer; 3, Injector; 4, Heater
or Injector; 5, Combustion Chamber; 6,
Oxidizer Evaporator; 7, Compensator

Plans of Rocket
Engines

In addition to
fie OR- 2 engine with
the extractive feed
system, F. A.
Tsander developed
several other
designs with injec-
tor fuel component
feed. K. E. Tsiol-
kovskiy believed
that injectors
could use a portion
of the energy
liberated in the
combustion chamber
to feed liquid fuel
components by means
of a stream of gas.
F. A. Tsander did
not produce such
systems.

F. A. Tsander made up a general engine plan with a turbine-
pump fuel- feed system, and suggested that a gas turbine be used
to drive the pump, the working fluid for which would be the com-
bustion products of the fuel, drawn away from the main combustion
chamber. At the present time, in order to produce turbine gas of
relatively low temperature, the turbines are supplied not by the
combustion chamber, but rather by gas generators.

O
0' :>

v

Diagram of Third Version of 5T Rocket
Engine After F. A. Tsander: 1, Into
Evaporator; 2, 02 or N2O4 or Ozone, 3,
Alcohol or Other Fuel; 4, Pumps; 5, Com-
bustion Chamber; 6, Gas Turbine; 7,
Evaporator; 8, Double Cone; 9, Water
Compensation Tank; 10, Water Pump

One interesting
design developed by
F. A. Tsanuer is an
engine plan in which,
in addition to the
usual liquid fuel,
powdered and liquid
metal fuel were to be
used. The powdered
metal was to be fed
into the combustion
chamber by an injec-
tor. The liquid
metal was to be
produced by melting
metal structural
elements of the
rocket no longer
necessary in flight.

41*

65

Di .gram of Engine 10 for Initial Version
of Liquid Fuel Rocket After F. A.
Tsander: 1, Powdered Metal Fuel; 2, Pump;
3, Solid Fuel; 4, Liquid Fuel

izing of metal were also unsucces
interesting ideas from the plans
injectors and of metal fuel, have
realized.

The studies
performed in the
30' s and 40* s with
injectors showed
that, in spite of
some promising
theoretical data,
they operate only
at very low effi-
ciencies.

Attempts to
develop an accep-
table engine design
using metal as an
additive to the
primary fuel were
unsuccessful.
Attempts to con-
struct units for
melting or pulver-
sful. Therefore, two very
of F. A. Tsander, the use of
not as yet been practically

Fridrikh Arturovich Tsander was one of the pioneers of
rocket technology, combining the talent of a great theoretical
scientist and that of a gifted experimenter and engineer. He
developed the principles of the theory and design of LRE and
performed detailed calculations of his first experimental speci-
mens. The theoretical and experimental developments of Fridrikh
Arturovich aided further development of research on LRE and
rockets, while the propagandistic activity of Tsander and other
enthusiasts led not only to the creation of GIRD in Moscow, but
also in many other cities of the country. Thanks to the practical
activity of F. A. Tsander at GIRD, the first Soviet rocket with
LRE was created and the OR- 2 engine was developed.

66

Move forward bravely, great and
small laborers of the world, and
know that net a line of your work
will be 1o*l, but will bear for
you great fruit.

K. E. Tsiolkovskiy

Chapter 2. The First Rocket Scientific Research and /74

Experimental Design Organizations in the USSR

The works of the early scientists of our country, their
students and followers formed a basis for the development of
scientific research and experimental design work on the creation
of rocket engines and rockets in the early 1920* s.

The reason for this development was not only the successful
results of the studies of our scientists, but also, to a signi-
ficant extent, the demands of various branches of science and
technology, particularly aviation and artillery

Thus, in the 20's the time had come for the transition to
experimental work, for the creation of creative teams md the
expansion of the range of scientific research work. Rocket
technology had to be brought from the area of theory to the
area of engineering practice, had to be given statewide signi-
ficance. "But we must undertake experiments. We must consider
nothing in our theoretical works to be absolutely true," 1 said
K. E. Tsiolkovskiy.

The development of the national economy, the rapid growth of
science and technology in the USSR, the successful fulfillment of
the first five-year plan allowed scientific research organiza-
tions to be sot up on the country for the development of rockets
and liquid- fueled rocket engines.

During these same years in the Soviet Union, many public /7S
organizations were developed which were of great significance
in the popularization and development of rocket and space tech-
nology. Some individual scientists made great contributions to
the development and popularization of the science of rockets and
engines.

The leading organizations in the USSR were the Gas Dynamics
Laboratory (GDL) under military auspices, which began its

tsiolkovskiy, K. E. , Collected Works . Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 274.

67

activity in the spring of 1921, the Group for the Study of
Rocket Motion (GIRD), a publically supported group begu- in
wiie fall of 1931, and the Reaction Scientific Research Insti-
tute (RNII), developed on the basis of GDL and GIRD late in
1935. The rocket organizations expanded, changed their purposes,
new large government enterprises were developed, solving complex
problems of the mastery of space.

Now, probably, it would be simply impossible to list all
of the large and small problems, the entire range of problems
studied by the subduers of space. There is no science or tech-
nology which has not been utilized to some extent in the study
of space; it is difficult to name a science whose development
has not been influenced by the results of the study of space.
The performance of such a grandiose program of difficult inves-
tigation, leading to the accumulation of a new wealth of know-
ledge by mankind, requires the harmonious development of all
areas of science and technology, leading tf an avalanche of
inventions and disco series.

However, things were different during the first years of
development of rocket technology. It vas impossible then to
begin immediately to solve problems of cosmic scale. Even
Sergey Pavlovich Korolev believed that dreams of flights to the
moon and of new speed records by rocket airplanes not yet in
existence were useless until scientists could create at least
small liquid- fueled rockets. For this reason, every researcher,
every worker in the area of rocket technology had to hold the
reaction engine at the center of his attention. If a reliable
engine could be built, Sergey Pavlovich believed, all other
problems arising in the process of work with the flight vehicles
could be successfully solved. "Our success," S. P. Korolev
wrote, "requires first of all a reliable, high quality motor. "* /76

It is therefore quite clear that the main task of GDL,
GIRD and RNII was the creation of LRE corresponding fully to the
requirements placed upon them.

2.1. The Initial Period of Development of GDL -- the
N. I. Tikhomirov Laboratory

Nikolay Ivanovich Tikhomirov and his immediate colleagues,
Vladimir Andreyevich Artem'yev, Geortiyey Erikhovich Langemak,
Boris Sergeye.ich Petropavlovskiy and many others made a great
contribution to the problem of creation of military rockets with
powder rocket engines.
_

*Korolev, S. P., Raketnyy Polet v Stratosfer e [Rocki t Flight in
the Stratosphere], ONTI Press, Moscow, 193'."

68

•v - fcork i the eteat ton - : rocl ■ • hu 5 . • s® •*••' •• s
.<-,:•? ul teJ many model v. o rockt : various or \ J

n poses * • h .*:•'--..-,. '• • $v - j '. • l€ iltdioy, tl new

. -• . •. » ipo ■ dev< I03 > 1 nt 19 i ,-' ale < - -<■■■. '•■. 1 ■- leeoi .1
"-■ jr Id fc'a i ;- lu ;u it J ' s i .- - : . •• .

Chora ical ijiinec i . Ti tomi -o\ ■- to 3 M perios sued

experimental work as earl v as 1894 to determine the forces of

ict ioi f •<■ • . • in order to t • ■ the react iv foi cc

tr, powei mi I it ar j lev ices , in 1 3 . f i ;hc lire ore: en te« *

i '■'. aat . !i ■ ipti •■ of - -'■, II lo ! Xav . ti s ry Ii ••'-12-

3 J . thi; plan -■■ ■ further te • oped and success full; passed

many tests, although conditions were provided for its production

only after the revolution In -\- ' ••'•' ■ hom>roi cut "o

\ .' . - --. • . • > o\ te • I .-' Pc , • ( dm Si a . V, P.

; . ••. B u) L< quest l call he at? t on »f ... ; euin

to t : • iceJ :i srod ice . i invent ioi in o da t< I i ud the

.- , , . ••• \ ■■ m ,* ■• -pel S\ I . Til mirov at I " :J ed

a do* cri i • '; . inven ■ • , te ?rt i f icate f Invent io?

*-J tch h »a: m kn in . i a .' ;he ; - t i 'c i epoi I •" the
Cha i rman o f t h -,-•,.■•'' in t ion c ■ ' ■ • , . %\ 1 i to ry
and ... > • • '•-'.'. te* iTK) , ! >.••>•,-,• , •• hu -•;»'»
:• i ded i i'Jio to hi lo i • e object v stu h of 1 te
pi in in the Committee o> riven i< n " \t liery ( .,-.',.,

hi Lo . .-.,* * in Chi >f all armed f -•-'-:. e . . . ,
5, S. * ,• -"• -, and ;., v-- h nan •• : ' i an loisnci of the

e| jbJ ii ordci .' rapid .' r iopment ol i*ork " I •-. invent on of
X, I. Tikis sii - ri ;*-.••" - .. '.'• curoen o. flic . lo .•>.- '.. iot<

t ; .• '■•.,-.• . .' I tO ; ,-- •-.' tlO« iif X. I .

Ti :h -' '•". an I point ' < ; thai i \ was
c< iderod > > • , s ignif ic; c< foi ' ; ?
;tat< fj Morel ol 192 • , *. I . Ti! h? mirov
, a: . iv n u'orJ »p; - •' in to o o* o > f ikhvin-

-.: ay;i '• • c t • two itor; ildioti Mo >) lo
s . u] . ; lahoi ••:': u j . I h , ■ ' ■• as
3 Lsc . wai ."••-, - he tcccss \ , " i >, : ' .

'-■ • ; -i r . ! -, t .? -'.-■ :•„'"';•.• < • * - in 4 Jg
and mat < . mpport . ••.■:• f the

an , a ■ ion »f tlu l.ohor.-itor; i a be
i ted a . ■' Ma :h -. ■ - 1

|ili^^»

i •• I - ••. - . ■•

Tikhomirov

In Ma • oi • ' ; , tb< .' '1 i t lory

Coairlttee seat V. A. Arteiu'vev, the "rocket

. : . - ., sper< ! • • ■ .- . .• '■ l Hi i n e
Vrti I or; L'omrni ttee, Lo -•;. i| > 1 .

! khoairov.

\ ! • An* -eye hi cm "ye ( 1885-

. hi ill ? /• - on :h< coo t riu " m

1 I'oj ■• :-' ' ; • I on i i -• in i . ti ton . - the GDI, ■ '■ • • : ■• • • ,: *
iaul SSH 191 : ■• . »p M) I • ; --'

§9

of military rockets in 1913. At that time, he was serving at /78
the Brest-Litovskaya fortress and, as head of a laboratory, was
studying the improvement of the three- inch rocket flares produced
by the Nikolayevskiy Plant and used by the armed forces.

In 1921, shops, a pyrotechnical laboratory and a chemical
laboratory were set up. This complex was called the "N. I.
Tikhomirov Laboratory" and was subordinate to the Military
Department. However, the work on the creation of military
rockets moved forward very slowly; difficulties arose primarily
due to the lack of high-eneigy, slow-burning powders. It became
obvious that development of self-powered mines would require sig-
nificant funds and time. Therefore, in April of 1923 the inven-
tors were ordered to perform experimental tests of the applica-
bility of reaction power for existing mines in order to increase
their range.

Between 22 March and 3 April 1924, 21 launches of these
rockets were conducted at the main artillery range in Leningrad
under the direction of V. A. Artem'yev, showing a ten- times
increase in range of the mortar shells used. These experimental
tests confirmed the promise of the new type of shell and the need
to perform further work in this direction.

The experience of preceding investigations had indicated
the inapplicability of available powders for the manufacture of
rockets, since they did not burn evenly, or were not sufficiently
effective. Smokeless pyroxylin powder, widely used in artillery,
did not yield positive results.

Smokeless powder (pyroxylin cartridge powder) was first
suggested for rockets in 1915 by I. P. Grave, but the rockets
being developed required slow-burning powder charges with great
top thickness. The preparation of such charges using known
formulas for pyroxylin powder based on volatile solvents encoun-
tered unsurmountable difficulties. The charges were warped and
cracked during drying, resulting in variations in burning time
and speed. Consistent results also could not be achieved in the
percent content of solvent remaining in the charges after drying.
During storage of the charges, the solvent evaporated, also
causing variation in the parameters of combustion of the powder
charges.

In order to avoid these shortcomings, N. I. Tikhomirov
decided to try smokeless powder with a nonvolatile solvent. The /7S
development of this smokeless powder was undertaken in 1922 under
the leadership of N. I. Tikhomirov in Leningrad with the partici-
pation of 0. G. Fillipov aid S. A. Serikov. This work was of
grea* scientific and practical significance for the development
of rockets and space technology. The first specimens of thick-
top powder drains of the new formula -- trotyl -pyroxylin powder

70

(using trotyl as the nonvolatile solvent) -- were produced in
1924. This powder, called PTP, was then Manufactured in the
powder shops of the Leningrad Steamship Port. These shops were
assigned to N. I. Tikhomirov, and became a part of the laboratory.
Powder testing was conducted at the Scientific Research Artillery
Range near Leningrad. Powder studies were continued at the
Military-Technical Academy imeni F. E. Dzerzhinskiy in Leningrad.

The basic model used in testing and experimental development
of charges was a grain with an external diameter of 24 mm and an
internal channel 6 mm in diameter. Later, the grain diameter was
increased to 100 mm.

The creation of a stable high-energy smokeless grain powder
with great top thickness was a great achievement, providing a
qualitative jump in the development of solid-fuel rocket design.

All of the most important work of the laboratory, related to
the development and manufacture of a smokeless powder, test
stand operation and experimental firing, was conducted at
Leningrad. As a result, in 192 5 the laboratory was transferred
to Leningrad completely.

After careful development and testing of grains and launching
devices, on 3 March 1928 the first firing of rockets with charges
of smokeless trotyl pyroxylin powder was conducted at the main
artillery range.

In his memoirs, V. A. Artem'yev wrote that no data have been
found indicating that foreign armies successfully tested rockets
using smokeless powder earlier than our own.

The creation of a smokeless powder rocket laid the founda-
tion for the design of the "Katyusha" military rockets.

2.2. The Gas Dynamics Laboratory /80

Following the successful launch of smokeless powder rockets
in 1928, the N. I. Tikhomirov Laboratory was expanded and
renamed the Gas Dynamics Laboratory (GDL) , subordinate to the
Military Scientific Research Committee of the Revolutionary
Military Council, USSR. The first task of the GDL was the
development of solid-fueled rockets utilizing high-quality
smokeless powder charges. Soon, GDL also undertook the creation
of powder takeoff assist and landing brake rockets for airplanes.
Based on the successful results of experimental work by N. I.
Tikhomirov and V. A. Artem'yev involving the creation of rockets,
the Main Artillery Administration of the Red Army decided to
send specialists to the GDL and to expand its production and
laboratory base.

71

The primary experimental research base of GDI was stationed

1928 at the Scientific Research Artillery Range (N1AP) near
the Design Bureau -- in a building at the Artillery
Research Institute, then at the Admiralty! the Admin*
- in Lengrad at No, 19 Khalturin Street, the powder
the steamship port, the aircraft testing base -- at

at Pctropavlovsk

in

l. en i ngrad ,

Scientific
i s t ra t i on
shop at

the military airfield, the mechanical shops
fortress and elsewhere.

"he Gas Dynamics Laboratory* was great 1 v

id*

the Lhair-

man of VSMKh, later People's Commissariat for Heavy Industry,
G. K. Ordzhonikidce and, particularly, by Marshall of the Soviet

Union M. N. Tukhachevskiy , the immediate superior of the GDI..

Boris
Petro;

1 - i a
the R

armed

Se

i rp

S-8
wi

lo

la

th

cyevic

vskiy

ne i n
from t

roc

:h

1932.
he I-
ket 1

In 1930, at the age of 70, N, I.
Tikhomirov died. On the 50th Anniversary

of the foundation of the GPL, a monument
to its creator and leader, patriot and
scientist N. I. Tikhomirov, was erected
at the GPL, The name of N. 1. Tikhomirov
is permanently inscribed in the history
of rocket technology in the USSR and has
been given to one of the craters on the
far side of the moon.

In 1930, the GPL was taken
talented tnttitnrv artillery engi
Boris Sorgeyovich Petropavlovksi
1930 to 1933, powder rocket bomb

various sizes were developed at
including 00, OS, 70, 82 and 132
in mid- 1931, the rockets produce
GM> were used as a basis for pla
aircraft takeoff assist devices,
practical fir* **» e tests of the RS
to-ground rod were conducted

n the summer c

ii ■

1 aircraft was eonuucted
tunc hers .

official f
, using an

over
neer ,
§1 i
s of
GPL,

cal
d at
n s to

and
-82 a
from
iring
airpl

py a
rout

i be i .

l or

i r-
an

of
ane

/81

Together with the improvement of air-to-ground rockets,
extensive studies were conducted on the use of ground- launched
rockets fired by special lightweight launching devices.

During the Second World War, a very significant weapon was
the barreliess mutliple-eharge mortar -- the BM-13-SN, BM8-48,
BM31-13 launchers and their modifications designed to fire
rockets. Poring the war, the people called this weapon "Katyusha,"

The development of the Katyusha launcher can be divided into
three main stages. During the first stage (1921-1929), the smoke -

7?

less powder was developed, the principles of design of solid-
fueled rockets were determined and flight testing was begun; in
the second stage (1930-1933), rockets were produced, passed
official testing and during the third stage (1933-1941), the
Katyusha rocket launcher was developed.

Experimental work with solid-fueled aircraft takeoff boos- /82
ters and landing braking devices began in 1927 using a powder
catapult, then later with the U-l training aircraft. Beginning
in late 1931, work on a solid-fueled takeoff assister was
conducted with the TB-1 aircraft. On 14 October 1933, the
TB-1 aircraft, equipped with a rocket-assisted takeoff device,
successfully passed state testing; the use of RATO reduce takeoff
run length by 77* with a flying weight of 8 t. RATO devices
were developed by B. S. Petropavlovskiy, G. E. Langermak and
other. V. N. Dudakov, pilot S. I. Mukhin and mechanic A. I.
Gritskevich assisted significantly in the development of
takeoff techniques.

In 1933, work was begun on a RATO device for the TB-3
aircraft, flying weight 20 t. In 1934, the Red Army Air Force
Command decided to conduct tests of reaction takeoff boosters on
three TB-1 aircraft. One test aircraft undertook a special test
trip from Leningrad to Moscow and back. On the whole, the tests
confirmed the effectiveness of the use of such boosters. The
advantages of aircraft takeoff with boosters became obvious to
all.

In addition to the development of rockets and rocket engines
based on solid fuels, beginning on 15 May 1929, GDL began to work
on the first domestic rocket engines: electric engines (ERE) and
liquid- fueled engines (LRE) , In 1931, GDL was divided into seven
sectors (called sections after 1932): 1 -- Powder Rockets (Chief
G. E. Langemak); 2 -- Liquid-Fueled Rockets (Chief V. P. Glushko) ;
3 -- Aviation Applications of Rockets (Chief V. I. Dudakov); 4 --
Military Rockets (Chief N. A. Dorovlev) ; 5 -- Powder Production
(Chief I. I. Kulagin); 6 -- Production Section (Chief Ye. S.
Petrov); 7 -- Administrative and Financial Section. Between
1930 and 1933, the number of workers increased from 23 to
approximately 200 persons.

The organizer and leader of the work of ERE and LRE, the
designer of the world's first electrothermal rocket engine and
the first domestic LRE, was Valentin Petrovich Glushko.

The development of the ERE and LRE involved the participa-
tion of V. I. Serov, A. L. Malyy, Ye. N. Kuz'min, I. I. Kulagin,
Ye. S. Petrov, P. I. Minayev, B. A. Kutkin, V. P. Yukov, N. G.
Chernyshev, V. A. Timofeyev, N. M. Mukhin, I. M. Pan'kin and /C3
others. The names of many of the scientists of GDL have been
given to craters on the far side of the moon.

73

V. P. Glushko was born on 2 September 1908 in Odessa. He
began to study problems of rocket flight in 1921. In 1923, he
began corresponding with K. E. Tsiolkovskiy, who mentioned
V. P. Glushko in the foreword to his books "Investigation of
Space with Reaction Devices" (1926), "Space Rocket Trains" (1929)
and other publications among those persons facilitating the
popularization of the ideas of star flight by their publications.

From 1922 to 1924, V. P. Glushko worked at the Odessa
astronomical observatory as an astronomical observer. The
results which he produced were published in 1924-1925 in the
Astronomical Bulletin and the Journal of "Nirovedeniye" Society.
The young astronomer was selected as an Associate Member, then
an Active Member of the Russian Society of Astronomy Enthusiasts
(ROLM) . Upon completion of his studies at the Department of
Physics and Mathematics of i cningrad State University (1925-
1929), V. P. Glushko began work at the Gas Dynamics Laboratory.
His thesis, dedicated to the development of rocket engines,
attracted interest and was approved by the experts of the
Department of Military Inventions (N. I. Tikhomirov and M. V.
Shuleykin) . The materials of his thesis served as the first
basis for the development of experimental ERE and LRE at the
Gas Dynamics Laboratory. V. P. Glushko is the author of a
number of scientific articles and fundamental works, including
the books "Rockets, Their Design and Application" (together with
G. E. Langemak, 1935), "Liquid Fuel for Reaction Engines" (1936),
etc.

A leading scientist in the area of physical and technical
problems of energetics, V. P. Glushko was selected in 1953 as
a Corresponding Member of the Academy of Sciences USSR, in 1958
as an Academician. He has been twice named a Hero of Socialist
Labor, is a Lenin and State Prize Laureate. Valentin Petrovich
has been repeatedly elected as a deputy to che USSR Supreme /84
Soviet. In 1972, the International Aviation Federation (FAI)
awarded V. P. Glushko an international certificate as a great
Soviet scientist in the area of development of rocket technology
and investigation of the physical and technical problems of
energetics. The FAI resolution is an international acknowledgement
of the great contribution of our coi ,r, try to the study and investi-
gation of space.

At the Gas Dynamics Laboratory, the possibility of practical
creatron of an electric rocket engine was proven in 1929-1930.
However, it was not possible at that time to solve the entire
range of problems related to the final development of ERE.

Therefore, the primary attention of the Gas Dynamics Labora-
tory was concentrated on the development of LRE and the investiga-
tion of processes of operation of these engines. In 1930, V. P.
Glushko suggested and subsequently studied various "uel

74

components: nitric acid, solutions of nitrogen tetroxide in
nitric acid, tetranitromethane, hydrogen peroxide, perchloric
acid, beryllium, liquid fuels and powders with dispersed
beryllium; in 1933, he suggested a mixture of liquid oxygen and
liquid fluorine as oxidizer, and solutions of pentaborane in
kerosene as fuel, as well as a fluorine-hydrogen fuel and many
others.

In 1931, he suggested hypergolic fuel and chemical ignition.*
The fuels used included gasoline, kerosene, toluene, benzene and
others.

During these same years, experimental development of indivi-
dual elements of liquid- fueled rocket engines was conducted.
Ceramic insulation based on zirconium oxide and magnesium oxide
was tested in the combustion chambers of experimental powder
engines (1930) . These combustion chambers were also used for
ballast pendulum tests to determine the most favorable for the
time exponential nozzle contour (1930). Measurement apparatus
was created for test stand studies of engines: spring and capaci-
tive pressure recorders and thrust recorders, inductive flow-
rate sensors and time recorders utilizing magnetoelectric /8S
oscilloscopes, etc. LRE devices with automatically controlled
variable thrust with constant pressure in the combustion chamber
were developed.

in 1930, the first LRE developed in the Soviet Union, the
ORM-1 laboratory engine was designed.

In 1931, flame testing of engines was begun at GDL. Unitary
fuels, solutions of a fuel (toluene or gasoline) in nitrogen
tetroxide, were tested in the ORM laboratory engine. 3 The ORM-1,
manufactured in 1930-1931, was designed to use nitrogen tetroxide
and toluene. Test stand operations were performed with liquid

A hypergolic fuel is a two-component liquid rocket fuel which
ignites at room temperature when the two components contact each
other.

Chemical ignition means ignition of the basic fuel in an
LRE, in which the basic fuel consists of hypergolic components
or a hypergolic supplementary starting fuel is used, introduced
to the combustion chamber only during the initial period of
operation of the engine.

2 The basic designation ORM was given to all LRE developed under
the leadership of V. P. Glushko in the GDL and at RNII.

^Nitrogen tetroxide is a high-boiling-point oxidizer for LRE. It
provides greater specific impulse than nitric acid, but is
inferior in the operational respect, since it has a narrower
liquid-state storage temperature interval.

75

oxygen and gasoline, since experiments performed earlier with the
ORM showed that it was very dangerous to start the engine with a
high-boiling oxidizer, particularly considering the complex
shape of the ORM-1 combustion chamber.

In 1932, eagines from ORM-4 to ORM-22 were developed, con-
structed and tested for experimental purposes. Liquid oxygen,
nitric acid, nitrogen tetroxide and solutions of nitrogen tetrox-
ide in nitric acid were used as oxidizers. Nitrogen tetroxide
was produced on a pilot-scale installation at the laboratory,
developed and put in use in 1931. Fuels tested included gasoline,
benzene, toluene and kerosene.

During the tests, start-up was developed and the organization
of processes within the chamber was improved, and methods were
developed for reliable cooling of the combustion chamber.

In 1933, experimental LRE from ORM-23 to ORM-49 were pro-
duced at the GDL and used to continue studies of problems of
LRE design. In order to create LRE providing sufficiently
high specific impulse and operating stably with identical indi-
cators in a scries of tests, i.e., reproducibly, reliably and ,«-
developing the required thrust, it was necessary to select fuel - —
components and the most favorable ratio of components, to develop
methods of feeding the fuel to the combustion chamber, and to
learn to organize the process of its combustion. This same
year, practical LRE were produced -- the ORM-50, 0RM-S1 and ORM-52,
burning kerosene and nitric acid both in pure form and mixed
with oxides of nitrogen. These engines used the principle of
chemical ignition developed at JPL, i.e., ignition by means of
hypergolic fuel. A number of experimental rocket -powered flight
vehicles were planned in 1932-1933 to test the engines under
flight conditions.

2.3. Liquid and Electrical Rocket Engines and Rockets
of GDL

The Gas Dynamics Laboratory studied and developed an electric
rocket engine (ERE), liquid-fueled rocket engines (LRE), called
at that time ORM, and experimental models of rockets, called RLA.

Step by step, the design of the individual elements and of
the engine as a whole was improved, which finally led to the
creation of a rather good liquid- fueled rocket engine for the
time, the ORM-52.

"Of particular promise," wrote M. N. Tykhachevskiy in
1932, "are the experiments at GDL on a liquid- fueled reaction

76

motor, which has recently been produced in their laboratory."

In the summer of 1932 and in January of 1933, GDL was
visited by S. P. Korolev, F. A. Tsander, M. K. Tikhonravov,
Yu. A. Pobedonostsev and other leaders and workers of GIRD,
who witnessed the operation of the LRE constructed at GDL.
Thus were the first meetings between the workers of GDL and
GIRD conducted.

Experimental Electric Rocket Engine

K. E. Tsiolkovskiy mentioned the possibility of using
electricity to drive rocket engines. In 1933, K. E. Tsiolkov-
skiy wrote, "The best transmission of energy is transmission by
means of electric current. But how can electric energy be con- /87
verted to mechanical work?... Electric current can be used to -
produce high temperatures and chemical decomposition of matter."

The designer of the world's first operating electrothermal
rocket engine was V. P. Glushko.

In 1928-1929, he developed a plan for a space rocket ship --
a heliorocket plane, driven by electric power produced by means
of solar batteries surrounding the ship in the form of a disc.
In April of 1929, the Military Scientific Research Committee of
the Revolutionary Military Council of the USSR received the work
"Metal as an Explosive Substance. A Reaction Engine with a High
Exhaust Velocity," by V. P. Glushko. 3 This work, a part of his
plan, served as a basis for the creation of the electric rocket
engine at GDL.

At first, the GDL division involved in the development and
testing of ERE (1929-1930) was colocated with the high voltage
laboratory of the Institute of Physics and Technology, directed
by Academician I. A. Ioffe. The laboratory itself, headed by
Academician A. A. Chernyshev, was located at Lesnoy near Lenin-
grad and in 1930 was reorganized as the Electrophysical Institute.
In 1932-1933, work on ERE was conducted on the territory of the
Ioanncvskiy ravelin of Petropavlovskaya fortress.

The experimental work was preceded by analytic calculation.
Then, engines of various types were made and tested; studies of

"Cosmonautics," Moscow, The Soviet Encyclopedia , 1970, p. 93.

2 Tsiolkovskiy, K. E., Collected Works , Vol. 2, Academy of
Sciences USSR Press, Moscow, 1954, p. 417.

3 GDL-0KV Archives, d. 1, pp. 1-16.

77

the properties of various conducting fluids and metals were
also conducted in order to determine the possibility of using
them as working fluid: .

In the Author's Certificate awarded V. P. Glushko, the
inventor of the LR1:, on 23 March 1931 l , various engine plans were
suggested in which the substance for the electric explosion was
introduced continually to the chamber. (Electric explosion
refers to the rapid conversion of the substance introduced to
the chamber to the gaseous state.)

in 1929-1930, two types of cant inuous- feed systems, called
carburetors were developed: a liquid system for a liquid working
fluid and a wire system for a wire working fluid.

/8S

•

External Appearance of the HI cc trie Rocket
Engine (F.RH) Designed by V. P. Glushko

In the liquid carburetor, the liquid was supplied from tanks
through tubes with interchangeable spray fittings of various
diameters. The liquids used included mercury, an aqueous solu-
tion of copper sulfate, a weak aqueous solution of nitric acid
and other substances.

In the wire carburetor, the feed mechanism consisted of
two steel rollers with a guide device. The rollers were driven
by an electric motor through a reducing drive. The working
fluid used consisted of metal wires of various metals (copper,
nickel, tungsten, lead, etc.) and carbon filaments. The fre-
quency of explosions was raised to several dozen per second.
The explosions were recorded by a photographic camera through
light filters. The gases produced by evaporation of the metal
were accelerated in an ordinary no;:le.

T

GDL-OKV Archives, d. 3, pp. 24~2b, 47.
78

Tli ■ ; '.•" wen upp i >•-.-,. .-■■■> by : hi h- if tci ?j ec-

tric »u s - *• k ric< :hc b< sic elements >1 ' ; '.'- en - i gh~

- tage transform Com ?ctif r: and oil {] ' - loodens -rs
.-' . h i ap . itancc o •* uf, charged up to 40 kv.

Tie- effect i n i rid i tdu I : . I • t -,., sion ••* 3 :er-
tninci 1 •- 19331 n .■"'+■ oi t bal istii >enduium. II was

lemon; - ' that lu • ' * : ic a I - "'. sro ic u . • ai - velo-
city oi several tens ..<:. kilometers per second.

In t ariy 1931 f Lore f i e« id chnolojj die

allow i ffe< i . Rl and <■ ■ t on- h *ard >le ri p< w -r
-.: •■ L) to bi . . iteU hei - • e, th •••■ - .•/'•- . . - • p ■• r-

;• - ' i tp : ; - " d V. P. . hi eg in rkio f I 1 m< o - -iO
studj and development of IRE,

A n '■ a e in t - den loj nc I ol el '-,•*. i< rockei -ngines
n the late 40' .- . : • ri) .'•"•. sen r! »vemeii . ri a
umb oi n . . : -. ie " were found t be s? I • c - - he

de¥elopnent of experimental work in the USSR and the USA, At
the : c ik , in cot leeti :h achievements i the a oi t se

• tt . ; *'. 1 ~v - i ' tieal need for ERE appeal ,-'

Me--' '".'., *,:,>;• .. si s • ■ , various les ip\$.

In - eie ottn ,• - : :ke1 .-.••■• the working fl - : . ,
olid or low-aol u ?ight hel iuro, hydro ••-•• et<

h • : -- ' to • : - tempo a i t i e h mean of an electric ai : raic

leal ."', or om« oth method • ■ •• • c heating ! '> l . he; *ed
work in fluid h ■' . : ! itc* in in on narv ic :1« tc toe . >
of nol om r 20 1 m/< e c .

A • ■• r sol ut ion i i . e< roraagnet is. asma j s -. ; • -

engine, in which the working fluid is converted to a plasma and

ce era ?d b ■- :-,.. of a . c lee ;. nnagiu -. field sc tni on the
plasm; f n th engii , hay .■>•<.. s of hundred of

k ti/ » ". , an b * at ?• i ev -1,

At thei mod n Ri is he iccti u - ■ , ' - : '. sgine,

in v ' - ' tl wo "•- >.• -.:-'-- m rn liun • : ir •• , on,
Is f irsi i ited mt i 1 ioni: : : " • - kind the as it ive
ions urine ■ • , -. ; - - , -• : i ■ • ror lectro: it ic field to
veloe it i es o tens or undreds of km/s ec . A •-' *ec ial " t : '^r
neu ! 1 - • • ' - ! read ion s - earn <? it I - yet ton

li : ;- •; . >^ MRl., tl pt ■■ , '■ • . :>" at pei ifi i-«pi i^e is
;e ected, he s/ali - oi which depends • • ,\ ; -- • ai ■'" ent on

tlu • •; j h -..' I | cm i ha , ..-:•> ; • - , . :he on-boi d electric

mpj < he parameter* J :he el rtrit co: ent conve rt . rs

in ,..•- : ?r pi 1 • •' the ,•'-•;:!-•. .on. In urdcr tt prodm ' e ect

■ iwer on >oa - : -. . -afi » I : lie; I n i< eai - i >lai pc er /'

plants ' are used',

?9

ERE generally develop low thrust, but can operate over long
periods of time.

At the present time, electric rocket engines are being
used by the USSR for the study of space.

A plasma magnetohydrodynamic ERE was first practically used
in the orientation system on the Zond-2 Soviet automatic space
probe, launched toward Mars on 30 November 1964.

An electrostatic rocket engine wa e first tested in orbital
flight in 1964 on the Voskhod spacecraft. The first test of an
ion engine in flight over a ballistic trajectory was conducted
in the USA on 20 May 1364.

In October of 1966, the Yantar'-l automatic ionospheric
laboratory, equipped with an experimental plasma™ ion ERE, was
launched in order to study the interaction of the reaction jet
of an ERE with the ionospheric plasma. The device was carried
on a geophysical rocket.

Thus, the developments on ERE performed in the late 20' s and
early 30' s at GDL properly predicted the development of technology
and preceded the actual demands of practice in this area by
approximately three decades.

Selection of Fuel for LRE

The most important problem to be solved in the creation of
LRE is the provision of high energy characteristics of the
engine. It is therefore understandable that the first stage
of investigations at GDL was the study of fuels, the investiga-
tion of various mixture formation plans.

An experimental rocket motor, the ORM, was created for pre-
liminary evaluation of the conditions of ignition and combustion
of liquid fuels in a chamber with a nozzle. The ORM consisted
of a thick-wall body, a nozzle cover and a screw plug; a membrane
was installed beneath the nozzle nut. A safety valve and crusher
device were attached to the body, fixing the maximum pressure
during the period of combustion. Two contacts screwed into
threaded holes in the ady carried a pyrotechnical igniter. This
experimental model (i.e., the ORM) was developed to test prepared
mixtures of oxidizer and fuel and underwent test stand testing in /91
1931. The charge of fuel to be tested (a mixture of benzene or
toluene or gasoline with nitrogen tetroxide) was placed in the com-
bustion chamber. When current was applied, the pyrotechnical
composition ignited, igniting the fuel. The pressure of the
burning fuel gases ruptured the calibrated membrane, and the
combustion products flowed out of the nozzle. If there was no

80

membrane, the chamber was freely connected to the surrounding
medium, the beginning of fuel ignition occurred in a semi -open
volume under atmospheric pressure.

Test., were
performed with noi-
ses from 3.4 to 12 mm
in diameter. in 1931 ,
46 flame tests of the
ORM wen performed,
the results of which
were used to eval-
uate the suitability
of various mixture
formation plans for
use in LRU. When
prepared mi.xtui cs
wer" used, explosions
wer. frequent; there-
fore, separate feed
of two -component fuels
was subsequently used
at GDI .

In 1930, GDI
first suggested
highly concentrated
nitric acid, its solu-
tions with nitrogen
tctroxide, hydrogen
pero.vide, perchloric
acid, tctranitro-
methano and other
substances as oxid.-
-crs, and beryllium
and a mixture of beryllium with hydrogen as fuels. Flame tests
were performed with nitric acid, solutions of nitrogen tctroxide
in nitric acid and liquid oxygen.

Cross-Section of Experimental Rocket
Motor (ORM)

Chemical ignition was first suggestedin 1931, then used in
a lumber of models of the ORM. iiyporgolie fuels were studied,
their corrosi veness for virions structural materials was tested,
the best methods for production of nitric acid oxidizers were
determined and experimental production of nitrogen tetrojfi.de
was undertaken to support the laboratory and test stand opera-
tion 1 ....

Due to the difficulty of producing nitrogen tctroxide in

large quantities, the most promising fuel type was found to be

highly concentrated nitric acid and kerosene. These components

are produced industrially, are not. explosive, can be stored for

/1A

81

extended periods and, consequently, allow rockets to be filled
long before launching. Nitric acid fuel provided good motor
operating stability and hhji specific impulse for the tin**

Engines with Annular Combustion Chambers

Thi., series of rocket engines includes the ORM-1, OHM- 2,
ORM-3, ORM-0-0, ORM-b and ORM-7, the combustion chambers of
which were annular in shape. They were created to study pro-
cesses inside the chamber and the regularities of changes in
thrust and specific impulse upon combustion of fuels in an
annular chamber. The annular shape of the chamber was selected /93
partially because it was found to be convenient to maintain
constant pressure as the thrust was cho^-d and the fiel com-
ponents sprayed into the chamber mixed better due to the elon-
gation of their path and rotation in the chamber by l o o .

ORM-1 was developed and constructed in 1^30-31. The engine
was designed for test i^tand studies of processes within the
chamber and was intended to be used repeatedly for short
periods of time. The ba^ic fuel called for by the plan was
nitrogen tetroxide and toluene, although the motor was tested
with liquid oxy«en and gasoline and developed a thrust of up
to IC kg.

The ORM-1 combustion chamber consisted of a cylindrical
steel body covered with copper; the mixing head of the chamber,
also clad with copper, was ma u e in one piece >»ith the outer
body .

The mixing head, also called the sprayer head, vas the
device used for mixture formation, i.e., spraying, atomizing
and mixing of the fuel components in the combustion chamber.
The design of the mixer went far to determine the mialitv of
the engine -- the completeness of combustion, stability of the
combustion process, reproducibility and stability of processes
within the chamber, etc.

The end cover was attached to th. outer body bv threading
on the opposite side from the head. The point where com-
bustion chamber and cover were joined was sealed by m ..- cJ a
circular knife seal .

A copper-covered steel nos:,le was also threaded to the
inner cup. In place of the ordinary cor, ical supersonic nozzle,
this nozzle had a rather long cylindrical port i.er with the
diameter of the critical cross section. Thir d^i s n was justi-
fied by the fact that the supeisonic portion o*' [lie nozzle is
not required to stuuv processes within the cha foe**, since the

82

supersonic stream Je v eIoping in the expanding portion of the

nozzlv does not perturb the subsonic stream in the constricting

,„,/-• - • : . - ,1 •■-• . •■ 10 1 inf] -
'.-• . st,* a within the sapbei Isc manul ctim s>f the gine
is s • • • " '• '• "• of thj scp nding po tioi oi ;>c
nozzle.

Si je [>e oae-pi? :pi n •■ • " fc ' •• -- - - »r,

! .♦-.-- the fuel) we we Idee it< ir< paced a l su \d

the head of the combustion chamber on a single circle. Bail
bad i ? es with screen filters ei <J id j :•> foi •: ?

spi 1 "- r - . ■ coj f.er s - m "... •■ o f re •? ■ t r; we v gaJ i?ani lly

coated with gold to make them corrosion resistant. During
tesi bt otor i i ubn ged water h'hii filled 1 : *. -

cool i r g ja :I at .

The fuel in. lifts
: 'liM - 1 was i ; ;?i i ted hv
tr- ■ I ' site c, : ui r : i ng
col ' c n, m t ci 'ih
all -. ' ■•: co : ■ i

wai> pj < ed n - h< :om-
bus - Lor hambej before
ear ; . te t, then
igt- ited , ; -. . kf< d

fuse. T $e ombt >t :'•; a
pre !uc . • - he main
fiw • . .'. ra d i -. t ic
am ulai sf i e • h :
ch; . ; v.. . ' -; • •- .. . . rd
the i c : i p] n ed i :
th< cent ci ■ > - he
eng ine chain Lrij . • ir
dii . ion oi • * - in
by 180°,

Cue! : >mpo-
rten ■ »*ere can c< . n
thi* -wa tan* and
f o i e J '.-is h e . 3in -
bui; •' : »■ • "by
cor ,,-«_•-* ;as p sure

A fui t lea level -
opr.*. of ' - esign ol

; ,. - jrjgiru •, the

ORM-: ( I '.131 I . 'ii con-
-. n i • • i OS 1-1 si it-tyj sprayers £ =•>' : ; i p *• :. i e i c

sed '- • * • • .- > - * . •"; ... -- s -■"•.' ' ! »-
s • , • -. • ' s.ber wa * n - re . ' - > :* Lntroduc ion of iyn: Lc

/94

/95

Cross-Section of ORM-1 Liquid-Fueled
/■/, kc ,; I :•■ : r»e

P3

cooling, and the design of certain individual parts and sec-
tions was simplified. The ORM-2 partially used the same type
of cooling as ORM-1, i.e., capacitive cooling (liberation of
heat into naturally circulating liquid surrounding the engine) ,
so-called static cooling.

By the time ORM-2 was finished, new and better designs
had been developed, and so ORM-2 was never tested.

The ORM-1 and ORM-2 engines were designed for element-
by-element testing of some of the main ideas upon which the
ORM-3 engine was based. This engine called for maintenance
of constant pressure in the combustion chamber with changing
thrust, an exponential nozzle, intensive (dynamic) cooling of
the combustion chamber by fuel, heat insulation of the com-
bustion chamber on the inside, slit-type sprayers and chemical
self-ignition.

The exponential nozzle developed at GDL is a profiled
nozzle in which the inner surfaces are given the proper geo-
metric shape to assure optimal flow characteristics of the
combustion products. The best contour of the nozzle is that
which achieves the extreme of specific impulse.

The methods of calculation of the nozzle were first
published in the USSR in 1957 by Yu. D. Shmyglevskiy and L. Ye.
Sternin. Simplified profiling methods are frequently used --
the nozzle contour is a circular arc, parabola, exponential
curve, etc.

The ORM-3 engine used hypergolic fuel, eliminating the
need for special ignition devices. Constancy of pressure in
chamber 1 was achieved by moving nozzle 5, sealed around two
belts 3 with a hydraulic or pneumatic device. As the nozzle
moved, the critical cross section changed, since the relative
position of the profiled projection at the center of the head
of the chamber which entered the nozzle was changed. In 1930- /96
1931, experimental and design work was continued on the
development of individual elements of this engine, in particular
using the ORM-1 engine.

The ORM-6 and ORM-7 engines were cooled by the fuel
components, had jet- type sprayers and represented further
development of the design of annular (slit-type) combustion
chambers. They were developed and produced in 1932.

The ORM-3, ORM-6 and ORM-7 engines were not tested, since
by that time the data from testing of the ORM-1 indicated that
annular combust i in chambers were undesirable, as was later
confirmed. Actually, the ratio of heated surface ^walls) to
volume where combustion occurs is greater in an annular

84

combustion chamber than in a cylindrical chamber; during combus
tion, the combustion products change their direction of motion
by 180°, which does not occur in cylindrical chambers. Both
of these factors cause overheating of the walls, particularly
the end portion, and complicate cooling conditions. The most
significant difficulty is in the organization of processes
within the chamber.

/97

Diagram of Regulation
ORM-3 Engine: 1, Comb
tion Chamber; 2, Cool
Fluid; 3, Sealing Bel
4, Controlling Gas or
Liquid; 5, Nozzle

development of design
individual processes.

Furthermore, the creation of
an engine with constant pressure
in the chamber but variable thrust
was found to be an independent
problem of some difficulty.

Experiments conducted in 1929-
1931 confirmed the possibility of
creation of reliable LRE. However,
it was also quite obvious that an
engine of constant thrust should be
created first, requiring that a
multitude of new problems be
solved; they included organization
of high quality mixture formation,
provision of complete fuel coniuas-
tion, assurance of high specific
impulse, organization of reliable
chamber wall cooling, etc.
Therefore, the program of further
studies called for stage-by-stage
elements and orderly, deep study of the

of
us-
ing
t;

Engines with Radially Placed Nozzles

This group of engines includes the ORM-4, ORM-5, ORM-8,
ORM-10 and CRM-13, developed in 1932.

These engines were created to study the processes of mixing
of fuel components, ignition, starting and shutdown. In order
to simplify the design of the combustion chamber and test stand,
to test the engine in the position with the head upward, the
nozzle was made in the for.ii of two radially placed apertures
opposite each other in the lower portion of the combustion
chamber wall.

The engines used pyrotechnic or electric-spark ignition,
with two spark plugs with massive copper electrodes installed
to i , >ease the reliability of ignition. In the first three
modt ; o of thi- group of engines, the fuel components were fed
directly into the combustion chamber, where they were mixed.

85

-3

'! i-\

The ORM-6 Engine

The ORM-4, ORM-5 and ORM-8
engines differed from each other in
sprayer design, with the sprayers
located on the hsad of the cylindrical
combustion chamber: the ORM-4 engine
had slit-type sprayers, ORM-5 was
equipped with jet-slit sprayers with
intersecting streams, while in ORM-8
the components were fed in through
jet-type sprayers, also with inter-
secting streams. In all three
models, the thick-wall steel body
of the cylindrical chamber was
attached by means of a threaded
joint at its end to the plate of the
test stand. The internal diameter
of the combustion chamber of these
engines was 40 mm.

These engines underwent flame
testing in 1932. Liquid oxygen,
liquid air, nitric acid, nitrogen
tetroxide and solutions of nitrogen
tetroxide in nitric acid were used

as oxidizers; gasoline, a mixture of gasoline with benzene and
toluene were rsed as fuels.

/98

T~~^ — T

The ORM-10 Engine

Electric spark ignition was
found to be unreliable. Metal-
nitrate caps were developed to
assure reliable ignition of fuels
with high-boiling oxidizers, while
trotyl pyroxyl caps, electrically
ignited, were used for fuels with
cryogenic oxidizers.

These tests yielded valuable
material on problems of safe starting
and stopping of engines, reliable
ignition and start-up when operating
with various fuels. The data from
testing of the ORM-4, ORM-5 and
ORM-8 engines allowed a comparative
evaluation of the quality of engines
equipped with jet and slit sprayers.

Jn the basic operating mode, the
pressure in the chamber reached
several atmospheres, the operating
time -- some tens c" seconds. During

86

individual, brief tests, the pressure in the chamber reached
50 a tin.

In order to study the possibility of high quality mixing
of fuel components in the liquid phase before they were fed to
the combustion chamber and atomized, thus increasing the com- /99
pleteness of combustion and the thrust per liter*, the ORM-10
and ORM-13 engines were planned with prechambers.

The fore-chamber or prechamber was a small chamber in which
preliminary mixture formation and partial combustion of the
fuel components occurred, after which the components were
delivered to the main chamber, where combustion was
completed.

In the ORM-10, the prechamber was made in the form of an
axisymmetrical channel; sections of identical length but dif-
ferent diameter alternated along the length of this chamber,
to improve mixing of the components which were fed in through
sprayers, the internal cavities of uhich contained spiral
snakes to spin the stream of liquid flowing from the sprayer.
The combustion products flowed out through two oppositely placed
radial apertures.

In ORM-13, the fuel components were fed in through concen-
trically placed slit sprayers into an annular prechamber. After
mixing, they were then sent to the spherical portion of the
prechamber and then, through the expanding portion, into the main
combustion chamber, which was cylindrical.

In one version, extra-rich or extra-pure liquid fuel mix-
ture of oxidizer and fuel, incapable of exploding due to its
composition, was fed in through one sprayer, while the other
sprayer supplied the remaining component required for complete
combustion.

The difficulty of manufacturing engines with three cham-
bers, the possibility of overheating of the heads and explosion
during start-up, led to new design solutions and stopped the
manufacture and testing of the ORM-10 and ORM-13 engines.

However, as we know, prechambers did come to be used in
certain engines produced in the first few years after the war,
particularly in the engines of the V2A and V5V geophysical
rockets. This resulted from the achievement of positive results

The thrust per liter refers to the ratio of thrust developed
by the engine in kg to the volume of the combustion chamber in
liters.

87

in a series c . •• si pei rmed ith p hashers in the

late 50*; •• id eai 1; 10" s.

hngiiu hit) Interna] roi ctivc Coat lags

/100

Tin t c .... • ■ 3f the pre
i in J ne* me ods to increas
It was -. und ■ ha this coul
surf;" :e • • lh :hamhei with
refr, tor] . tsul , • •

foum I i th< cyl indr il
istio! iaml r as re -
the c iamb* - c ! • •, . ad si
expei met ■ ■ • '• " .' to me
■•• ■ . let hods of >;niti

;r of l scs ot sprayers

cedin •■ ..•> ind

e the

d be do r by coa

cuprite or appl
cover in i . Furtb
sh *oe is the mos
ds orj :at i on
nip! . •■ i t> o .* prod
asurc • rest and
on an I ■■ l ., i t-up

were c mpa red t

icate i i he a ed to
t ime o f ; n -. - ! -- inc.
ting the im er
ieat i >.-: i t" a
erinore, it was
t favc ran le i o r a
ot" p roc i : -e; within
uct ion , Fui t her

dete_ . - < -, ;•-■* •; i fie
were hsptc \ - i d; a
o sel> ! " . best.

Th - • e cai I • - - ;« ting o orab i ■ : * fibers

with • : "i srsonic izzlt - ?ec it t»a: known that i no:: can
mere i : ' th : s i ai ., c m-, u ti> .-•• - i • ; . imj . Only

tests ol . ■•' •" • hasubers in coral lation u i t h n • . tould
be used full • •,>-. tp t\ '• engine a? hole

and ol Individual unit i presence of th 10 ~; hould

influi ce the eled no - « < :' 3d of ignit ion ■ i h< fin ;
and 1 Ik n ot o : . ; ti i t up mode .

/101

The ORM- ,• Engine

The ORM-12 Engine

88

The ORM-9 engine had a combustion chamber with an internal
diameter and height of 90 mm, covered on the inside with a layer
of ceramic heat insulation 10 mm thick of zirconium dioxide or
magnesium dioxide mixed with binder materials. The nozzle of
the engine, located in the flat cover, was clad with a layer of
cuprite 8 mm thick; its critical cross section was 15 mm in dia-
meter. The entry to the nozzle was rounded, the exit plane
was of the critical cross section. The two- component sprayers
were located in the head of the combustion chamber. Fuel
(gasoline) entered through a center channel with several out-
put apertures, while the oxidizer (liquid oxygen) entered
through a multiple-jet sprayer, the channels of which were
located around the central channel, parallel to its axis. The
output apertures of the central channel were tilted to make the
streams of fuel components intersect. The engine was placed
in a steel cup in the test stand and tested with nozzle upward.
Several firings of some tens of seconds each were performed in
1932. One of these was visited by Professor V. P. Vetchinkin /102
(TsAGI), who concluded: "The most important part of the work for
the manufacture of a rocket -- the production of a liquid- fueled
reaction motor -- has been performed at GDL... From this
standpoint, the achievements of the GDL (primarily of Engineer
V. P. Glushko) must be considered outstanding.'* 1

In ORM-11, the chamber and nozzle were also clad with
cuprite. The sprayers were also two-component jet type sprayers
with concentric placement of the fuel-feeding channels. They
provided fine, even atomization of the fuel; two-component
sprayers were found to be the best and are successfully used
in a number of LRE designs to ay.

The ORM-12 engine had the same dimensions as the ORM-9.
The chamber and nozzle in this engine were also clad with
cuprite, but the fuel components entered the combustion chamber
through individual snake sprayers located opposite each other
approximately at the middle cross section of the chamber. Back
valves were placed before the entry to the sprayers. The
ORM-11 and ORM-12 engines were tested on oxygen- gasoline and
nitric acid-kerosene fuels.

The ORM-14 and ORM-15 engines were planned but not manu-
factured, since their design, similar to certain foreign
models, was considered to be clearly unpromising. The primary
shortcoming of the engines was the fact that the fuel components
were fed into the combustion chamber from the direction of the
nozzle rather than toward the nozzle as is usually done.

"Cosmonautics," Moscow, The Soviet Encyclopedia , 1970, p. 93.

89

The next model was the ORM-16 engine. It has a supersonic
conical nozzle. The fuel entered the chamber through an
improved centrifugal sprayer. ORM-16 underwent flame testing
in 1932.

The ORM-17-ORM-21 engines, developed in 1932 on the model
of the ORM-16, differed only in length of cylindrical portion of
the combustion chamber and were designed to study the influence
of chamber volume on processes within the chamber.

The ORM-23 engines with two centrifugal sprayers, the
delivery of which wa.« regulated by a hydraulically moved
needle, had a combustion chamber placed between the sprayers
and could be repeatedly started. An air-gasoline mixture was /103
fed to the chamber and ignited by two spark plugs. This engine
was successfully tested with nitric acid fuel in early 1933.

The centrifugal
sprayer, first used by
GDL in rocket engine con-
struction, allowed a sig-
nificant improvement in
the quality of LRE and
practically almost cor-
pletely solved the
problem of preparation
of the fuel for complete
combustion. In
centrifugal sprayers,
the fuel components,
fed under pressure, are
twisted as they pass
through a nonmoving
multipass spiral in the
inner cavity of the
sprayer or by tangential
injection of the liquid
into the inner cylindrical
cavity of the sprayer.
As they fly from the sprayer into the combustion chamber, the
components form a so-called atomization cone, consisting of a
thin film which rapidly breaks down into tiny drops of various
diameters. f'".is new sprayer, used on the ORM-12 and ORM-16,
assured f ii •■ . tomization of the components and good mixing and,
as a result, complete combustion of the fuel. Due to this
property, spiral sprayers later became widely used and were /104
firmly fixed in domestic rocket engine construction.

At the same time, it was established that even when ceramic
heat insulation is used, the operating time of a rocket engine
is quite limited, and that it is more promising to use copper

The ORM-16 Engine

90

alloys with good heat conductivity for the manufacture of
nozzles* particularly in the area of the critical cross sect: t,

However* in either case unchanging design temperature of
the chamber wall can be achieved only if a portion of the heat
is carried away from the outer surface of the wall. Therefore,
studies of combustion chambers and nozzles cooled from without
were planned.

Engines with External Cooling

A. Air Cooling

This series includes the first LRE beginning with the
ORM-24 developed and tested in 1933. Experiments with preceding
engine models confirmed the need to equip the LRE with a cooling
system which would carry the heat away from the walls of the
chamber continuously during its entire operating time to provide
stable thermal conditions for the engine.

/IPS

The ORM-24 Engine

The ORM-26 Engine

At first, attempts were made to cool the engine with an
air stream. Therefore, the ORM-24, ORM-25, ORM-26 and ORM-30
engines were made with air-cooled nozzles. The chamber of the

91

ORM-24, like the ORM-16, was cylindrical in shape; the subsonic
portion of the nozzle was conical and ended in a flat nozzle of
critical diameter. The upper portion of the nozzle carried a
ribbed cuprite radiator. Spiral sprayers with ball back valves
were used to feed the fuel components. At the center of the
head was a device to detei^iine the maximum pressure in the
combustion chamber.

ORM-26 had a shaped nozzle with a well -developed super-
sonic portion and longitudinal external fins to cool the air
stream drawn by the gas stream of the operating engine. The
cooling fins encompassed both the subsonic and the supersonic
portions of the nozzle. The ORM-29 and ORM-30 engines had
massive, short nozzles with air cooling. In ORM-30, the inner
surface of the nozzle was not coated and was protected from
rupture by a film along the wall created by additional fuel
sprayers installed at the entry to the nozzle. This method of
heat protection of the nozzle walls was found to be effective
and has been widely used in practice.

Tests of the ORM-24,
0RM-2S, ORM-26, ORM-29 and
ORM-30 engines showed that
air cooling could not pro-
vide for long-term operation
of noz-les.

B. Liquid Cooling

An external dynamic
liquid cooling system is
capable not only of assuring
reliable operation of the
engine, but also of improving
the conditions of processes
within the chamber due to the
heating of one of the fuel
components in the cooling
cavity.

The first representatives
of sue. 1 ; engines -- ORM-2 with
fluid cooling of the head by
fuel and ORM-3, ORM-6-0,
ORM-? and ORM-7 -- had
practically complete cooling
by the oxidizer and fuel.
Due to the difficulty of manufacture of engines with fluid cool-
ing and the necessity in the first stage of repeated short-term

/106

The ORM-30 Engine

/107

92

start-ups to develop fuel spraying and ignition systems, start-
up and shut-down modes, the development of a reliable cooling
system for long-term operation was delayed to the second stage.
Element -by- element development of engines accelerated its
creation.

ORM-27 is also a fully cooled engine. The nozzle of
ORM-27 had longitudinal finning; the combustion chamber had
external fluid cooling. The internal wall was made massive
and had an elongation temperature compensator.

Beginning with model ORM-34, all nozzles of engines devel-
oped had flowing fluid cooling. In ORM-34, the region of the
critical cross section of the nozzle was cooled by liquid flow-
ing through a line at insufficient speed. In order to improve
cooling, the contour of the fluid-carrying portion of ORM-35
was somewhat improved, and the speed of the liquid was increased.
The nozzle of ORM-39 had an initial section with transverse
finning, cooled by liquid. The fully nitric-acid-cooled nozzle
of ORM-40 was found to be more stable in tests. In ORM-40,
the cooling fluid flowed in a spiral pattern through a thin
cooling jacket over the ribbed nozzle wall. Heat transfer from
wall to cooling fluid was increased by further increasing the
flow speed and its turbulization, a result of the ribbing in
the flow line.

As the design of ORM- series engines improved, the pressure
in the combustion chamber and specific impulse increased, and it
became possible to increase the operating time and thrust of
the engines. For example, ORM-39 and ORM-40 developed thrusts
of 100-150 kg. The critical cross section of the nozzles of
these engines were 25 mm in diameter, the pressure in the com-
bustion chamber reached 20-25 atm.

The nozzle of ORM- 4 4 and
all subsequent engines had
spiral ribbing, washed with
nitric acid. In these
designs, in order to give
the fluid-carrying portion
the necessary shape, a split
aluminum insert was installed.
A gap was formed between the
outer surface of the nozzle
wall and the inner surface of
the insert, through which the
cooling fluid flowed. The
diameter of the critical
cross section of the ORM-44 nozzle was 32 mm. The engine
developed a thrust of 250 kg. The ORM- 4 5 and ORM- 46 engines,

/108

Diagram of External Liquid
Cooling of ORM-44 Nozzle

93

designed for the same thrust, were sealed by the temperature
expansion of the nozzle.

The combustion chambers of all the engines mentioned from
ORM-34 to ORM-46 were cylindrical in shape with an internal
diameter of 120 mm and were cooled from without by the fuel
components, fed by centrifugal spray pumrs.

The ORM-47 engine utilized four supercritical mechanically
controlled centrifugal sprayers with back valves and filters.
Studies performed with ORM-48 allowed the concepts of the
nature of the distribution of pressure over the length of
the nozzle to be refined. The experimental installation on
which this engine was tested was simple and quite convenient;
these installations were later widely used in scientific
research organizations and educational institutions.

The ORM-49 engine had centrifugal sprayers with plate-type /109
back valves. In order to assure soft start-up of the engine,
some of the output apertures of the sprayers were sealed with
low-melting Wood's alloy.

Ignition in the ORM-24 and immediately subsequent engines
was by 7- second metal -nitrate pyrotechnic caps, suitable for
all oxidizers. Furthermore, in 1933 5 -second chlorate caps
40 mm in diameter and height, consisting of 501 Berthollet's
salt and 501 sugar were developed, which left no residue upon
combustion and were also suitable for all oxidizers. The
chlorate caps were also suitable for chemical ignition, since
they ignite spontaneously upon contact with nitric acid. These
caps were used in 1933 in a number of ORM-series engines for
chemical ignition by early oxidizer feed upon engine svart-up.
Starting in mid-1933 (ORM-44, ORM-50, etc.), chemical ignition
was provided using a start-up fuel developed at GDL which
ignited spontaneously when mixed with nitric acid. This
fuel included a solution of phosphorus in a mixture of carbon
disulfide and turpentine. The hypergolic fuel was first
carried in a starting tank on the muin fuel line near the
entry to the combustion chamber; -ater, it was supplied only
throtigh the lines feeding the kerosene sprayers.

Chemical ignition, developed and 1. . used at GDL, later
became common in rocket engine construct .on.

GDL Engines for Flight Vehicles

The result of the scientific research and experimental
design development at GDL prior to 1933 was the creation of the
ORM-50, ORM-51 and ORM-52 rocket engines.

94

; ' *• • d scribing h< ope t ton - id ' *•

let u recall i course of planning ., • n and
work --,.- -• -.:■ i - the - '- i Lou o ei gin • . t GDL
cribt- :• he p ecedinj sections. the wo k was
with tl Ir < Lo| ne t of eqi i] nen , r - i I ;1 stan
the basic .'.-,,<.-< t risti is of LRI , f] is was fol
tion of fuel :on ■'■* sm ' to- t stand tudi *> and
basi*. •• ..,.-," . . ii ngi te th .. r rate : eed o
mi •'■• $1 s \fi ?r . : !.'. ■• -■■:•■:- te < nt, ft

. •" (iopmont •-•• LRI Followei Our attenti n is
logic,- • jquen c of Che ■ •. ige m« "■, , ment
the h\ ad range pi \p< die y of he >] ction
invo. ed ir e cr . ? oi * gin« ■ : ,■ lis cla

al »o no i extensi : experimental j nduct

the t< ". anc :ould i t oi •• • a ; . •< to th

hich arose u he ie; - ;r c E first LRI

of t .- • - ' RE,
exf • - ent i
» brief 1 des-
- • on in 929
d mca ;ui £ isent of
Lowed b • : lec-
; c it 1 '-.■:■ j f a
f the I uel
age-b: • * ago
.--,'•• the
of de, Lgns,
of pr-. bloi -
ss. W »h aid
ed, i in : nly
e man ■ |uc: I • 31

/110

I

i

■^fenWB

Cros: tic a " - ; •, j

cu lit ions depondo
thermal constants
products and did
:ulat ions w. \

engines which he

el ■• idol £ • wa;

thci - 1 vcre <

Glushko in 1933-1

two series of stu

• : were ubJ h-d

Th • once .f c :'
: - ' •- nra -.■-.--. 1 1

d ..>:. t

of th
not re
r forme

develo

utl me

954 1 1

dents

he re
e f.-e
suit
d by
pod,
ct of
d in
the
speci
art ic

: he vt v .- Loj . ■: nt

of si] ,""•-• began ri th
ca I culal v • .
use .■'' the I -•;- of
the 4 :., m •■■•'-• -
cl , :• ; it • ; th >rmc
d] ■ im ics ■ i loi ed I he
th nal h < • , • ■
of fuel- :.-• a ■ t , ; «; the
; < .. c\ iracteri 1 ics
of i • : eng ■'-',-.- •■ -*i as
tlv u t , ^ress 1 r ? ii .' .
comb is t ion ;) - ie "and
al mg I he nc : .. , uel
co'v : nc iit ;-•/-,.- nipt ion
as ' un . • "- •• . sn * i ;:•
officii lime tons

of cri t ic .;. ; ind en ; t
p] s n , r ;s ses . i •-. of
th-- n »2zl« : be
determined. The
:; cct - '-. '. -.- ol the? 1 • al-
ii bi i ty >f d te m , 1 j 1 oi the
1 cc-!- pon< •■> ; ; ■ * / ' . bust ion
in si ;■■ - i leant >. r ro 1 - 1 hese cal »
V. P. • .' -:. :i 929 li i --or the
- ••• ,*" em .nt .cal : a
s i « • •. i e a 11 1 cm - * • , r h the s

•*-' - :ture? 1 a\ J by V. P.
. /. ikovskj , Lit ry ; • .- * ul •-•.- * to
alising in rocket technology, and

i vs ■■ *': 1- :>'. k; .1

111

Rocke-
eted Works,

■ si ■■ •'.. chinkin iu« ''- ; . \ Ko?

: ;:■:-. 01 , I 9' 2 .

95

The situation was quite different at that time as concerns
reliable calculation of liquid fueled rocket engine cooling
systems. In his report, "Heat Losses and Cooling of RM,"
published by V. P. Glushko on 2 July 1931 1 , the author pre-
sents the results of theoretical and experimental work on the
cooling of ORM performed at GDL up to that time. The report
presents a method for calculation of the cooling . f the ORM
combustion chamber, and the author notes that "the nature of
thermodynamic calculation of cooling of rocket engine combus-
tion chambers with liquid is well known. However, the
problem cannot be solved by theoretical calculations for a
single specific case in which heat transfer from the gas to
the internal wall occurs at the pressures and temperatures
which are found in rocket engine combustir chambers. Our
lack of knowledge of the heat transfer coeit icients makes the
cumbersome thermodynamic calculations useless and forces us to
turn to experimentation as the only satisfactory method for
solution of these problems."

Actually, we know that overestimation of the accuracy of
analytic calculation of cooling systems resulted in destruction
of both of the LRE of F. A. Tsander (OR-2 and 10) during their
first test stand operation (in 1933).

The cooling system developed at GDL by the experimental
method allowed engines to be operated repeatedly.

The ORM- 50 engine was developed at GDL for the 05 rocket >
planned and built at GIRD.

The ORM- 50 engine burned nitric acid and kerosene, had
a relatively short, spirally finned, oxid. zer-cooled nozzle
producing a gas pressure at the exit plane of 1 atm. The /111
liquid oxidizer heated in the cooling jacket was fed to two
spiral sprayers, placed radially on the cylindrical portion
of the combustion chamber. The fuel entered the chambe-. , also
radially, through t«o centrifugal sprayers. All sprayers had
back valves. The middle cylindrical portion of th? combustion
chamber i ad no external liquid cooling, Lut was cooled by an
internal curtain; ignition was chemical.

Thu ORM- 50 engine, of widen a single model was built,
passed three reiinement, operating life and acceptance stand
tests i.n 1933. Then, in 1934, five test launches of the 05 /113
rocket vere conducved, powered by this engine, to test the
fuel feed system.

The Pioneers of Rocket Technology. Vetchinkin, Glushko,
iCorolev, Tikhonravov, Selected .'orks, Moscow, x972, pp. 208'
212, 770.

96

Wht n '■■>-■' 05 ro< ket

was lai i t h id as : • e fir-
! "i' r • • * n vioscow, low

ressure J -on
the I auks a i • ■■■. the
eng *•- t d 2 eloi I ess
than full hn t (des i »r
: rust I5f I | and , he
ngi le 01 ,. a ed or

■ sc« ••. is ii the a inch
stand until the tanks
emptied, without lifting
the r. • "■ -■ ■ We si >u d
note particularly the
• ib •. • : of 1 he - n ; ine,
which •• • ■• r ed : ( •• tarts .

Th ' exj ;• i -.-ce
gained * '• - • lo] r g the
01 M-5C tfas ,: d : ! • te

• )i oi t isoi ■• wer-
■ • ;.. -'.el -- the i IM-51
engi te - desi| s .-• 3
develop 250 kg thrust.
In ' hi ORI 51 I i • Lne,

f uel eompo it ',-ere
de I . . ,-,. - I t< wo c i i ;ular
colle tors. Local rd
. :>< s te ' ->■ , -•- * :al
I ?ad : : th< c on bust ion
charnb* , . • ; .-
x lizei pa: ed "'-, sg he ik • • - trig .;et. Fr-. ' . i

;■, ", •• toi . :.- •■<:. '■ • :>.. ,,■ ■' fuc • e sent t« : ■, • sntrifugal
sprayers located around the base of the hemisphere arid directed
upwa rd al an ; i| ] e of 2 5°,

The ORM-50 on the Test Stand

F rther .-• t tie esea ! and »xperiti ntal
resulted in < a - QRM-! engim

earli ' odels, burned n tri< u . kei • ;.• fuel,
designed £o ' I .. Rl I , RJ a " ; -'■ < ;■ rod ts pi

GDL, n endi . ; t< be used boi ', - • roarin< :o : '-;
. o I )oos :e i . k *' " iii raft. 0£-'. -al te

• - - - > ■ - , ; i *. ■ • ; ; • ; ■ ' _ • tion

pressuie of 2-- .. atm . ievelope<; a n st of up
300 kg. An GRM-52 manufa ture mt t sted :-. the
1938 developed a thrust of 300-320 kg with a feed
35 atro, a pressure n '. l -., coml s s . m haml e of 20
specific impulse - L0 e< n is stil >perati
start > and a tal operating tin i ec.

development
h , 1 i •■. e
I ? was
anned at

and s a

oi the
• .,-' • , ?r
to 250-
stand in
pressure of

'la
ng after 29

97

In this engine » the steel cylindrical combustion chamber

(inside diameter 120 «») with spherical head had a conical

nozzle. The fuel components were fed in through 6 centrifugal
sprayers -- three for each component. Back valves were placed
before each sprayer. Ignition was chemical, using hypergolic
fuel consisting of the basic oxidizer, nitric acid, and a
starting fuel --an active liquid poured into the fuel line
from a feeding collector ring before the start.

/115

ZM1

The ORM-51 Engine

The combustion chamber had no external liquid cooling, but
was cooled by an internal curtain. The nozzle was cooled with
nitric acid, fed from the tank to a collector in the lower por-
tion of the nozzle cor*:ing jacket. The liquid flowed from
there through the gap between the jacket and the nozzle, then
flowed along the finned ozzle through a spiral channel and
exited through three connections, each of which was connected
to one of the sprayers. The nozzle was surrounded by a
properly shaped aluminum sleeve to provide the correct nozzle
shape an! size.

The ORM-52 was the best engine of the time as concerns its /116
basic characteristics -- thrust, specific impulse and operating
life.

98

Fuel

Fuel Feed Systems and Stands

Beginning in 1929, together with the search for efficient
combustion chamber designs, work was performed on the creation
of stand measurement and fuel component feed systems. In 1930,

based on analysis of
weights, it was estab-
lished that the most
efficient type of fuel
feed for low -thrust LRE
is an extractive (cylin-
der) system, using either
compressed gas from a
pressure accumulator or
liquefied gas evaporated
in an evaporator. It
was clear in the 30' s
that a pump feed system
was preferable for high-
thrust LRE. Let us
recall that K. E. Tsiol-
kovskiy planned this
type of fuel feed system
in his theoretical
studies.

The development of
compact turbine -pump
units and the applica-
0ff *J tion of the latest struc-

tural materials has
allowed pump feed systems
to be used not only in
large engines, but also

Cross Section of the ORM-52 Engine

in LRE producing relatively low thrust, in recent years.

At GDL, a feed system was developed both for flame test
stands and for engines installed in flight vehicles of
various types. In 1930-1932, LRE were tested at GDL on a stand
in which the fuel components were driven from their tanks by
compressed nitrogen. The test stand containers for oxidizer and
fuel were large-caliber artillery cartridges, lined on the
inside with aluminum if they were to contain nitric acid or
other corrosive fuel components.

The 20-liter liquid oxygen tank was placed in a sealed
brass jacket, made from the cartridge of a 12-inch shell; the
gap was filled with carbon dioxide and activated charcoal.
When the tank was filled with liquid oxygen, the gaseous
carbon dioxide was frozen, and the other gaseous products

99

present as impurities were absorbed by the charcoal, creating /117
a high vacuum to insulate the tank.

During 1931-1932, work was performed at GDL on a special
fuel feed system using piston pumps. In 1931, a fuel feed
system was developed using a piston unit consisting of four
double-acting piston pumps placed radially around the combustion
chamber. This pumping unit was planned for use with ORM-3.

In 1931, the ORM-A engine was constructed according to
a plan suggested by B. S. Petropavlovskiy. This engine had a
pumping unit driven by the combustion products; a charge of
smokeless trotyl pyroxylin powder was burned in the chamber
for the first few seconds in order to produce the products for
engine start.

In 1931-1932, a piston pumping unit was developed, manu-
factured and tested to feed a nitrogen tetroxide- toluene
engine with a thrust of 300 kg.

One common feature of pump fuel feed systems is the use of
a portion of the energy of the gases in the combustion chamber,
causing a certain increase in the efficiency of the entire
engine. However, these systems have been found practically
inconvenient, primarily due to the unevenness of fuel feed
during the course of one cycle. Therefore, in 1933 the devel-
opment was begun of a turbine-pump unit for a nitric acid-
kerosene engine developing 300 kg thrust with a fuel componet
feed pressure of up to 75 atm (shaft rotation speed 25,000 rpm) .
A design plan was selected for the TPU [turbine pump unit],
consisting of a gas turbine with one stage and two single-
stage centrifugal pumps (for oxidizer and fuel) seated on a
common horizontal shaft.

The vanes had bidirectional input to relieve the axial
forces. The body and vanes of the pump were made of an aluminum
alloy. The turbine was powered by the combustion products of
the fuel at a temperature of 500° C and a pressure of 15 atm.

During testing of an experimental model at a test stand
of the metal plant, a single-stage pump produced a guage
delivery pressure of 75 atm, which many had considered impos-
sible at the time. The gas turbine rotor was taken from a /119
supplementary marine engine. According to an air force order
(1932), this TPU was designed to be installed together with a
300 kg-thrust combustion chamber on the 1-4 aircraft.

100

A Diagram of the ORM-A

chute and an automatic device for e
after completion of the flight test
ponents were fed into the engine us
through the hollow gimbal rings, wh
journals around them. The lower po
carried a compressed gas cylinder,
the fuel tanks, while the nose port

The Rocket of GDL

In 1932, 1933, RLA
rockets were produced
at GDL for flight
testing of LRE -- the
RLA- 100, RLA-1, RLA- 2
and RLA- 3. They were
preceded by development
of individual RLA sys-
tems.

The basic design
parameters of the RLA-
100 rocket, the plan
for which was developed
in 1932, were as fol-
lows. Flying altitude --
up to 100 km, launch
mass -- 400 kg, fuel
mass -- 250 kg, engine
thrust -- 3000 kg,
pay load mass -- 20 kg,
operating time --20
sec. The rocket con-
sisted of two steel
bodies interconnected
by the nose portion.

Nitric-acid LRE 2
was installed above
the center of gravity
of the rocket on a
gimbal support, was
gyroscopically stabil-
ized and served not
only as the driving
power source, but also
as the actuating element
of the control system.
The nose portion of the
rocket 1 carried weather
instruments, a para-
jection of the instruments
program. The fuel com-
ing an extractive system
ich were sealed into the
rtion of the rocket body 4
the upper portion 3 carried
ion 1 carried the oxidizer

/118

101

tanks. The duralumin fins 5 in the tail portion of the rocket
assured that its center of lateral resistance was lower than the
center of gravity.

A test stand with a gimbal support was made to test the
engine and determine the stabilizing influence of the exhaust
stream. Working drawings of a motion picture camera with a
time recording system to be installed in the tail section of
the rocket in order to determine the trajectory of flight of
the rocket were developed. In 1932, three rocket bodies were
under construction at a machine building plant.

The RLA-1, RLA-2 and RLA-3 rockets were designed for
flight testing of LRE with up to 250 kg thrust. They were to /120
fly vertically to altitudes of 2 to 4 km. The design of these
rockets called for rigid mounting of the engines in the tail
portion of the rocket. The fuel feed system was extractive
using compressed gas from a pressure cylinder. The fuel tank
was located concentrically within the oxidizer tank. The
launch was to be vertical, without a guiding support, from a
launching stage.

The simplest design was that of the RLA-1 rocket, with the
ORM-47 LRE. The body of the rocket was steel, but its nose
portion and tail fins were made of wood. The extractive feed
system had no pressure reducer. The length of the rocket was
1880 mm, the diameter of the body -- 195 mm.

The RLA-2 rocket, like the RLA-1, was uncontrolled, but
differed from the RLA-1 in that it had a duraluminum nose
cone, in which were located a parachute and weather instru-
ments, and an automatic device for opening and ejection of
the parachute; furthermore, the central portion of the
rocket body carried an equipment section with a pressure
reducer, assuring even fuel feed to the combustion chamber;
the rocket had duraluminum tail fins. These rockets were manu-
factured in the shops of the National Mint. Preliminary test
stand operation of the RLA-2 rocket with the ORM-52 engine
(not shown on the figure) was conducted in 1933. /122

The RLA-3 rocket was a controlled rocket, and differed
from the RLA-2 in that the body contained an instrument
section with two gyroscopic devices with air pressure power
(gyroscopes from a marine torpedo were used) ; they controlled
two pairs of rudder fins at the tail of the rocket by means of
pneumatic servo drives and mechanical linkages.

Munk aerodynamic profiles were selected for the rudders,
providing the minimum displacement of the center of pressure
upon movement of the rudders. The RLA-3 was never completed.

102

In early 1934, the documenta-
tion and materials section of the
RLA project was transferred to
RNII, where a section for development
of liquid-fueled rockets was set up.
Since by this time RNII already had
an approved plan of operations, the
RLA rockets were never developed
further.

Thus, the basic result of the
scientific research and experimental -
design work performed at GDL in
1929-1933 wps deep and comprehensive
study of the processes occurring in
LRE, the development of good, eco-
nomical and reliable engines (for
the time) and the solution of a broad
range of problems related to rocket
engine contruction. Liquid rocket
fuels were developed and studied,
as well as methods of fuel feed to
the combustion chamber, conditions
of mixture formation and preparation
of fuels for combustion, and methods
and means were developed for pyro-
technical and chemical ignition in
engines, as well as the start-up
and shut-down of engines, processes
within the combustion chamber were
studied, methods of cooling of com-
bustion chambers were developed,
the conditions of flow of the com-
bustion products from nozzles of
various shapes were studied, and
factors influencing the thrust and
specific impulse were determined. Finally, GDL mastered
techniques of experimentation and operation of LRE, developed
test stand equipment and apparatus for recording of parameters
during testing and developed the design of engines developing
thrust up to 300 kg with specific impulses of up to 210 sec at
ground level with repeated start-up capability.

The viability of LRE was convincingly proven by extended,
reliable and economic operation of the ORM-50 and ORM-52. The
path was shown for further improvement of engines. The creation
of these models was of decisive significance for further develop-
ment of Soviet rocket engine construction.

Diagram of the RLA-100
Rocket

/124

103

/121

D ., ;ram of the RLA-1
Rocket

Diagram of the RLA-2
Rocket

104

'■»**J \
• ti.»l.» ► •* ..." i,„V . - - < -- * -.»• '

Jl -

W10'

■•... • -^b^^Mm -%^^-->y^^c ' .-. -, , >•

• " -^F Ti / tf '•'*!• '-'51 ■ '>"'

',- \Fj^jf^f.^t-^?.!5 - ' 'A,.-' - •

/123

J

Memorial Plaque Installed on the Building of
the Ioannovskiy Ravelina of Petropavlovskaya
Fortress. [Translation of Plaque:

In 1932-1933, here at Ioannovskaya Ravelina were located
the test stands and shops of the USSR's first experimental-design
organization for the development of rocket engines -- the Gas
Dynamics Laboratory (GDL) of the Military Scientific Research Com-
mittee of the Revolutionary Military Council, USSR. Here were con-
ducted test stand operation of the world's first electrothermal
rocket engine and the first Soviet liquid-fueled rocket engines,
developed by GDL in 1929-1933. GDL laid the foundation for domes-
tic rocket engine construction. The team which grew out of GDL, a
part of the twice awarded Experimental -Design Bureau, created the
powerful engines of the booster rockets which placed satellites in
orbit around the Earth, mo n and sun, sent automatic spacecraft to
the moon, Venus and Mars, and launched the manned spacecraft Vostok,
Voskhod and Soyuz.]

105

It has been 44 years since the subdivision for development
of ERE and LRE was created at GDL (1929-1933), beginning the
long and difficult path of development through subdivisions in
RNII (1934-1938) to the formation of the independent group
(1939-1940), which in 1941 was expanded into the Experimental
Design Bureau. This was the creative path of development from
GDL to Experimental Design Bureau of the organization called
GDL-OKB. The foundations of domestic rocket engine construction
were laid down at GDL. Most of the workers who held creative
positions in the twice-awarded Experimental Design Bureau GDL-
OKB, which created the powerful liquid-fueled rocket engines
for all Soviet booster rockets which have flown in space, came
from these walls.

V. P. Glushko, the great leader of GDL-OKB, was the
designer of these engines.

In celebration of the 40th anniversary of GDL-OKB (1929-
1969), memorial plaques were installed on the buildings of the
Main Admiralty and the Ioannovskaya Ravelina of Petropavlovsk
Fortress (Leningrad), where GDL was located in the 1930' s when
the ERE and LRE were invented.

2.4. The Moscow Group for the Study of Reaction Motion,
CS Osoaviikhim USSR (MosGIRD)

By the early 1930's, efficient forms of participation of
society in the solution of practical problems of astronautics
had been found. Party and state organizations provided great
aid to individual clubs and groups involved in the study of
reaction equipment.

A significant step in development of work on rocket tech-
nology in the USSR consisted of the organizational measures
performed by Osoaviakhim USSR, which cooperated greatly in the
development of new military technology.

From the very beginning of the activity of Osoaviakhiir,
its theme and structure included the conduct of scientific
research work, which was then broadly developed. In particular,
the Scientific Research Center of the CS Osoaviakhim included
the Bureau of Air Technology (BVT) , the task of which included
scientific research work and the development of new types of
flight vehicles. Design bureaus, shops and laboratories were /12S
set up for this purpose.

In particular, serious attention was given to the study of
problems of rocket technology, based on the works performed
since 1921 in the laboratory of N. I. Tikhomirov, and somewhat

106

later in the sections, clubs and societies of rocket technology
enthusiasts.

The first public group for the study of reaction moti n
began forming in Moscow in connection with the works of F. \.
Tsander, who was discussed above. In December of 1930, working
at TsIAM, F. A. Tsander attempted together with CS Osoaviak'.iim
to create such a group of rocket technology enthusiasts which
could solve independently the great scientific research prob-
lems and perform the necessary planning and experimental work.

On 18 July 1931, the first meeting of the new Osoaviakhim
organization, called the Bureau for the Study of Reaction
Motica (BIRD), was held, under the chairmanship of F. A.
Tsander. The plan for the work of BIRD called, in particular,
for organization of BIRD cells at enterprises, and a report by
F. A. Tsander at a general meeting of members of cells on the
conditions of interplanetary voyages.

Thus, BIRD, which later grew into GIRD, was a fully formed
organization by this date (18 July 1931).

The name GIRD is first encountered on 20 September 1941 in
a letter by one of its members, comrade Fortikov, to K. E.
Tsiolkovskiy, who was familiar with the practical and organiza-
tional affairs of GIRD.

According to another point of view, GIRD was founded on
18 August 1931, on the initiative of F. A. Tsander and
N. K. Fedorenkov, who spoke to Osoaviakhim USSR on the crea-
tion of an "Interplanetary Society." N. K. Fedorenkov
announced through the press late in 1930 and early in 1931 that
all those interested in problems of interplanetary voyages were
invited to joint together and wrote in a letter to Ya. I.
Perel'man that "the <jroup for the study of reaction motion"
was organized on 18 August 1931. This date is mentioned in
the article "The Rocket and its Development" (1935) .

Finally, a third point of view is defended by those who
consider the date of founding of GIRD to be the day of the
beginning of practical work on reaction equipment, namely 18
November 1931, when F. A. Tsander, who at that time headed the
study of reaction motion in Moscow, concluded a "socialist
agreement for strengthening the defense of the USSR" with the
Bureau of Air Technology of the Scientific Research Section of
CS Osoaviakhim for planning and development of working draw-
ings, manufacture and production of models of a reaction engine,
including installation of this LRE on an aircraft.

We note that it is this date, 18 November 1931, which was
selected by a group o r veterans of rocket technology of the

107

Soviet National Union of Historians of Natural Science and
Technology, Academy of Sciences USSR, to hold a creative meet- /126
ing dedicated to the 40th anniversary of the organization of
GIRD in Moscow. This disagreement in the determination of the
precise date of organization of GIRD is explained by the fact
that the group was created gradually, its organizational forms
changed, were improved and strengthened with each new step.

In 1932, CS Osoaviakhim adopted a resolution calling for
broad development of work in the area of aviation technology.
In particular, the Tsander group was encourated throughout
1932 to complete work on the creation of a reaction engine for
an aircraft. In June of 1932, the Praesidium of CS Osoaviakhim
adopted a resolution calling for the organization of an experi-
mental scientific research base (GIRD) , which was given the
task of planning, construction and testing of engines and
rockets of various types.

Thus, the group, which worked up to June of 1932 by popular
support, was converted to a scientific research and experimental -
design organization with its own staff and base. Financing was
both through Osoaviakhim and through the Administration for
Military Inventions (UVI) of the People's Commissariat for the
Navy.

In 1932, GIRD was given space for the creation of a
scientific research production design base beneath No. 19
Sadavo-Spaskaya Street in Moscow.

By July of 1932, the basic trends in the activity of GiRl*
and its structure had been determined. An order of CS Osoavia-
khim of 14 July 1932 names Sergey Pavlovich Korolev as the head
of GIRD, beginning 1 May 1932.

The structure of GIRD which had developed by mid- 1932
reflected the trends of its activity. Four interrelated
trends of work are characteristic:

-- scientific research and experimental work on the applica-
tion of reaction engines;

-- broad technical popularization of the application of
reaction engines;

-- training of workers in rocket technology;

-- leadership and coordination of the activity of the
GIRD created across the country*, allowing the Moscow group to
be called the central group (TsGIRD). 2

By this time, some 100 groups had been formed for the study of

reaction motion.
2
The name TsGIRD is first encountered in official documents on

31 March 1932.

108

The work of GIRD
the ' • hi :a mcil

the < Lef of i RD, S.

was headed by
, cha • , ' by
>lev.

/127

' he Chi : >i GIRD,
korolev (1933 Photo)

5erg< l iv lov ■ . -' o] -• wa • born
3C De -.'•:.> L! ( n Zhil Offl h , : ; .• son
of a tea I e n I' h .... di - ted

from loscow Highc • • ti - • > -• ->1
• , mm and at the same time

from . Moscow Pilot , ool . S. P.

Korolev creata ! n ml ;t . ; •,"■• gns
of glide '■ ■ - 5 fully.

. e : .- ,' Liariz ',-,".= ■•.,,' the

works of K. E <• ; , >lkovs* i> . te -is
attracted t tl p d Lit ; • iing
liqui i- fueled rockei ngines -r
aircra t, . • • de < • md o< c age-

ther with i • I sand e md < her
spec ial ists Li lit a rea o I :ket
techr ology :••• took pa? in t hi or *an-
ization of 6IED in CS Osoaviakhim USSR,

After isee g F. A ' J .'.-.• P - i i hi the

sciei : i • . dea of ridri h ■ " :ui • • , ed vvith Ji »wn

ideas to a reat jxtcnl '■ ■ ■• that F. A. Ts idei who sad

stud, . problem. c o , t onautic; foi many ■ . i • • , i I nu more
oxpe'* '.'•.' ■- •' ■ •■■', ?dg< . h< a i : a f - 3e • • : technolog I han
he. Serge) Pavlo 1 ; ;onsidered It te ; sai c hrin I c
lik crin d< opment of F, A. TSai • I " . oon as

tble S, P. KoroJ ' con red that in orde for tl* .deas

of reat tors motion to be • cep • i a i i tion flight v< . • - : e
would r , ; : te 5 • Ic »n, u tha tl . fould »qui e ;.hat the
v -2 ocket rig n< plan by instructed.

Und ; ' md Lng : ; ' n< h :'•-. n -
tal work, S. P. KoroJ ' iched gr< a t

Icvelopmeni ' ' - >; ;•' . 'opaj ar . • . He
seri .• popula - I u

I ms • : u riagec e gi vc t in e • his 1

in 9 .= Gosvoyenizt " Press
of S \ Korole v '■ 'Rocket Id g i

in which th need and ans or mastei
clearl> shown , . tions f high alt i
and tl c pe< ul iaritic >j ;rai with

desc : - -:■ , this - ■ '<• '.- "- "- '. • • Jescr
whic;- •- -'•;• ired u] ' f th< • me, at
of the theor) of *eaci m flight, mvl
formula! thrust : iausl -elocity

nif . • ' • . -e
reamed . ; ting a

sp ze ■ q pm< i ile

itera work, '*or
rintei i,00l ics

in t • S ratosph -c"
, • • . ?trat - • xt ■ - n

tude flight were studied
reat : on ens in -• were

•• - : oi i nui be r >i i Ri

kc 11 as cert aii ' em< > 1 >
iding - i] si • • lie

and ••■'•;. ! ■• cy.

/128

109

"From the shores of the universe, which our Earth has now
become," Sergey Pavlovich said, "Soviet ships will repeatedly
fly far into space, lifted by powerful rocket boosters. Each
flight and return will be a holiday for the Soviet people, for
all forward- thinking mankind --a victory of intelligence and
progress."

The outstanding organizational capabilities of Sergey
Pavlovich, the brilliant mind of this great scientist, allowed
him to solve a number of important problems of rocket construc-
tion. During the post war period, S. P. Korolev directed the
work of the design, scientific research organizations and
test firing ranges for many years.

In the history of the study and mastery of space, the name
of S. P. Korolev is connected to epochal achievements. The
scientific and technical ideas of Sergey Pavlovich have been
broadly realized in practice. Many ballistic and geophysical
rockets, booster rockets, manned spacecraft and automatic
interplanetary spacecraft (AIS) and artificial Earth satellites
were created under his leadership. Sergey Pavlovich Korolev
directed the launching of the world's first artificial Earth
satellite, created the space rocket systems used for the first
manned flight in space, the first flights of automatic space-
craft to the moon, Venus, Mars and the landing of an AIS on the
moon.

S. P. Korolev was made a Corresponding Member of the
Academy of Sciences USSR in 1953, an Academician in 1958.
Sergey Pavlovich Korolev, a CPSU member, is a twice Hero of
Socialist Labor and a Lenin Prize laureate.

The name of Korolev, one of the founders of astronautics,
has been given to the largest formation on the far side of the
moon.

GIRD consisted of four planning-design teams, combined
into section I, production shops and a test station (section IV),
an administrative division (II) and the organizational and
mass operations division (III). GIRD was subordinate to CS
Osoaviakhim. Sections I, II and IV were located in the base-
ment of No. 19 Sadovo-Spasskaya Street and were a secret enter-
prise; section III functioned as an open and somewhat inde-
pendent organization in Osoaviakhim.

The first team was headed by F. A. Tsander. The team
included L. K. Korneyev (who later became the team leader in
March 1933), A. I. Polyarnyy, L. S. Dushkin, A. V. Salikov,
S. S. Smirnov, V. V. Griyaznov, Ye. K. Moshkin, I. I. Khovanskiy, /130
N. M. Vever, L. I. Kolbasina and A. I. Podlipayev. This team

110

tested the OR-1, worked on the preparation of suspensions of
metal and kerosene, experiments on the ignition of metallic
fuel in air. A suspension of magnesium and kerosene was sug-
gested for the engines designed by F. A. Tsander as fuel.
The suspension was produced using ball mills, and also by means
of an electric arc. The OR- 2 engine was tested with liquid
oxygen and gasoline, the LRE 02 aviation-type engine was
planned and tested, burning liquid oxygen and ethyl alcohol
as well as the LRE 10, designed for the GIRD-X rocket.

Party
Organization

Chief of GIRD

Technical
Council

i

/129

X

Section I, Sci
entific Research
and Experimen-

iai

I

iBSfr.

I

Section II,
Administrative

B
A
V

H

I

section ill,
Organizational
and Mass
Operations

Section IV,
Production

E
A
V

T

to

a.
o

si

SI

"O

o

u
a.

in

C
«8

a>
H

S

Organizational Plan of GIRD

The second team was headed by M. K. Tikhonravov. It
included: V. A. Andreyev, V. N. Galkovskiy, Ya. A. Golyshev,
N. I. Yefremov, V. S. Suyev, Z. I. Kruglova, 0. K. Parovina,
Ye. I. Snegirev, V. A. Fedulov, N. I. Shul*gina and F. A.
Yakaytis.

Under the leadership and according to the plan of M. K.
Tikhonravov, the second team developed the GIRD- 09 rocket with
the 09 hybrid- fuel engine. The second team developed the 07
rocket, flight tested in 1935. This team attempted to create
an aviation engine with pump feed of liquid oxygen and gasoline.
Other developments were also conducted.

Ill

Mikhail Klavdiyevich Tikhonravov was born 29 July 1900. He
began his creative activity in 1923 when he was still a student
at the Military Air Academy imeni Zhukovskiy. After graduating
from the Academy in 1925, M. K. Tikhonravov was sent to work at
the Aircraft Design Bureau of N. N. Polikarpov. In 1930, M. K.
Tikhonravov was transferred to work at the Central Design Bureau
imeni Menzhinskiy, where he used his work on aircraft motor
equipment as a basis for his brochures "Aviation Tanks" (1934)
and "Aviation Motor Supply and Lubrication Systems" (1936) . In
1932, M. K. Tikhonravov, after meetings and discussions with
S. P. Korolev, was transferred to GIRD. At RNII, M. K. Tikhon-
ravov, together with a team from the Department of Wingless
Rockets, began the development of a rocket to carry man into
the stratosphere. Then M. K. Tikhonravov headed the Laboratory
of Alcohol -Oxygen LRE. As a result of his scientific stucies on
LRE, Tikhonravov published the articles "Use of Rockets for
Investigation of the Stratosphere" (1936) , "An Oxygen Rocket
Engine " (1937) , and "Principal Characteristics of a Rocket
Engine" (1938) in the collections Raketnoye Tekhnika [Rocket
Technology] and Raketnoye Dvizheniye [Rocket Motion] .

In 1938, M. K. Tikhonravov began to study the stability
of flight and reproducibility of trajectories of uncontrolled
solid-fueled rocket weapons. The results of his studies were
published in Raketnoye Tekhnika under the title "Study of
Factors Influencing Firing Accuracy of Rocket Shells."

When he was leading the work on the investigation cf /131
flight conditions of the artificial Earth satellite in 1950-
1951, M. K. Tikhonravov was one of the authors of "Principles
of the Theory of Flight and Elements of Planning of Artificial
Earth Satellites." M. K. Tikhonravov also wrote many other
works on problems of rocket technology.

The government of the USSR has evaluated the works of
Mikhail Klavdiyevich Tikhonravov highly, awarding him orders
of the Soviet Union, and giving him the Lenin Prize and the
rank of Hero of Socialist Labor.

In January 1970, Mikhail Klavdiyevich Tikhonravov was
selected a Corresponding Member of the International Academy
of Astronautics.

The third team, headed by Yu. A. Pobedonostsev, studied
and developed air-reaction engines.

Yu. Alekseyevich Pobedonostsev was born in 1907 and became
a Doctor of Technical Sciences and Professor. He participated
in the organization of GIRD. In 1932 he was transferred to GIRD
as a full-time worker, where he led the development of direct-
flow air-breathing reaction engines using solid fuel. Working

112

at RNII, he contributed to the crea-
tion of the Katyusha rocket launcher.
In 1968, he was selected as a Corres-
ponding Member of the International
Academy of Astronautics.

Mikhail Klavdiyevich
Tikhonravov (1971
Photo)

The third team successfully flight
tested models of direct flow air-
breathing reaction engines (DARL) .
The first domestic supersonic wind
tunnel, created with the participation
of M. S. Kisenko, an engineer in the /132
third team, allowed the production of
an open air stream from 40 to 60 mm
in diameter at a velocity of 480 to
900 m/sec; working at reduced pressures,
the gas stream could be increased to
1100 m/sec. The axisymmetrical
nozzles used to produce the supersonic
stream were designed by a method
suggested by Professor F. I. Frankel.

Winged rockets we
leader of which was S.
developed a glider for
flew, prepared for tes
GIRD RP-1. The rocket
weight without the LRE
arising in development
being flight tested wi

re developed by the fourth team, the first
P. Korolev. Designer B. I. Cheranovskiy
the OR -2 engine, which S. P. Korolev

ting as a rocket plane, later called the
plan had wind span of 12.1 m Its
was 200 kg. However, difficulties
of the OR- 2 prevented the RP-1 from

th the engine.

GIRD had experimental shops equipped with machine tools and
various specialized devices. The production process was
headed by P. S. Aleksandrov, I. A. Vorob'yev and Ye. M.
Matysik. Flame testing of engines and flight testing of
rockets were performed at the range in Nakhabino.

One important area of the activity of GIRD was propaganda
and popularization of reaction motion.

This area was headed by the third section of GIRD, the
organization and mass operations section. For reasons of
secrecy, section III was placed separately from the other
sections of GIRD in an open territory.

The work of section III involved not only the GIRD members,
but also people working with popular support, not included as
a part of the GIRD staff.

The activity of GIRD in the area of scientific and techni-
cal propaganda corresponded to the resolutions of the communist

113

party on problems of mastery of technology. We have in mind
here the resolution of the CC VKP(b), adopted in 1931-1932 and
designed to encourage broad development of technical propaganda
in which, in particular, the need for comprehensive encouragement
of all types of initiatives advancing the development of domestic
scientific and technology was emphasized.

Between 30 January and 4 February 1932, the 17th Conference
of the party gave particular attention to the need for the
development of extensive scientific and technical propaganda.

Courses organized by GIRD in 1932 on rocket technology and /133
the history of astronautics were particularly significant in
the training of specialists in the new technology. The course
on the theory of rocket engines was read by F. A. Tsander, the
course on the dynamics of reaction apparatus by V. P. Vetchinkin,
the course on the theory of air breathing reaction engines by
B. S. Stechkin, the course on hydrodynamics and gas dynamics by
B. S. Zemskiy, while N. A. Zhuravchenko read the course of
lectures on experimental aerodynamics.

In order to activate work in the field, the organizational
and mass operations section of GIRD developed a program of
courses for propagandists in 1932, designed for 40 hours. The
training plans of the courses were sent out to peripheral organ-
izations.

In April of 1932 there were six communists at GIRD,
organized into a party group. The first party group organizer
was L. K. Korneyev. In early 1933, an independent party
orr^.iization was set up at GIRD. The first secretary of the
paity bureau was the Deputy Chief of the second team of GIRD,
Nikolay Ivanovich Yefremov.

The communists of GIRD were the first combat detachment
of the organization. The communists actively influenced the
scientific and production life of all subdivisions of GIRD,
and were leaders in the shock movement and in socialist competi-
tion. When difficulties arose in the work of any team, the
part- organization always mobilized the communists and gave
h'^lp to lagging sections.

During the time of most intensive work, the communists
gave personal examples, working diy and night, as for example
during the time of the first launching of the 09 and GIRD-X
rockets.

114

2.5. Liquid-Fueled Rocket Engines and Rockets of GIRD

The primary results of the work of the first and second
teams of GIRD were the 02 rocket for the RP-1 glider, the 10,
09 and 03 engines for the GIRD-X, GIRD-09, GIRD-07 and GIRD-05
rockets. Furthermore, experiments were performed with OR-1 and
individual LRE units.

/134

Cross Section of 02 Engine with Prechamber

The 02 Engine

Sergey Pavlovich Korolev (even before the organization of
GIRD) attached great significance to the creation of a piloted
flight vehicle with an LRE. This is indicated by his interest
in the plans of F. A. Tsander, his great support of the work
performed in the first team of GIRD on the OR- 2 engine, the
creation and personal leadership of the fourth team of GIRD,
which developed the rocket plane flight vehicle, on which the
OR-2 liquid fueled rocket engine was to be installed.

The 02 engine was first tested in the OR-2 version, i.e.,
the form in which it was planned by F. A. Tsander.

After three tests (18, 21 and 26 March 1933), in order to
improve the operating capacity of the 02 engine, further testing
was performed with a fuel with lower heat content, consisting
of liquid oxygen and 85% ethyl alcohol. Furthermore, the
design of the liquid-carrying portion of the cooling system and
of the combustion chamber itself was simplified; the cooling
agent used was the liquid oxygen, the heating and partial
evaporation of which in the cooling chamber had a favorable
influence on processes within the chamber; the chamber was
equipped with ceramic inserts, requiring studies on the selec-
tion of refractory heat insulating materials. Thus, the 02
engine differed significantly from the OR 2 designed by F. A.
Tsander.

In many documents this engine is called the "ORD-2."

115

During its development, the design of the 02 engine changed /135
from model to model. According to the special program of
investigations, in July of 1933 a chamber was tested with a
graphite insert, which burst during the 55th second of opera-
tion due to the presence of impurities in the graphite mass.
In October, the chamber was tested with an insert made of
carbon electrodes; the insert burned out during the 62nd second
of operation. The insert or lining was a separate part (of
graphite, aluminum oxide or magnesium oxide), placed tightly
in the chamber and nozzle during assembly. In many cases, the
refractory insulating material was applied in the form of a
thick mass to the inner surfaces of the chamber and nozzle,
then subjected to the required heat and mechanical treatments.
In subsequent experiments, the graphite facing was covered by a
protective refractory mass in order to avoid oxidation of the
carbon.

By December of 1933 when the first team of GIRD had
become a part of the RNII, it was finally established that the
chamber should be lined with corundum, the nozzle with magnesium
oxide, and on 20 December 1953 a chamber with this insulation
operated 2 minutes 40 seconds without damage.

At GIRD, the development and testing of the 02 engine were
conducted by A. I. Polyarnyy (Chief Designer), L. S. Dushkinym,
L. K. Korneyev and other members of the first team. The devel-
opment of heat insulating refractory coatings involved the par-
ticipation of Ye. K. Moshkin. Final development of the engine
was performed in the oxygen team of RNII, headed by M. K.
Tikhonravov. Testing of the 6 main versions of the 02 engine
on the stand of the third laboratory of RNII was conducted by
L. S. Dushkin, A. I. Polyarnyy, B. V. Frolov and others.

The first version of the 02 engine was a cylindrical com-
bustion chamber made of sheet copper 1.5 mm thick. The com-
bustion chamber was lined on the inside with aluminum oxide,
the nozzle -- with magnesium oxide. The shell of the chamber
and the nozzle were made of low-carbon steel. The head of the
engine carried a plate (called the jet plate) acting as a
sprayer. The plate had 35 apertures 0.5 mm in diameter, through
which the alcohol was sprayed. The oxygen, heated in the cooling /136
section and partially vaporized was fed into the combustion
chamber through two tubes welded to the assembly ring in the
area of the entry to the cooling section and apertures (windows)
located in the cylindrical portion of the chamber wall near
the head. Ignition was by sparkplug, introduced to the combus-
tion chamber before start-up through the nozzle.

116

, *

MUiMMiyMtmiMTrrrnimiiiuiiiniiimkuM,

Cross Section of Final Version of
02 Engine

The second version had a shaped nozzle, calculated by the
method of Professor F. I. Frankel. Considering the complexity
of manufacture of shaped nozzles and the multitude of problems
not yet solved, GIRD did not continue to use this type of
nozzle. Shaped nozzles became widely used only during the
post war years.

The third version of the engine had a nozzle like the first
version, but with a broader cone angle. The fourth and fifth
versions were equipped with the nozzle of the third version and
a prechamber. After a long series of tests performed in 1934-
1935, the final version -- the 02-s engine -- was designed.
This engine underwent testing in 1955.

The basic data of the final version of the 02-s engine
are as follows. Length 570 mm, outside diameter 90 mm, diameter
of critical nozzle cross section 26 mm, volume of combustion
chamber 930 cm^. The liquid oxygen consumption was 0.338 kg/sec,
the consumption of 96% ethyl alcohol was 0.162 kg/sec. With a
feed pressure of 20 atm, the pressure in the combustion chamber
reached 11 atm. The engine developed a thrust of 100 kg and
operated without damage up to 60 sec. The cylindrical portion
of the combustion chamber was lined with a refractory heat
insulating material based on aluminum oxide, the nozzle was
lined with magnesium oxide.

Thus, an LRE was created as a result of work beguu at
GIRD and completed at RNII.

The 02-s engine was tested in 1936 on the 216 winged
rocket. This rocket was launched from a catapult truck accel- /137
erated by solid fueled engines. Four tests were conducted;
in two cases, the 216 rocket left the truck normally, climbing
one time on an inclined, straight trajectory to an altitude of
about 500 m.

117

The 10 Engine

The first team created the 10 engine for the GIRD-X rocket.
It was designed to develop a thrust of 60-70 kg for a duration
of 30 sec with a chamber pressure of 8-10 atm. The work on the
engine was begun in January of 1933 under the direct leadership
of F. A. Tsander.

The first version, developed by F. A. Tsander, was an
engine which burned liquid oxygen and gasoline with the addition
of metal, which was to be fed into the combustion chamber in
powdered and melted form. In parallel with the planning of the
engine, studies of the feeding and ignition of metal fuel were
conducted, as a result of which it became clear that the prepa-
ration of metal fuel for combustion and use in the engine
involved too great technical and operational difficulties.
Therefore, the first version of the engine was not manufactured,
and the second version was designed only for liquid oxygen and
gasoline, without the addition of metal fuel.

The second version was an all-metal welded
The inner wall of the chamber was made of stainl
outer wall (jacket) of ordinary struc* ral steel
was pear-shaped and featured external liquid coo
consisted of a mixing chamber with sprayers, a d
central portion, i.e., the combustion chamber it
nozzle. Liquid oxygen was fed to the lower port
nozzle through a collector into a cooling cavity
washed over the outside of the chamber wall and
chamber through jet- type sprayers. Gasoline was
upper portion of the mixing chamber through jet-
formed by drilling holes into the side surface o
The working mixture thus formed passed through a
the central portion of the chamber.

structure,
ess steel, the

The engine
ling. It
iffuser, and a
self, and the
ion of the

3 mm wide, then
entered the

fed into the
type sprayers,
f the chamber.

diffuser into

Oxygen

♦

Pressure measurement

Spark plug

Gasoline

*

/138

Second Version of the 10 Engine

The testing of the 10 engine, begun in August of 1933, and

the improvement of its design were performed by L. S. Dushkin,

L. K. Korneyev, A. I. Polyarnyy, V. P. Avdonin, M. G. Vorob'yev

118

and others. During flame testing, changes were made in the
design of the chamber. A chamber with a prechamber with a
shaped contour was used; the prechamber was connected with the
chamber by means of a diffusor. The engine was tested on
liquid oxygen and gasoline. During flame tests, the excess
pressure in the chamber varied little and did not exceed 2.5
atm.

The assigned time for fulfillment of the plan came to an
end, and the engine had not yet been developed. Rupture of the
combustion chamber required that further testing be performed
using a fuel consisting of liquid oxygen and ethyl alcohol. The
concentration of the alcohol (most frequently an 85% solution
in water) was selected as a function of the assigned operating
mode of the engine, and oxygen was used as before as the cooling
fluid.

The third version of the 10 engine had a mixing chamber,
i.e., a prechamber with a flat bottom, carrying the jet- type
sprayers for alcohol feed. The fuel used was 781 ethyl alcohol.
The oxygen sprayers were located on the cylindrical surface of
the chamber, closer to the component mixing zone. The cooling
of the central portion of the combustion chamber was intensified
by additional input of liquid oxygen to the cooling cavity in the
region where the combustion chamber was joined to the nozzle.
During flame testing, the combustion chamber burst due to
excessive thermal stresses.

The fourth version of the engine, made of SKh-8 steel, was
tested on 2 October 1933 on the powder test stand at RNII.
The pressure in the chamber reached 8 atm, the thrust -- 75 kg.
During the test, the peak of thrust was recorded when the oper-
ating mode was reached, then the thrust decreased during the
16th second. The engine was shut down after 21 seconds. An
inspection revealed a crack in the inner wall of the central
portion of the chamber.

/139

Oxygen

Oxygen

Spark plug hole Fue l

Pressure

measurement

Temperature
measurement

Fourth Version of the 10 Engine

119

The next model of this engine was made of ENERZh-7 steel.
It was tested together with the fuel feed system on a balance
frame which carried the tanks, elements of the feed system
and combustion chamber. The force developed by the engine was
transmitted by this frame to the thrust-measuring device.

The basic data of the engine were as follows. Length
312 mm, outside diameter 92 mm, nozzle critical cross section
diameter 24 mm, volume of combustion chamber 450 cm^. The
consumption of 85% ethyl alcohol was 0.280 kg/sec. With a
pressure in the chamber of 10 atm, the thrust was 65-75 kg.
The specific impulse, according to the data of three successive
tests, was 162-175 sec.

Based on the results of the testing, the decision was made
to install, the engine in a rocket. The test report included the
following: "Since the design data have been exceeded and a
thrust of 75 kg achieved, with a pressure in the combustion
chamber of 10 am, operating time of 20 sec, and keeping in mind
the slight, easily repaired damage to the chamber occurring
during-two tests, it is considered possible to launch the 10
rocket burning liquid fuel into the air using the motor
tested."

Work with the 10 engine was continued at RNII. Beginning /140
in Febraury 1934, adjustment tests and further studies of this
engine were conducted on the RNII test stand. The fuel was
fed into the combustion chamber through jet-type sprayers.
Two specimens were developed: an all-metal and a ceramic, i.e.,
with ceramic lining. The all-metal chamber differed little
from the last GIRD chamber.

The other version of this engine had a nozzle with a
refractory ceramic insert. The oxygen cooled only the central
portion of the chamber and the mixing chamber. On 25 November
1934, during testing of the engine at RNII, it was considered
possible to use the 10 motor with ceramic nozzle to launch
rockets with powered flight times of 25-30 sec, since the
thrust produced experimentally was 70 kg. The flame resistance
of the nozzle made of ceramic was considered satisfactory,
since no melting was observed after 25-30 sec operation of the
engine.

The features of the 10 combustion chamber included the
use of liquid oxygen and ethyl alcohol as fuel components, the

1 GIRD Archives, d. No. 3-050, p. 3.
2
At GIRD, the rocket was called the 10. It was given the name

GIRD-X later.

120

presence of pre'hambers, the pear shape of the combustion
chamber, and the external liquid oxygen cooling.

The 10 liquid-fueled rocket engine was the first Soviet
LRE tested by rocket flight.

The 09 Engine

The second team developed the 09 engine for the GIRD-09
rocket. After long search for the most expedient design, i.e.,
the most reliable design providing for the most rapid develop-
ment, the team selected a hybrid fuel engine. This was facili-
tated by the suggestion that solidified (gelled) gasoline be
used as fuel. This gel was produced by dissolving colophony
in gasoline. Liquid oxygen was used as the oxidizer. The
entire fuel reserve was placed in the inner cavity of the com-
bustion chamber, while the liquid oxygen was poured into the
fuel tank.

During planning of the GIRD-09 rocket, use of this plan
allowed a reduction in rocket weight, simplified the design /141
of the fuel feed system (only the oxidizer had to be fed into
the chamber). True, the development of the mode of processes
within the chamber was made more complex, and the evenness,
stability and reproducibility of combustion of the fuel were
reduced.

The chamber of the 09 rocket engine was made and tested in
various versions differing in the design of individual elements.

The first model suitable for testing was completed on 31
December 1932, flame tests were begun in April of 1933.

The chamber of the rocket engine consisted of a sprayer
disc, a cylindrical portion (combustion chamber) with a screen
and the nozzle.

The sprayer disc was a disc with tiny apertures through
which the liquid oxygen was sprayed into the combustion chamber.

The combustion chamber included a cylinder with apertures
called the screen. The diameter of the cylinder was less
than the diameter of the combustion chamber. The solidified
gasoline was placed in the cavity between the screen and the
chamber wall before start-up. The oxygen flowed through these
holes in the screen to the gasoline and the combustion products
flowed back through these holes into the central portion of
the combustion chamber and to the nozzle.

121

The chamber did not have external
liquid cooling. The walls of the
chamber were protected from burning by
a layer of asbestos and by the fuel
itself, which burned radially, i.e.,
in the direction from the screen
toward the chamber wall. Thus, if
the gasoline burned evenly, the com-
bustion products could contact and
heat the chamber wall only during the
last instants of motor operation.

The nozzle was fastened to the
cylindrical portion o r ' e chamber;
it also had no extern iquid cooling.

The basic data 01 engine are

as follows. Thrust 25-33 kg, liquid

oxygen feed pressure 13.5 atm, pressure

in the combustion chamber 5-6 atm,

length of chamber 320.5 mm, maximum

diameter 145 mm, diameter of nozzle

n. „, u«««.i«„ n r *u„ critical cross section 26 mm.
Final Version of the

ngine Flame tests of the engine were

used to develop its individual units. In May of 1933, the
team tested the combustion chamber for strength under static
load and hydraulic shock conditions. In April and May of 1933,
the oxygen valve, reduction valve and other units were
tested.

/142

In April of 1933, work was performed on the selection of
the type of ignition. First, pyrotechnical ignition was
tested. The igniter consisted of gun powder, wood charcoal
and a (third) ballast component. Considering the insufficient,
reliability of the pyrotechnical ignition system, it was
decided to use electric spark plugs powered by a magnito
for ignition.

The engine was tested under the leadership of S. P.
Korolev and M. K. Tikhonravov with the participation of
N. I. Yefremov, V. S. Zuyev, Yu. A. Pobedonostsev, Z. I.
Kruglovaya and other members of the group.

The chamber was first made of structural steel, then of
copper. Seven tests performed beginning in July of 1973
showed that these chambers did not provide the required oper-
ating time, even when lined with asbestos. Chambers made of
ENERZh steel were then tested. The final version of the
chamber was made of brass,

122

The nozzle, first made of structural steel, was replaced
with a copper nozzle, then with nozzles of ENERZh steel. These
nozzles worked fairly well, although in some cases they were
burned through in the region of the critical cross section.

Chambers with screens of various types were tested --
with ribs, to hold the solid gasoline in place and facilitate
even burning, and without them; the material of ths screen
was varied (celluloid, aluminum, structural steel, chrome
steel, etc.).

The chambers tes.ed were equipped with heads of various /143
types, differing in the direction of the jets, structural
material and number of apertures, which was varied from 5 to
14. Each number of apertures produced a different thrust
leve 7 .

Some tests resulted in explosions, as for example on
28 April 1933.

The pressure in the oxygen tank was developed by evapora-
tion of a portion of the liquid oxygen in the tank due to
heat exchange with the surrounding medium, and the design
pressure was maintained in the tank by means of a safety
valve.

Stable and reproducible pressure was not achieved
immediately, resulting in repeated changes in the design of
the oxygen valve, safety valve and other elements.

By mid- August 1933, the rocket engine, in the form in which
it was installed in the rocket, passed final adjustment testing.

The 03 Engine

The 03 engine was designed for the GIRD- 07 rocket devel-
oped in the second team by M. K. Tikhonravov, N. I. Yefremov,
V. S. Zuyev and other workers of GIRD. It was constructed in
1933.

In this engine, spiral sprayers similar to the sprayers
used in 1932 in the ORM-12 engine were used to inject the
fuel (gasoline) . The combustion chamber of the 03 engine was
connected to the nozzle by a threaded joint. The combustion
chamber did not have external liquid cooling. Testing of the
engine began 17 Od >ber 1933. Beginning in February of 1934,
ethyl alcohol was used as the fuel rather than gasoline.

The engine had the following design data: pressure in
chamber 18-20 atm, thrust 80-85 kg, operating time 22-27 sec.

123

After a number of unsuccessful tests, work on the 03
engine was halted, and the 10 engine was installed on the
GIRD- 07 rocket.

One of the peculiarities of the operation of GIRD was
the assignment of each LRE developed to a given rocket in order
to power its flight. Therefore, in analyzing the activity of
GIRD, we cannot limit ourselves to analysis of work performed
with LRE on the test stand. GIRD set itself the primary task
of achievement of rocket flight with LRE. It was therefore /144
frequently called the rocket organization.

The GIRD-09 Rocket

The first Soviet experimental rocket, with the 09 hybrid-
fuel engine, was created in the second team of GIRD under the
leadership of M. K. Tikhonravov.

In August, 1933, the GIRD-09 rocket passed preliminary
tests and attempts were made to launch it, unsuccessful for
various technical reasons. After elimination of individual
problems, on 17 August 1933, under the direct leadership of
S. P. Korolev, the rocket was launched and the first hybrid
fuel rocket in the world flew.

This date has entered the history of astronautics as the
day of the launch of the first Soviet liquid-fueled rocket.

The basic data on the GIRD-09 rocket are 1 : length 2.4 m,
diameter 0.18 m, launch weight 19 kg, including 5 kg of fuel,
payload weight (parachute and several instruments) 6.2 kg.
The 09 engine installed on the rocket developed a thrust on
the order of 25-33 kg.

The main parts of the GIRD-09 rocket were made of aluminum
alloys. The rocket consisted of a body, the lower portion of
which carried four stabilizers. Inside the body was the
oxygen tank, made of a pipe. The large annular clearance
between the tank and the body acted as thermal insulation.
Between the oxygen tank and the chamber of the rocket engine
was a manually operated starting valve. The nose portion of
the rocket carried the parachute and its ejector.

The launch of the rocket on 17 August 1933 was conducted
as follows. After it was placed on its vertical guides, the
rocket was filled with liquid oxygen. Heat exchange with the

At GIRD, this rocket w called the 09. The name GIRD-09 was
given the rocket significantly later.

124

surrounding medium caused a portion of the oxygen to vaporize,

increasing the pressure in the tank. When the design pressure

was reached, the oxygen start valve was opened and at the same

time power was fed from the raagnito to the spark plug; the fuel

in the chamber ignited, and the motor came up to power. During /14 5

this launch, the rocket left its guides and rose to an altitude

of 400 m.

The entire flight, from launch to landing, lasted 18 sec.
During a second launch in the fall of 1933, the engine exploded
after the rocket reached a height of about 100 m.

In 1934, now in RNII, several GIRD-09 rockets were made,
with slight changes in design, and a number of successful
flights were conducted. The greatest altitude reached by the
rocket was 1500 m.

The GIRD-X Rocket

The first experimental Soviet rocket with LRE was the
GIRD-X rocket with the LRE 10, burning liquid oxygen and
ethyl alcohol. The initial development of the rocket was
performed by F. A. Tsander, the working plan, supplementary
testing of units, assembly of the rocket, finishing of the
entire complex of equipment and launching of the rocket were
by the members of the first team of GIRD.

The principal plan of the rocket was developed by the
members of the first team, based on materials of F. A. Tsander.
Then a rough plan of the rocket was made up and ballistic
calculations were performed - the cet ^r of gravity, center
of lateral effort, dimensions of tail fins and stability over
the entire flight trajectory were calculated.

Using the methods of LRE design developed by F. A. Tsander,
A. I. Poiyarnyy, L. S. Dushkin and other workers calculated the
basic version of a rocket engine combustion chamber in detail.
Experimental data produced in the development of the versions
of the 10 engine described above were used.

The launch weight of the rocket was 29.5 kg, 8.3 kg of
which was fuel, 2 ig -- the nose portion.

The rocket consisted of five sections. In the first
section, the nose portion of the rocket, was the parachute
and its ejection device; in the second section was the oxygen
tank and equipment; the central, third section, carried a
cylinder with air compressed to 150 atm. Here also was the
start-up equipment, including a pressure reducer, starting
valves, etc. The fuel section contained a tank of alcohol. / 1 4 7

125

The bottom, motor section carried the engine and the start
valve on the fuel feed line.

'he 09 Rocket

The first
launch of the rocket
under the direct
leadership of S. P.
Korolev, was held
on 2 5 November 1935

and went as follows.
After test ing of all
equipment, the
rocket was placed on
its guides and the
pressure accumulator
cylinder was charged.
Then the tank; were
filled with the i'uvl
components: Prst the
fuel , then the oxi -
dicer. After shocking
the pressure, the
start valves were
opened and the magnito
fed current to the
electric s pa r k p 1 u g .
The rocket smoothly
1 i f ted off from the
launch support and
>egan to climb
vertically with
rapidly increasing
..peed . At an alt itudc
of 7 5-80 m , due to
damage to the motor
mount, the rocket
changed its flight
direction sharply and
fel 1 to F.arth at a
distance of 150 m
from the launch point.
The engine, which
fired for 12-13 sec,
developed a thrust on
the order of 65-70 kg
in flight.

t ion

J-edoro

The 1 . hing of the (ilRD-X rocket involved the participa-
of L. K. kori --'•.', L. S. Pushkin, A. J. I'olyarnyy, k. K.
L. N Kc b, ; ina and other.

126

m

The GI

First

Rocket

Ins pec
and al
of 17
rocket

RD-X,
Domes t
with

tion a
co hoi

Novemb
. The

The design of the GRD-X rocket was
further developed in later, improved
Soviet rockets.

The GIRD- 07 Rocket

This rocket was developed by the
second team of GIRD in 1933. It had an
unusual shape, a result of the search for
improved aerodynamic characteristics, since
the GIRD-07, like earlier rockets, was not
planned to be equipped with a control
system. The diagram of placement of the
engine in the rocket and the design of the
stabilizers were selected on the assumption
that flight stability of the rocket would
be increased by moving the point of appli
cation of the thrust vector closer to the
nose portion and further from the center
of lateral effort of the rocket.

In its final version, the rocket was
driven by the LRH 10 developed in the first
team of GIRD. Liquid oxygen and ethyl
alcohol were carried in tubular tanks
mounted one in each of the four stabilizers.
The fuel components were fed to the combus-
tion chamber under compressed air pressure.

The rocket wa
at RNil. M. K. Ti
Yu. A. Pobcdonosts
part in the test.
after the command
sparks burst from
rocket. Immediate
the stabilizers caught
ic tinued 2"" seconds,
LRU was switched off.
from its position
fter the test revealed tha
tank had burned through.
er recommended further dev
07 rocket was later teste

s tested 17 November 1954
khonravov, V. S. Zuyev,
ev, L. S. Dushkin took
Five to eight seconds
for ignition was given,
the lower portion of the
ly thereafter, one of the
fire. The burning con-
after which the engine
The rocket did not move
and remained in the guides
t the combustion chamber
The resport on the test
elopment of the engine and
d in flight.

Many ideas and promising design solutions were embodied in
the GIRD-09 and GIRD-X rockets; their flights of 17 August and
25 November 1933 laid the foundation for flight testing of
liquid- fueled rockets in the USSR.

/148

/ISO

127

General Vic;-: of the 0? Rocket

In Nakhabino, near Moscow, on the site where the G1RO-09 and
G1RD-X rockets were launched, an obelisk has been placed,
carrying the names of S. P. Kcrolev, F. A. Tsander and M. K.
Tikhonravov.

Air Breathing Reaction Engines

The theory of air breathing reaction engines, developed
by Boris Sergeyevich Stechkin, allowed practical fcork to be begun
on the creation of air breathing reaction engines. The first
experimental studies of direct-flow air breathing reaction
engines were performed at GIRD, in the third team, led by Yu. A.
Pobedonostsev. The theoretical calculations were followed by
practical work -- experimental study of models of ARE and
individual elements on the III- 1 test stand, built in March of
1933 for this purpose.

On 15 April 1933, the first ARE was started, and operated
for 5 minutes. It was noted in the conclusions that the tests
of the engine fully justified the theoretical assumptions.
These tests served as the basis for experimental studies of ARE
in the USSR.

As work on the testing of ARE models developed, the
methods of study were also improved. Beginning in June of
1933, tests on the IU-1 included the measurement of the thrust
developed by the engine being tested.

Since Yu. A. Pobedonostsev suggested that slowly burning
solid fuels be used, after studying' a large number of fuels,
he selected white phosphorus and solid gasoline. On 12 July 1933

128

/149

At the Launch Site of the First Domestic
Liquid- Fueled Rocket I Inscription Reads:

On this site in 1953, the first Soviet rockets,
the "09" and "GIRP-X" were launched. Korolev, S. P.,
Tsander, F. A., Tikhonravov, M. K. » to the UIRI>
workers from the Komsomol members of Nakhabinskaya
Middle School, No. 2]

at a firinr. range near Moscow, a fully successful test of a
combustion chamber burning phosphorus was conducted. During
the tests of this same ARl- with solid gasoline, ignition was
achieved by a chunk of phosphorus placed in the axial channel
of the combustion chamber. In order to develop the most
successful means of ignition of the fuel in the combustion
chamber of ARl:, a number of tests of powder ignition devices
were conducted in July of 1933, including tests with a conical
chamber burning ethylene.

/ I d I

To increase the effectiveness of ARE not only at supersonic,
but also at subsonic velocities, a search was conducted to find
plans in which the incoming air was not only compressed in a
diffusor by the velocity head, but also by some other sort of
additional device; a plan was drawn up for a pulse jet, in
which the air stream was periodically admitted to the chamber
by a valve moved by the pressure drop between the diffusor and
chamber. An experimental combustion chamber was constructed
at GIRD in June of 1933 to study the possibility of creating
pulse jets.

The tests of pulse jets performed at GIRD in 1933 indicated
the basic problems arising in the design development of engines
of this type, and allowed the volume and difficulty of solution
of these problems to be estimated. It was decided to direct
all attention toward the study of ram jets, as the most promising
type, but the study of pulse jets was reactivated at RNII in
1936-1939.

The successes of the first experimental studies allowed
flight testing of ram jets to be undertaken. Yu. A.
Pobedonostsev suggested that the engine to be tested be included
in the body of an artillery shell, so that the jet could be
tested at supersonic speeds, i.e., in the area where ram jets
are most effective. The first series of flight tests in
September of 1933 confirmed the calculation data. In February
of 1934, now at RNII, a second series of tests was held and in
1935 a third series of jet flight tests was conducted. Six
more versions of engines were planned for these tests, placed
in the body of a 76 mm shell. Some versions included several
groups, differing in the dimensions of the diffusor inlet cross
section and the critical cross section of the nozzle and fuel
reserve.

The ram jet engines designed by Yu. A. Pobedonostsev were /152
the first reaction engines which operated in the supersonic
range. The experiments confirmed the primary theoretical
conclusions of B. S. Stechkin on the efficiency of engines of
this type and the need for further improvement of their design.
The experience of later years showed that the scientific direc-
tion developed in the third team of GIRD was fruitful. In
later years, designers concluded that future large booster
rockets could use ARE as first stage engines.

Thus, GIRD demonstrated broad capabilities fur practical
realization of the leading ideas of Soviet rocket scientists
in a short time: work on the creation of IRE (OR- 2, 02) for
piloted vehicles, rockets with LRE (GIRD-X), with hybrid fuel
engine (GIRD-09), study of direct flow air breathing reaction
engines, The development of reaction and rocket-space technol-
ogy indicates the correctness of the scientific trends selected

130

at OliU), the timeliness o£ the .statement and solution of com-
plex problems of the new area of technology by CARD.

1. 0,

The Leningrad Croup for the Study of Reaction
Motion (LenGIRD)

An important role in the development of studies on reac-
tion technology was played by the Leningrad Group for the
Study of Reaction Motion [LenGiRD), organized 15 November 1931
on the suggestion ot the well-known aviator P. ?-', Fedoseyenko^
in the Leningrad oblast Soviet of Osoaviakhim. The initiative
group, in addition to P. !•' . i-'edoseyenko, included: Professor
U. A. Ryu in, Vu. 1. Perel'man, Pngineers V. V. Razumov, M. V.
Gazhala, A. N. Shtern, V«. , Ve. Cltcrtovsk iy „ i. X. Samarin and
M. V. Machinskiy. B. S. Pet ropav lovskiy and V. A. Artem'yev,
workers of GDL, were very helpful in the organizat ion of the
group and i ts work.

Professor Nikolay Alekscyevich Rynin (1877-19421 was a
comprehensively trained engineer, a great scientist and an
outstanding lecturer. During the initial period of his scien-
tific activity, his primary attention
was concentrated on determination of
the aerodynamic loads acting on various
structures. A significant portion of
his activity was dedicated to aviation,
problems of flight in the stratosphere
and interplanetar voyages, which he
considered to he a logical continuation
and completion o( aviation. His
scientific, engineering am! pedagogic
ac t i v i t y we re comb i ned w i t h g r ea t
organizational work. in 1909, he
participated in the creation of an
aerodynamics laboratory at the Peter-
burg kikolayevskiy Institute of
Railroads; he organized the first
aviation competitions and flights in
Russia; he himself took part as a
pilot in f 1 i ght s o f a i re ra ft , cent rol I ed
aerostats, and flew in balloons. In
1920, with his participation, a
Department of Air Voyages was organized
at the Petrograd Institute of Railroad
Lngineers. The department was later converted to the Civil
Aviation Institute, then to the Leningrad Military Air

/1 53

Nikolay Alekscyevich
Rynin

P. P. Pedoseyenko was the commander of the Osoaviakhim- 1
stratostat, which set a worl altitude record in 1954.

131

engineering academy, now the Military Air Engineering academy

tilers i A. F, Mozh ;■ skiy.

N. A
organiser

created in
fleers, whe
to his act
theory of
carefully
and 193:,
down from
of theoret
scientists

Rynin was an interplanetary flight enthusiast, the

man of the : .- * m for Int ■■ met ry .-. iges »

1928, in the Leningrad Institute of Railroad hngi-
re he began work as a prof esso i 1921 Is : ion
ive worl oi LenGIRD, Jc ■ lo| - : > • bleros of he
interplanetary voyages and was published. He
coile • •; and ; ••: • bed in nin be ■» ; be •. i • : - 8
ideas 3 flight i s nit • . up ths . nd - id - ; - s ed
ancient ti', , ? dr im< . ore . : tin est ts
ij and -'- r intent * studies oi domes t ie aiu for* gn

etc.

live i tod •: ■'.'•'.• - vf N.
interest to historians and those

technology.

A. Ry 'mart : *' -. it

ntei est ,••• n . >eki t and -pace

of Rynin h;t h i s s '" rater on

; ' 1 s < the m > a i .

the far

Yakov is ■ « - -• •; crel i . > i 2

19'-? was a ' = ' let ;. - enl s . ; t , t w 'II-
kno* ii >opular >f matli .mcs,

physic?, chemistry and astronautics.
'■ ; hool "Ini ' ' ;. •. lathemal - »"
"I tore ■■ in| • - oi ■ •• •• d i t her; , in
wh . - Va. I . Pe re I 'roan desc hi the
many interesting technical problems
involved in the development of various
sciei i i: t are : : '. * in I em. "J i*J « .'-day.

"•; w i - a >ab Le oi ■ howi \\g
ph ««.-. n ski *etningly qu : nro >n nd
or-.- : rsa y in a complc f c I > new ind

une: pec ted I ; -'."it.

In 191' / «'.'.'., book h>

Ya. I . P 'l'man ei titled 'InterpJ n<

ta I'oy ages' - : - > ; --' i . . ii 'd. It went
thr - t»l LO pi in ings is - ears.
Tl t »oi '■ . •• it ten el< r iy : i s zing
the tethods of it t glu oi ma '<
space then I >cuss< : ; .' the i eratui * ' - N . ■■■ . Lent if - and-
point. Ya. I. Perel nan ;howed os incing ; that a
present st ge I ievelopm m >f ien< - aid I »c tn< • . -aly
the 5 ., kel ec I be • . ns .,- • . • in U means of rry i

into space.

Yakov Isidorovich

/154

13.

Yakov Isidorovich wrote books on Konstantin Eduardovich
Tsiolkovskiy, Galileo Galilei and Thomas A. Edison.

In 1928, Ya. I. Perel'man took part in the work of the
Leningrad Section for Interplanetary Voyages, and beginning in
1931 he directed the work of the Scientific Propaganda Group
of LenGIRD. Beginning in 1934, after LenGIRD was converted to
the Section for Reaction Motion of the Leningrad Bureau of
Aviation Technology of the oblast Osoaviakhira Soviet, he
continued his propagandists work until the beginning of the
Great Patriotic War. The name of Perel'man has been given to
one of the craters on the far side of the moon.

On 13 November 1931, a general meeting of LenGIRD activists /1S5
was held in the District Red Army and Navy Hall. After the
introductory words of Professor N. A. Rynin concerning the goals
and tasks of the organization, V. V. Razumov presented a report
on his plan for a high altitude rocket and the immediate
possibilities for interplanetary flight. The meeting was
ended by the election of officers of LenGIRD, which included
V. V. Razumov (President), Ya. I. Perel'man (Vice President),
N. A. Rynin and Engineer M. V. Gazhala. Later, M. V. Machinskiy
was selected as Chairman of the Technical Council of the
organization.

Vladimir Vasil'yevich Razumov was

born on 15 June 1890 in Peterburg.
After graduating from the Marine
Engineers School in Kronshtadt, he
worked at the Admiralty Ship Repair
Plant until 1931; he was a scientific
consultant for the Leningrad Division
of Dirigible Construction and in 1933
headed a design bureau for the con-
struction of an all-metal Tsiolkovskiy
dirigible. V. V. Razumov headed the
planning and design group of LenGIRD;
under his leadership, eight rocket
plans were developed in 1932-1933.

Vladimir Vasil'yevich
Razumov

of the primary
created -- the
Gazhala), the
the Scientific
the Laboratory
Port Group (he

The next meeting of LenGIRD,
involving about 40 persons, was held
21 November 1931. This meeting dis-
cussed practical measures related to
the development of work on the study
problems of reaction motion. Five groups were
Scientific Research Group (headed by M. V.
Planning-Design Group (headed by V. V. Razumov),
Propaganda Group (headed by Ya. I. Perel'man),
Group (headed by I. N. Samarin) and the Rocket
aded by Ye. Ye. Chertovskiy) . Each group included

/1S6

133

five to six persons. The scientific secretary of the oblast
Osoaviakhim Soviet, V. I. Shorin, took part actively in the
formation of LenGIRD.

In 1932, courses on rocket technology organized by LenGIRD
were conducted.

In 1932, three rockets were planned at LenGIRD with powder
engines (a photographic, illumination and recording rocket), as
well as a recording rocket with LRE, and in 1933, high altitude
rockets with LRE were planned.

In order to develop rocket engines, two sections were
organized in the Planning-Design Group of LenGIRD. One of these,
headed by V. A. Artem'yev, created a number of solid-fueled
rocket engines between 1932 and 1935, which were installed on
all the experimental rockets of LenGIRD which were successfully
flight tested. The second section, headed by A. N. Shtern,
developed a rotary reaction LRE, the LRD-D-1, which burned
liquid oxygen and gasoline. However, this engine was never
completely constructed.

LenGIRD maintained communications with MosGIRD. MosGIRD
had as many as 400 members.

The Powder Rockets of LenGIRD

The photographic rocket, planned on the order of the
Leningrad section of the Scientific Research Institute for
Geodesy and Cartography, carried foru SRE designed by V. A.
Artem'yev.

Calculated data: altitude of flight 10 km; total weight
26 kg, including 6 kg powder; total length 1.32 m; diameter of
body 0.25 m; launch thrust 270 kg; engine operating time
4.33 sec; fuel -- smokeless trotyl pyroxylin powder.

The illuminating rocket was designed to supplement or
replace searchlights, and also to blind enemy aircraft, as an
air defense measure. The nose portion and stabilizers were
made of aluminum, the combustion chamber and nozzle of the /157
SRE -- of heat-resistant steel.

The calculated data of the illuminating rocket were:
altitude of flight 5 km; total weight 18 kg, including 3 kg
powder; total length 1.2 m; body diameter 0.15 m; launch
thrust 81 kg; operating time of engine 4.35 sec; fuel --
smokeless trotyl pyroxylin powder.

134

The LenGIRD
Recording Rocket

The plan for the rocket was completed
in February of 1932. September of this
same year, several experimental models were
made at the Leningrad Mechanical Plant,
which successfully passed flight testing
at the Osoaviakhim range.

The recording rocket was designed to
record data on the pressure and tempera-
ture of the atmosphere at altitudes of up
to 10 km.

The rocket consisted of a nose portion
with the required instruments, a body with
stabilizers and rudders and four V. A.
Artem'yev SRE.

The calculated data of the recording
rocket are: total weight 30 kg, including
10 kg powder; total length 2.11 m; body
diameter 0.23 m; launch thrust 148 kg;
engine operating time 12.7 sec; fuel --
trotyl pyroxylin powder.

The plan for the recording rocket
was produced in March of 1932 for the
Leningrad Geographical Institute. Later,
the design of the rocket was simplified,
the dimensions were reduced and three
versions were built: a high altitude
rocket, an agitation rocket (with leaf-
lets) and a shrapnel rocket. They were
flight tested at the firing range in the
Osoaviakhim camp. After summarizing the
experience produced, the group of M. V.
Gazhala planned, then manufactured in the
mechanical shops another 20 rockets with
similar SRE. The rockets, designed to
reach an altitude of 1 km, were tested at
the Aerological Institute in Slutska.

/158

Liquid- Fueled Engines

For the recording rocket, LenGIRD developed a plan for two-
chamber LRE, the LRD-D-1, which was to use liquid oxygen and
gasoline. The nozzles of the two chambers had an inclined
cross section, causing the rocket to rotate about its longitu-
dinal axis; the centrifugal force caused the fuel components
to enter the combustion chamber. The engine was called a
rotating reaction engine. The basic elements of the LRD-D-1

135

rocket were made of steel. The walls of the combustion chamber
and nozzle were to be cooled with the liquid oxygen, evaporating
in the cooling spaces.

The calculated data of the rocket with LRE are: maximum
altitude 5 km; launch thrust 200 kg with exuaust gas velocity
2000 m/sec; total weight 90 kg, including 17.8 kg oxygen, 4.89 kg
gasoline; weight of engine 16.0 kg; total length 2.665 m; diameter
of body 0.35 m.

The rocket was manufactured in 1932. Individual parts of
the engine, combustion chamber and nozzle were exhibited during
the first All-Union conference on the study of the stratosphere,
hela 31 March- 6 April 1934 at Leningrad. Since the engine was
never completely constructed and developed, the rocket was
launched to determine its aerodynamic characteristics late in
1934 using the V. A. Artem'yev SRE.

In 1933, the group of V. V. Razumov began the development
of the design of two recording rockets with design altitudes of
60 and 300 km with LRE burning liquid oxygen-gasoline fuel.
The combustion chamber and nozzle were cooled with liquid oxygen,
evaporated in the cooling space. The fuel component feed system
was by compressed gas cylinder.

The calculated data for the rocket designed to reach an
altitude of 60 km are: tot.^l weight 90 kg, including fuel 43.7 kg;
total length 3.62 m; body diameter 0.35 m; launch thrust 1000 kg;
engine operating time 28 sec.

The calculated data for the rocket designed to reach alti- /158
tudes up to 300 km: total weight 150 kg, including 110 kg fuel;
total length 5.9 m; body diameter 0.5 m; launch thrust 1571 kg;
engine operating time 51 sec.

Since the necessary production base and funds were not
available, this rocket was never manufactured.

During these years, a great deal of attention was given in
LenGIRD to the selection of fuel for LRE, the search for the most
favorable flight trajectories, the search for efficient rocket
and engine element (combustion chamber, nozzle) forms, the gas
dynamic studies of LRE, and the selection of materials for rockets
and engines.

The workers of LenGIRD constantly conducted extensive
explanatory work and gave consultation and practical aid in
problems of reaction motion both to various teams which arose
within the walls of military and civil e'ucational institutions,
and to individual enthusiasts.

136

In 1934, LenCIRD was converted to
the section for reaction motion, which,
under the leadership of M. V.
Machinskiy, continued propaganda work,
performed experiments on the effects of
accelerations on animals and continued
development and testing of LRE and
rocket models right up to the beginning
of the Great Patriotic War.

2.7. The Work of the Society

Problems of interplanetary voyages
attracted the interest of many
specialists. In addition to the litate
enterprises and groups for the study of
reaction motion, individual persons,
societies, sections, and clubs worked
across the USSR, making no small contri-
bution to the development of domestic rocket engine construction,

Diagram of Rotating
Reaction Engine

On 20 January 1924, at a session of the thejtetical section /160
of the Moscow Society for Astronomy Enthusiasts, F. A. Tsander
read a report "On the Design of an Interplanetary Ship and Flights
to Other Planets," and suggested that a "Society for the Study of
Interplanetary Voyages" (OIMS) be formed in the USSR.

In April of 1924, students at the Vi I ' . .'y Air Academy
imeni N. Ye. Zhukovskiy created a section *»a ^uterpJanetary
voyages in the Military Scientific Society of the Academy. The
founders and most active participants in the section were V. P.
Kaperskiy, M. G. Leyteyzen and M. A. Rezunov. The work of the
section was supported by K. E. Tsiolkovskiy, F. A. Tsander and
V. P. Vetchinkin.

On 30 May 1924, in the Great Auditorium of the Polytechnical
Museum, a lecture was read by a great engineer and widely educated
scientist, Mikhail Yakovlevich Lapirov-Skoblo, the subject of
which was interplanetary voyages. The lecture showed how modern
science and technology were capable of solving this problem.
Then, members were signed up for the "Society for the Study of
Interplanetary Voyages" (OIMS) .

First, the society had some 200 members. They were located
in the building of the Astronomical Observatory of the Moscow
Division of Popular Education -- at 13 Bol'shaya Lubyanka (now
F. E. Dzerzhinskiy Street). The society set very difficult tasks
before itself -- the unification of all organisations, all
scientists involved in problems of the study of interplanetary
voyages, and the creation of a scientific research laboratory.

137

The first, organizational meeting of OIMS was held on 20
June 1924. The officers of the society were elected at this
meeting --a presidium consisting of: President-- the then
well-known publicist and old Bolshevik G. M. Kramarov, Secre-
tary -- M. G. Leyteyzen, members -- F. A. Tsander, V. P.
Kaperskiy, M. A. Rezunov, V. I. Chernov, M. G. Serebrennikov.
K. E. Tsiolkovskiy was elected as an honorary member. The
society attracted the attention of talented scientists,
engineers and designers to problems of astronautics and helped
to popularize the ideas of rocket building and interplanetary
voyages. K. E. Tsiolkovskiy, V. P. Vetchinkin, M. Ya.
Lapirov-Skoblo and other famous scientists took part in the »• rk
of the society.

OIMS systematically held scientific-popular lectures. When
it was reported that the USA planned to launch a shell designed /161
by Professor Goddard to the moon to celebrate Independence Day,
July 4, OIMS held a debate on 1 October 1924 on the theme "Fligth
to Other Worlds." Although the auditorium was large, it was not
sufficient to contain all those who wanted to attend. Therefore,
the debate was repeated twice -- on 4 and 5 October --in the
Great Auditorium of the First University Physics Institute. F. A.
Tsander appeared on 4 October 1924 to report on a new ship which
he had invented for space flight.

The society worked for comprehensive expansion of its
propaganda activity. On 31 October and 2 November 1924, V. P.
Vetchinkin read lectures in the Great Auditorium of the Poly-
technical Museum on the possibility of interplanetary flight.
Here also an informative report was read by V. I. Chernov on the
construction of a rocket which he had designed. Lectures were
read on interplanetary voyages at aviation plants, in the club
of the Moscow Higher Technical School imeni N. E. Bauman, at
the Astronomical Institute imeni Shternberg and elsewhere.
Journeys by specialists were organized to read reports and
lectures in other cities: Leningrad, Khar'kov, Saratov, Ryazan'
and Tula.

OIMS existed but a single year, then broke up due to the
fact that the tasks which the society had set before itself could
not be performed with the funds available or the help provided by
ether organizations.

In June of 1925, Academician D. A. Grave spoke on the
subject "A Request for Clubs to Study and Master Space." That
same year, D. A. Grave, together with the great scientists Ye. 0.
Paton, B. I. Sreznevskiy, K. K. Seminskiy, V. I. Shaposhnikov and
other enthusiasts, created a "Club for the Study of Space (the
Cosmos)" in Kiev. The efforts of this club resulted in the open-
ing of an exhibit dedicated to problems of the study of inter-
planetary space in the section of inventors of the Kiev Associa-
tion of Engineers and Technicians on 19 June 1925.

138

In April-June of 1927, the world's first exhibit of models
and plans for interplanetar) apparatus and mechanisms was held at
the Moscow Association of Inventors. This exhibit displayed
interesting and unique materials on the work of Russian and
foreign researchers.

'■o'

The organizers of the exhibit were M. S. Belyayev, G. A. /162
Polevoy, Z. G. Pyatetskiy, 0. V. Kholshcheva. In late January
1927, persons interested in or working on problems of inter-
planetary voyages received invitations to take part in the
organization of an exhibit. In a short time, many scientists and
inventors sent in manuscripts, plans, drawings and models. The
exhibit was held in Moscow, at No. 68 Tverskaya Street (now
Gor'kiy Street) and was quite popular. The main portion of the
exhibit consisted of the following sections: astronomical,
aviation and air flight, science fiction, where the works of
Jules Verve and H. G. Wells were presented, science-realistic, a
significant portion of which was dedicated to the creativity of
N. I. Kibal'chich, then an inventors section, in which the
central position was occupied by materials describing the crea-
tivity of K. E. Tsiolkovskiy. The final, design section,
presented plans for rockets of various types and their methods
of flight. Here also were exhibited models of rockets and
rocket apparatus designed by K. E. Tsiolkovskiy, F. A. Tsander,
A. Ya. Fedorov, G. A. Polevoy (USSR), Eno-Pel'try (France),
Goddard (USA), G. Oberth and Max Walier (Germany), F. A.
Ulinskiy (Austria) and other. The exhibit was constructed so
that it was all located around the Tsiolkovskiy section, the
center of the theoretical division.

A significant role in the development of reaction motion was
played by the scientific society and research orgniazations of
'eningrad. For example, a scientific research section on
interplanetary voyages was set up at the Leningrad Institute of
Railroad Engineers on the initiative and under the leadership
of N. A. Rynin. This section considered its main task the
detailed development of problems related to "reaction flight."

The section held sessions quite regularly, discussing
various problems of space flight. At one session, on 25 February
1929, Ya. I. Perel'man read a report, noting the first practical
steps which needed to be taken by members of the section: namely
the construction of rockets with engines burning liquid fuel
(petroleum and its derivatives); experimental launchings, /163

u ginning with small powder rockets and gradually going over to
more powerful rockets, in order to end this stage with the
creation of a "stratosphere rocket," capable of reaching an
altitude of 100 km or more. Professor N. A. Rynin took active
part in the work of this section.

In connection with the organization of RNII in 1933,
absorbing Moscow GIRD, the public activity of the latter

139

organization was continued b. the reaction group of the
Military Scientific Committee of CS Osoaviakhim, founded in
January of 1934. The reaction group, soon reorganized as the
reaction section, was subordinate to the Military Scientific
Committee in the Osoaviakhim system.

On 6 January 1934, the first meeting of the reaction group
was held, headed by I. A. Merkulov.

From 31 March through 6 April 1934, the reaction group,
on the initiative of the Academy of Sciences USSR, held the
first All -Union Conference on the Study of the Stratosphere in
Leningrad. Primary attention was turned at this conference to
problems of the creation of high altitude rockets.

In 1935, the reaction section held the first USSR con-
ference on the application of rockets and rocket planes for the
study of the stratosphere. In 1935-1937, exhibitions on rocket
technology at the planetarium, Central Park of Culture and
Rest imeni M. Gor'kiy and Central Hall of the Red Army were quite
successful.

The lectors group, created in the reaction section read
several hundreds of reports on reaction motion and interplane-
tary voyages during the time of its existence. The work of the
section was publically supported, and involved the active
participation of A. I. Polyarnyy, I. A. Merkulov, L. S. Dushkin,
0. S. Oganesov, L. E. Bryukker, G. V. Overbukh, as well as
professors B. S. Stechkin, V. P. Vetchinkin, F. I. Frankel ' ,
A. V. Kvasnikov, K. L. Bayev, B. M. Zemskiy and others.

Between 1935 and 1938, the reaction section published three
collections on "reaction motion," including the articles of
domestic scientists -- K. E. Tsiolkovskiy, M. K. Tikhonravov,
V. I. Dudakov, Ye. S. Shetinkov, V. S. Zuyev, I. A. Merkulov,
F. D. Yakaytis, N. G. Chernyshev and others. A textbook on
LRE was prepared. The USSR's first textbook on the design of
liquid-fueled rocket engines was written by Ye. K. Moshkin and /164
published in 1947. It was used in many higner educational
institutions in the country for some 10 years.

One of the primary conditions resulting in the successful
activity of the reaction section of CS Osoaviakhim was the
scientific leadership of the leading scientists of RNII.
Especially helpful were G. E. Langemak, M. K. Tikhonravov, V. P.
Glushko, S. P. Korolev, Yu. A. Pobedonostsev and A. P. Vanichev.

The reaction section also performed design development.
For example, in the fall of 1934 under the leadership of A. I.
Polyarnyy, a weather rocket with LRE was planned. The fuel com-
ponents used were ethyl alcohol and liquid oxygen.

140

The first test launching of this rocket was conducted in
1935. The rocket was later modernized and began to be called the
R-06. The first successful launch was conducted on 11 April
1937 near Moscow; six more launches were subsequently conducted.

The design data of this rocket are as follows: length with
stabilizers l".b45 m; diameter 0.126 ra; launch weight 9-iO kg;
dry weight 6.5-7 kg; engine operating tiwe 11 sec; engine Thrust
40 kg; design vertical flight altitude 4.5 km; velocity upon
leaving launch support -- 21 in/ sec.

The R-06 Rocket

The "Aviavnito" Rocket

141

In this rocket, liquid oxygen was supplied from the tank to
the combustion chamber under its vapor pressure; alcohol -- under
compressed nitrogen pressure, with nitrogen occupying 65% of the
volume of the alcohol tank. The LRE was made of stainless steel
and cooled externally by its fuel. The fuel was electrically
ignited. The nose portion of the rocket was opened when the
proper altitude was reached and a parachute was ejected from the
nose, returning the rocket smoothly to Earth. The tail section
of the rocket carried four stabilizers.

In 1937, the reaction section created a second rocket with
LRE, also burning ethyl alcohol and liquid oxygen, but water was
sprayed into the combustion chamber to improve cooling. This
naturally reduced the specific impulse and increased the weight
of the rocket.

In 1938-1939. the reaction section planned the first
Soviet two- stage rocket. It was made and tested under the
leadership of I. A. Merkulov. The first stage had an engine /166
which burned a solid fuel -- smokeless powder. The second stage
utilized a*i air breathing reaction engine (ARE). The launch
weight of the rocket was 7.07 kg, the first stage weighing 3.51
kg, the second stage -- 3.56 kg.

The first successful launch was conducted 5 March 1939.
During a flight on 1 September 1939, the engine of the first
stage lifted the rocket to an altitude of 625 m, and achieved a
flight velocity of 105 m/sec. After this, the first stage was
separated by aerodynamic braking and the ARE of the second stage
was ignited. It lifted the rocket to 1800 m; the rocket achieved
a velocity of 224 m/sec. In 1939, these rockets were launched
16 times. All launches were conducted from a special vertical-
type launch support unit with four guides.

During this time, there was yet another reaction section
in Moscow, a part of an independent organization called the
Stratosphere Committee of the All-Union Aviation Scientific
Engineering and Technical Society "Aviavnito." This public
organization was also involved in the study of the stratosphere
and the development of the problem of reaction motion.

The reaction section of Aviavnito was involved in scientific
and technical propaganda and the development of a rocket. trans-
ferred there from RNII, called the 05 rocket until 1935 . At
first, the rocket carried the ORM-50 engine designed by GDL, and
utilized nitric acid and kerosene as fuel. In the reaction
section, the 05 was renamed the Aviavnito rocket, and a type 12K

oxygen engine was installed.

Planning of the 05 rocket was begun in the second team of GIRD
under the leadership of M. K. Tikhonravov.

142

The Aviavnito rocket had a streamlined shape, the nose por-
carried a parachute and pyrotechnical device, while the tail por-
tion carried the engine and equipment. The middle portion of the
rocket carried four tanks made of duralium tubes: two tanks for
ethyl alcohol and two for liquid oxygen.

The design data of the rocket are as follows: length 3.2 m,
maximum diameter 0.3 m; launch weight 97 kg; dry weight 64.8 kg;
fuel weight 32.6 kg (ethyl alcohol 13.4 kg, liquid oxygen 19.2
kg); engine thrust 300 kg; engine operating time 21 sec; flying
altitude 10 km.

The walls of the engine were protected from overheating by /167
a ceramic lining, consisting of a mixture of magnesium oxide and
aluminum oxide. The fuel was ignited by an electric spark plug.

The first launch was conducted 6 \pril 1937, the second --
on IS August 1937. During the second launch, the rocket climbed
smoothly upward, after which it lost stability and began to
descent rapidly with the engine operating. The rocket utilized
parts and assemblies from earlier rockets. A launch support
48 m high was constructed to launch the rocket.

The reaction section of Aviavnito planned two more liquid-
fueled rockets. One at a maximum design flying altitude of
40 km, the other --65 km. Subsequently, work was continued
only on the second plan, but the rocket was never built due to
the lack of sufficient funds.

Interplanetary and reaction sections and groups were devel-
oped in many higher educational institutions and other organiza-
tions.

For example, in 1930 a student aviation builders club met
at the Polytechnical Institute imeni M. I. Kalinin.

In 1938 a reaction section was organized at the Moscow
Institute for Mechanization and Electrification of Socialist
Agriculture (MIMESSKh) , involving some 50 students in the senior
classes. One result of the work of this section was a plan for
a motor vehicle with an LRE.

The beginning of the Great Patriotic War hindered the
continuation of experimental work. After the war, an engine was
constructed and utilized in certain higher educational institu-
tions for a number of years for the performance of scientific
research and laboratory work.

The society was very effective in its work of scientific and
technical propaganda, publication of scientific literature and
training of engineering and technical workers in the area of
rocket technology.

143

/lt»9

A group . -.'. -'•' ■• . k«, •- it tin* StUli '* adings:
(h'ei . r i j> h 1 1 V. I . ,-Uek: uul \i, V. M.
C.alkovskn ' . \ I'ikl mm v, U*. J Matysik,

0. I - Pan*' ,,.i, Yu. ,\. I'ii l< *• -.•*-, Ye. k.

kin lul I , •■-•',- v

hi l«>(i*l, v* ' '• tiii • ■ ', > * ', 'orms - members of" Moy cow f.JRD,
a U1R0 worker'- •. ' - • • r% -• for ivo years, the group

worked fruit 11? on - •• >epulari itioi M re \ef technology,

facilitating t m erg; m at ion i -"■ -\ el pniom -• .' museums and
exhibits, and worl I e In col h •' i ;:. • •. ten ti ration of
.ii «. iu v e d.v wment -■

Based on thi i uij i . the '•■ p oi V- terans of Rocket

Techno logv of th Ii K kiis cr ited, i .'•--'•:♦ in addition to the

tUUO workers, tht ' orl r at in I let '• iMet organisations.

Young spocia . work of the group.

144

Memorial Symbol in
Honor of 40th Anni
versary of the
Group for Study of
Reaction Motion

the meeting, repor
Raushenbakh, I. A
others.

The chairman of the group is Yu. A.
Pobedonostsev. At a general meeting o**
the veterans, honored members of the
bureau of the group were elected:
outstanding Soviet scientists in the area
of rockets and space technology. The
following sections are included in the
group: organizational, editing-
publishing, propaganda and agitation,
youth work and cooperation with museums
and exhibits.

On 18 November 1971, the group of
veterans held a Creative Jubilee Meeting
dedicated to the 40th anniversary of the
organization of GIRD in the USSR. At
ts were read by Yu. A. Pobedonostsev, B. V.
Merkulov, Ye. K. Moshkin, B. M. Matysik and

145

Without doubt, the organization
of this Institute became possible
onlg due to the conditions
created by the struggle of the
Soviet working class under the
leadership of the Communist Tarty.
I. T. kleymcnov

Chapter 3. The Reaction Scientific Research Institute (.RN11) /170
3.1. Creation ot" the Institute

In 1931, the administration of GDL, and beginning in 1932
the leaders of Moscow GIRD and Leningrad GIRD repeatedly put forth
the suggestion that the world's first State Scientific Research
Institute for Rocket and Space Technology he created. The leading
specialists in the area of rocket technology knew clerkly that
successes on a statewide scale could he achieved only by concen-
tration of the forces of scientists in a large scientific research
and experimental-design organisation.

This suggestion was supported by the Deputy Commissar for
Miliary and Naval Affairs, M. N. Tukhachevskiy , and an order of
the Revolutionary Military Council of the USSR of 21 September
1953 called for the organization of the Reaction Scientific
Research Institute (RNII) as a part of the People's Commissariat
for Military and Naval Affairs. The new institute was based on
OUL and Moscow GIRD.

A resolution of the Council for Labor and Defense of the
USSR No. 4 dated :>l October 1933 transferred RNI1 to the People's
Commissariat for Heavy Industry, and as of 4 April 1934 it was
directly subordinated to the Scientific Research Section of this
Commissariat. The Chief of RN1I was Ivan Terent 'yevich Kleymcnov,
his deputy was s. P. Korolev until January of 1934, after which
C.eorgiy lirikhovich t.angemak took over.

ivan Terent 'yevich Kleymcnov (1989-1938) was one of ti.e
organizers and leaders of work on rocket technology in the USSR.
In 1932-1933, he was the Chief of the Gas Dynamics Laboratory,
in 1933-193? -- Chief of RNII. His name has been given to a /JU
crater on the far side of the moon.

Gcorgiy lirikhovich Langemak (.1898-1938) was a Soviet

artillery engineer, the designer of rocket weapons burning
smokeless powder. He was one of the principal leaders of the
development of rocket weapons at GDI. and RNII, including those
later used in the Katyusha rocket launcher. In 1934-193™, he
served as Deputy Director and Chief engineer of RNII.

146

■•/ • Chic i --. RNII,
Ivan ■•••:.• . evich
." y <■ •-.ay

, ; , i tu we •• : n ettiai I n \ n i

to a crater on » fa) • . • ot tin n< on.

>ui ' ; tli : I ■.'. • jxistence o this
inst : - . : , its stru tur , »ik3 t h ' m : of

it*- ubdi vi; . ■ • re c ,-/ g • ! r< > a tod ty.
During the initial period/ the institute
co-' I of I eet \ • . • v . $

co- • ■ • .'. H sectors, tl .i rs being

Ji ktecl nto tcai . Later, th ?
in - ' .' • > was divideJ tito gros ps. Fur-
th . rm< re, the ir tititt* :J .:•. ed • >; s?ri-

raei :a pr< - ; :,- r i< n tc 1 ities, i sb< •- tto ies
an<. - .• ■. •- . • at •-■ - si< »$.

The first section of the institute
st d - 1 d , tier- £uel« d r< . •' - ; ind
la--.-;- ; ;.•••• L ton: foi * hem, i.e. ,

CO, 1 --. : '.-.-• : ht • . :"- s S I i Otl

wo - ; dl '•' Rl he nl • ' -acid t< m Jk el-

op- ocl t engines \ Hi ins nonvo (tile

o. I - - I • - - -. ihe ox t tcai

dL •. pod rocket "• • to s burning • • •: ' 1

oxygt (i and e i h) '. al< o.hol.

The l lirtl »nd f< uri I section de srel -
oj " >' aged re kc t ui breath i •• < .. -
ti- - mginc snd *th< t de\ ices.

' 5 - he >",' ul is oi ; i .='.', ts
<** .. * : . ;, - id i U a ' v/ it all t ,- o u • - »f
0IR1 • i t) f the it rl c r s c GPL, id o
vc-: ■ .•• j * .-• I from . in : rad . ti 'Ing
it.*» I i - t rri >j f '; of < i.st <"=•. e, th<
institute hi red ne$ •.•'".;.--.• el 1 .

• • in I ::■ i t \ in gi min o * ;• ••• f iv-
it , : , ' • . "' i - - I • pes .en

wi I h ' In f ••;•■ ler o astronaut ic: K\ U.
Ts o3 ov5kiy I woi •• r< i • hist i -

tu1 risit [oust am i ! luais -.-. ich
rep ■ edl i In! it : :oi jpondenee with
him, ut : Is fvl his consuli at i on in - ;e •:

Geori , ..--.-- -h w k to hii . review.

Langetnak

On } chi w ■ L934, - • a general

meeting of the workers ot" RNJ1 dedicated to the 15th anniversary

of the '; I krmy $ K. E, T: iolko . ; , w;t? > -,-•' .- an h sc •• y

merohei 3f * ,. * fechnical Conn . ." the [nsi tt f,

R« ognizing th< service? ; • . * . * I > >] thi '■ nienl
Coun* .- : >f RNII iliaed th name Tsiolkoi ki •- inula '«•• the

/172

14'

basic equation for the flight velocity of a rocket and "Tsiol-
kovskiy nuiaber" for the ratio of the mass of fuel reserve to the
final mass of a rocket, suggesting that this ratio be represented
by the letter "Ts."

3.2. The Activity of the Institute

The institute developed scientific research and experimental
design work on solid-fueled rocket engines, LRE and flight /173
vehicles, most of which had been begun at GDL and GIRD.

Powder Rocket Weapons

Powder rockets of various types and launch installations were
developed at RNII under the leadership of G. E. Langemak by
subdivisions headed by L. E. Shvarts, K. K. Glukharev, I. I.
Gvay, V. I. Aleksandrov and others. During the initial period of
existence of RNII, the workers of the institute were aided in the
solution of many theoretical problems by scientists from the
Artillery Academy such as D. A. Venttsel', M. Ye. Serebryakov,
I. P. Grave and others, who took part earlier in the work of GDL.

The search for the most effective and economically favorable
types of solid fuel (powde*) for various models of reaction
weapons was conducted in the powder shop, which was first located
in Leningrad, but was transferred to Moscow in the first quarter
of 1936.

Problems of the theory of interior and exterior ballistics
of powder rockets were also studied in the late 1930' s by Yu. A.
Pobedonostsev, M. K. Tikhonravov, M. S. Kisenko, V. G. Bessonov
and others.

By 1934, work had been widely developed on the creation of
solid-fuel rockets of various sizes both for field artillery and
for anti-aircraft purposes.

Solid- fueled rockets differ from rockets with LRE in their
simplicity of design, high reliability, safety for the users and
convenience of operation. Furthermore, the level of technology
achieved in the 1930' s was quite sufficient to support rapid
development of mass production of solid- fueled rockets.

In July of 1937, the RS-82 air-air and air-ground reaction
devices were fired. The military tests of the RS-82 were com-
pleted in November -December 1937 with group firing against surface
targets on a training firing range from 1-15 aircraft.

Late in 1937, the RS-82 was adopted for armament of the 1-15
fighter. The airborne launchers were developed by A. P. Pavlenko

148

and N. G. Belov. Improved launchers were later developed by I.I.
Gvay, A. S. Popov and others. In July of 1938, military tests /174
of the RS-132 missiles, to be installed on bombers, were con-
ducted. The tests were successful, and the RS-132 was also put
in military use.

Air-to-air powder-fueled missiles were used in combat for
the first time on 20 August 1939 by Soviet troops fighting the
Japanese militarists in the region of the Khalkhyn-Gol River,
when 5 1-153 fighters ("Chayka") , each armed with eight missiles,
attacked a larger detachment of Japanese fighters.

The five first Soviet missile planes were led by test pilot
Captain N. I. Lvonarev. This group, on five missions, shot down
10 fighters, 2 heavy bombers and 1 light bomber, without losing
a single aircraft.

In 1938, RNII began working on a surface launcher for the
RS-132 missile. The first models, with a capacity of 24 missiles,
were mounted across the chassis of a truck. In the summer of
1939, considering the experience accumulated, a 16-missile
launcher with guides directed along the chassis of a three-axle
truck, was created. By late 1940, RNII had constructed six such
installations. The missiles were fired, after jacking up the
vehicle, in the forward direction, and the launcher was loaded
from the rear. These devices, developed by engineers I. I. Gvay,
V. N. Galkovskiy, A. P. Pavlenko, A. S. Popov and others, were
prototypes of the BM-13-16 or Katyusha launchers.

A resolution of the State Defense Committee calling for
series manufacture of rocket launchers was signed in June of 1941.

lne BM-13-16 launcher was first used in combat on 14 July
1941 in the battery of Captain I. A. Flerov, a graduate of the
Artillery Academy imeni F. E. Dzerzhinskiy. The German fascist
troops occupying the railroad station at Crsh were quite sur-
prised by a barrage of uncommon force at 15:30 hours. The entire
station went up in flames, and powerful explosions went off one
after another.

During the years of the Great Patriotic War, combat rocket
launchers were used successfully in massive numbers, carried by
wheeled and tracked vehicles as well as combat aircraft.

The rocket artillery fully confirmed its high combat
qualities -- mobility and maneuverability, the capability for
sudden concentration of fire at high densities over large areas
with a rapid rate of fire.

149

Liquid- Fueled Rocket Engines

As we have noted, the second section of the institute worked /17S
on the study and development of LRE.

The nitric acid team, headed by V. P. Glushko, continued the
study and development of LRE begun at GDL, using nonvolatile
nitrogen-containing compounds, primarily nitric acid with oxid» s
of nitrogen, as well as tetranitromethane, as oxidizers.

Between 1934 and 1938, this team developed engine models
from ORM-53 to ORM-71, plus the ORM-101 and ORM-102.

The primary task of this team was the creation of rocket
engines and supplementary devices. Considerable attention was
also given to problems related to the use of promising materials
such as stainless, heat resistant, aluminum and other materials.
New methods of welding and soldering were introduced, and experi-
ments were conducted on increasing service life by chrome plating
of worn surfaces. Since the engines were designed for both
manned and unmanned flight vehicles, one important task performed
by the team was reduction of the period required to reach nominal
operation and automation of the launch.

The ORM-53 through ORM-63 engines were planned in 1934 and
developed in 1935, followed by the ORM-64 and ORM-65.

The ORM-65 engine successfully passed adjustment and official
testing in 1936, followed by surface testing on the RP-318 rocket
plane and the 212 winged rocket in 1937-1938.

In 1939, the ORM-65 engine passed flying tests on the 212
winged rocket quite successfully and was highly evaluated.

After processing of a great deal of experimental data and
conducting a series of scientific research operations in 1936-
1938, the team developed the ORM-66, ORM-67, ORM-68, ORM-69 and
ORM-70 engines with higher characteristics.

Furthermore, the team created various systems for LRE:
turbine pump units, gas generators, automatic control elements,
etc. In 1935-1936, for example, the first domestic gas
generator, the GG-1, designed for production of the working
fluid for the TNA turbine, was developed under V. P. Glushko.
This gas generator passed official interdepartmental tests /176
successfully in 1937.

In 1939, V. P. Glushko was made the leader of an independent
subdivision of the Aviation Motor Plant, separated from RNII.
Therefore, the work of V. P. Glushko at RNII ended in 1938.

150

By 1939, after further vesting, the RDA-1-150 liquid-
fueled rocket engine for the RP- 318-1 engine was planned on the
basis of the ORM-65 engine, under the leadership of L. S.
Dushkin.

However, the RDA-1-150 engine, due to its low thrust, was
found to be unsuitable for unaided takeoff of an aircraft.
Therefore, a more powerful nitric acid engine, the RDA-300, was
planned and manufactured in the first half of 1939. During this
same year of 1939, the RDK-1-150, burning alcohol and oxygen,
was created.

The oxygen team, headed by M. K. Tikhonravov, developed
engines burning liquid oxygen and an aqueous solution of ethyl
alcohol. Means were sought to assure the most complete possible
combustion of the fuel and increase the -hernial efficiency. The
results of theoretical and experimental work indicated that this
required an increase in combustion chamber pressure. Therefore,
even chough increasing the pressure complicated the cooliag
problem, new engines were designed for combustion chamber
pressures of around 15 atm, in place of the 5-8 atm used earlier.

Several versions of the 12K engine were first tested; in
1936, the oxygen team began development of the 205, 206, 207 and
208 engines, designed, like the 12K, for installation in
rockets. The technical assignment for planning of the engines
noted the need to eliminate the shortcomings of alcohol -oxygen
engines developed earlier. It was also required to increase the
reliability end reproducibility of test results and the j^Gcific
impulse.

In early 1934, the group of L. K. Korneyev, working on the
development of GIRD engines in order to increase the reliability
and reproducibility of test results, was separated from RNII.
Some of its workers later took part in tht v.~ k of Design Bureau
No. 7 (KB-7), organized as a part of the Main Artillery Adminis-
tration of the Red Army and headed by L. K. Korneyev.

Air-Breathing Reaction Engines

In 1934-1935, RNII performed experimental work with direct
flow air-breathing reaction engines (PVRD) . Preliminary calcu-
lations and testing of PVRD models were performed at GIRD in 1932-
1933. The experiments performed at RNII confirmed that PVRD,
based on the theory of B. S. Stechkin, were suitable for use for
flight at supersonic speeds. This work was performed under the

_

The symbol RDA- 1-150 stands for "Rocket Engine, Nitric Acid,
No. 1, Thrust ISO kg."

151

leadership of Yu. A. Pobedonostsev, with the participation of
M. S. Kisenko, A. V. Salikov, I. A. Merkulov, U. S.
Oganesov and A. B. Ryazankin.

PVRD Installation Designed by I. A. Merkulov on
an Aircraft

In 1936-1939, the institute studied pulse jet engines. How-
ever, these engines were not further developed.

In 1937 1940, under the leadership of V. S. Zuyev and Ye. S.
Shchetinkov, PVRD models were tested. Preliminary work on
improvement and development of experimental methodology was per-
formed using hydrogen fuel, after which an extensive PVRD testing
program using devices burning gasoline was undertaken. Based
on the experience accumulated, V. S. Zuyev designed a ram jet to
be installed on an aircraft.

In 1942, flying tests of the jet engine designed by M. M. /178
Bondaryuk were conducted on an LAG-3 aircraft. At this time, the
Design Bureau was still not a part of the institute. Later, in
1946-1947, a ram jet engine for subsonic speeds was developed at
RNII under the leadership of M. M. Bondaryuk. It was designed to
be used as an accelerator by the LA-7 and LA-9 aircraft. In 1948-
1950, a dual-loop aircraft PVRD was developed.

Late in 1944, an experimental turbojet en^'ne, the S-18, was

developed at RNII under the leadership of A. M. Lyul'k. Subse-
quently, the experience gained in working on this er.gine was used
as the basis for the plan for the Soviet turbojet engine (TRD) .
which posses state testing in March of 1947. This work served as
a basis for the development of air-breathing reaction engines in
the USSR, which engines have been widely used by various aircraft
since the war.

152

Flight Vehicles

RNII, under the leadership of S. P. Korolev, continued work
on winged rockets -- air torpedos -- with both solid- and liquid-
fueled engines, following the work undertaken on his initiative
at GIRD. Preliminary calculations of the flight stability of
winged rockets were performed by Ye. S. Shchetinkov and A.
Mar kin under the leadership of A. V. Chesalov. The first rockets
with LRE, called the 06 rockets* were flight tested in early
1934.

During the process of work on unmanned winged rockets,
several flying versions of the 216 winged rocket with the 02
alcohol -oxygen LRE were created, then (1936) the improved 212
rocket, with the ORM-65 engine. An extensive program of stand
testing of the various units of the engine and 212 and 301 rockets
was undertaken, followed by flight testing of improved versions.

The 212 rocket, an all-metal device, consists of the follow-
ing sections: nose section, carrying the pay load and parachute;
instrument section, for the stabilization and control system
apparatus; fuel section, carrying the tanks; nitrogen section,
carrying the pressure cylinder; and the engine section.

The fuel and oxidizer tanks, tubular in shape, were located
within the wing. The fuel components were fed to the combustion
chamber by compressed nitrogen pressure. The pressure reducers /179
for the nitrogen which was fed to the tanks from the pressure
cylinder and the fuel valves were located at the plan center of
the rocket.

The ORM-65 engine was carried in the tail portion of the
rocket on a frame and covered by a fairing with a metal sleeve
located above the nozzle exit plane to protect the rudders from
the flame.

The device was launched from a catapult truck powered by a
powder- fueled rocket, the combustion chamber of which contained
packets of trotyl pyroxene powder (IS packets measuring 75 x 10
x 92 mm). The catapult truck rode on rails ISO ra in length.
The takeoff run required for the winged rocket during flight
tests was 26 ra.

The planned flight range of the winged rocket, with a launch
weight of 210 kg and a fuel reserve of 30 kg, was 50 km.

The design of the RP-318 rocket plane was as follows: wooden,
free flying monoplane, fuselage of oval cross-section with mid-
section area 0.7S m z ; length 7.44 m, wing span 17 m, bearing
surface of wing 7.85 ra 2 . Initial flying weight 700 kg. Launched
from Earth as normal for gliders.

153

The steel fuel tanks located behind the Metal back of the
pilots seat carried 7S kg of fuel, sufficient for 100 sec con-
tinuous operation of the engine at a thrust of 150 kg. The
capacity of the fuel tank, located directly behind the pilot's
seat, was 20 i, while the two oxidizer tanks, located at the
center of gravity of the aircraft, had a capacity of 40 t. In
case of leakage, the oxygen tanks were contained in duralumin
baths with a drain leading outside the aircraft. The oxidizer
and fuel were fed to the engine under compressed air pressure,
with the air carried in four tanks of 5 liter capacity, two 4 n
each wing. The air was fed to the fuel tanks through a pressure
reducer. The engines were started by rotating a control lever,
which Mechanically opened the fuel valves located in the tail por-
tion of the fuselage iMMediately before the engine. The fuel
valves were opened when a signal laMp installed on the pi lot* s
instrument panel lit up.

The RP-318 rocket plane designed by S. P. Korolev was
tested with LRE about 40 times.

The engine was carried on a frame in the tail of the
fuselage and Mounted beneath a metal shield to protect the tail
section from the flame. For this same purpose, the portion of
the rudder closest to the engine was covered with a sheet of
stainless steel 0.3 mm thick.

During the pre-war period of activity of RNII, almost all
of the creative workers and specialists in the area of rocket
technology labored within its walls. The principles of the theory
of rockets and engines were developed, operating models were
created, which later saw practical application and development.
RNII made a significant contribution to widely varied areas of
rocket technology, thus providing a reliable foundation for
Soviet rocket science.

3.3. Nitric Acid LRE

The ORM-53 - ORM-63 Engines

The nitric acid team worked on the creation of engines,
utilizing the last LRE of the Gas Dynamics Laboratory, the ORM-52,
as a prototype. The basic fuel components utilized in the
engines developed, as before, were nitric acid and kerosene.
Summarizing the experience of the work of GDL, the designers came
to the conclusion that the reliability of engine starting in all
nozzle positions would have to be improved, by using chemical
and pyrotechnical ignition, that the fuel feed system would have
to be developed to bring the engine up to full design thrust
more rapidly, that the operating time of the engine would have
to be increased, as well as the specific impulse, by improving

1S4

Mixture formation. In order to decrease the weight of the engine,
the feed pressure had tc be reduced by improving the hydraulic
characteristic* while conserving the same pressure in the combus-
tion chamber.

Taking these initial ideas, RNII developed a series of
engines from 0RM-S3 through ORM-63 in 1934-1935.

In the ORN-53 engine, a number of design elements were
improved over the ORM-52. The ORM-S4 had external cooling of
the nozzle by the oxidizer and higher spiral ribbing; the spray
head and combustion chamber, as before, were protected from the
effects of the high heat fluxes by an internal film (vapor cur-
tain).

The ORM-57 8- sprayer high-thrust engine had a critical »o»zle
cross section diameter of 40 mm, an exit plane diameter of 100
mm, with a cone aperture angle of 20*. The aluminum nozzle insert
consisted of 6 parts. This engine was planned but not manufac- /181
tured. The first domestic two-chamber engine was the 0RM-S8,
designed for a thrust of 600 kg.

Summarizing the experience gained in planning the engines
from ORM-53 to ORM-62, the designers selected the best features
and created the ORM-63 engine.

The ORM-63 was a fully cooled experimental engine developing
a thrust of 300 kg. It underwent element-by-element technological
development in production in order to assimilate a number of new
technological operations: roller electric welding of the compen-
sator, stamped from a sheet of stainless steel, to the nozzle and
its jacket, butt electric welding at the critical cross section of
the nozzle, high temperature hermetic soldering of various joints
with high- temperature solder, etc. Particular attention was
given to the quality of manufacture of parts, testing of sub-
assemblies and the quality of assembly of the entire engine.

The combustion chamber of the CRM-63 utilized membrane-type
hydraulically controlled spiral sprayers. The corrugated mem-
branes were stamped of sheet stainless steel.

The ORM-64 - ORM-70 Engines

In early 1936, tactical and technical requirements were
developed for an engine for use in the RP-318 rocket plane and
the 212 remote controlled winged rocket. The engine was to
develop a thrust of 150-160 kg, to operate continually for at
least 75 sec per start and develop a specific impulse of at least
180 sec; its weight was limited to 10 kg. The variation in
mean thrust from start to start of the engine during the period

23 a

of stable operation was not to exceed ±3 kg; the difference
between values of mean thrust and aaxiaua and miniaua thrust
during a single start of the engine during the period of stable
operation should not exceed ±3 kg; the fuel feed pressure was
not to be over 35 ata. .he engine should operate normally in
the horizontal and vertical position, and also with the inlet
pressure choked froa 35 to 12 ata by aeration of fuel flow rate.
Particular attention was given to the assurance of high relia-
bility of starting and operation. According to these requi re-
sents, the ORM-64 engine, as an experiaental version, and the
ORM-65 engine, as the basic operating version, were planned,
constructed and tested in 1936.

/182

The ORM-64 Engine

was 10 kg. With a pressure
and a feed pressure of 27.5
iapulse of 216 sec.

The ORM-64 was an experiaental
engine with a thrust of ISO kg,
siailar in design to the ORM-52
e* line; it was a four -sprayer
engine, coabustion chaaber voluae
2.23 *, diaaeter of nozzle critical
cross section 20 aa, exit plane
diaaeter 40 aa, nozzle expansion
angle 20°. At the center of the
head was a device for ignition con-
sisting of a sleeve carrying a
current conductor (ES-Kh sparkplug),
an electric cap and a 6-8 second
aetal -nitrate ignition cap, seated
on a rod. The aaterial of the
chaaber was carbon steel, the nozzle
was aade of EYa36 steel.

During test stand operation of
the ORM-64 engine in the vertical
(nozzle downward) and horizontal
positions, the required technical and
technical characteristics were
achived, including the weight, which
in the combustion chaaber of 22.5 ata
ata, the engine developed a specific

/183

The coabustion chaaber operated for a total time of 502 sec
without defects; start-ups were shock- free, the engine operated at
its design mode stably, without oscillations. With a continuous
engine operating time reaching 120 sec, the cylindrical portion of
the coabustion chaaber, due to the intensive process of fuel coa-
bustion, glowed bright yellow. This was due to the fact that the
combustion chamber had not external cooling, the coabustion chaaber
walls being cooled only by the spraying of the fuel components on
its inner surface. In later designs of ORM, in order ^o assure
higher reliability of the cylindrical portion of the coabustion

156

chaaber, it was cooled by a flow of nitric acid on the outside.
Based on analysis of the results of these tests, the sain ver-
sion of the engine, the ORN-65, was developed, and successfully
passed official stand tests in 1936, also in the vertical
(nozzle downward) and horizontal positions. The ORM-65 engine
was the aost highly developed engine of its tiae.

The aain data produced in the tests of 1936 were superior
to the assigned tactical and technical requireaents, except for
weight, and were as follows.

Thrust at ground level in aaxiaua aode 17S kg, in noainal
■ode 155 kg, in ainiaal aode 50 kg; specific impulse in aaxiaua
aode 195 sec, in noainal aode, average aode for the entire tiae
of stable operation, 215 sec; coabustion chaaber pressure in
aaxiaua aode 25 ata, in noainal aode 23 ata and in ainiaal aode
8 ata; fuel consuaption in aaxiaua aode 0.900 kg/sec, in noainal
aode 0.738 kg/sec. Method of start-up aanual on signal laap or
autoaatic.

The ORM-65 coabustion chaaber, with a voluae of 2.01 t, con-
sisted of three steel aain parts: the spray head, chaaber nozzle
and jacket, sealed with an asbestos liner. The chavber head,
designed to prepare the fuel for coabustion, with internal
fila cooling, had an operating surface teaperature of 300-400° C.
The coabination chaaber and nozzle consisted of the cylindrical
portion of the coabustion chaaber, made in one piece with the
nozzle. It was equipped with external flow cooling; in order to
increase heat transfer, the chaaber -nozzle had spiral ribbing in
two places. The pressure drop through the cooling fluid line was
3.5 ata when operating in noainal aode.

The necessary jacket gap at the nozzle was provided by the /186
installation of two shaped aluminum inserts.

The nozzle was equipped with a coapensator --a lead liner,
held under pressure by a threaded ring. This coapensator allowed
theraal elongation of the chaaber and nozzle relative to the
cooler jacket (with the lead flowing into the circular cnp
between the jacket and chamber-nozzle), while maintaining the
tightness of seal. After each test, the pressure ring hid to
be tightened up to restore the seal.

The fuel coaponents were sprayed into the coabustion chamber
through centrifugal -type sprayers (three oxidizer sprayers and
three fuel sprayers alternating at intervals of 60°). The
oxidizer sprayers were installed in the head portion of the
chaaber at an angle of 60° to the axis and directed opposite to
the nozzle. The fuel sprayers were installed in the head
perpendicular to its axis.

157

jm2

Cross -Section of the ORM-65 Engine

starting device. When the
fully* which was the signal
start or put the automatic
distance of the shunt from
that it would burn through
be well ignited.

The ignition device
consisted basically of a
current conducting plug,
cartridge with an electric
cap and pyrotechnical
igniter (metal -nitrate) cap.

When the ignition
circuit was closed, the
wire in the electric cap
burned out, igniting the
charge of smokeless powder
in the cartridge. The
hot powder gasses, flying
out through the channels in
the cartridge, ignited the
cap. The ignition cap,
which burned from one end,
was cylindrical in shape,
24 mm in diaaeter and 40 mm
long with a central inner
channel protected by a
duralumin tube. The fuel
components were fed to the
chamber only after good
ignition of the igniter cap.

This was achieved by
connecting a low-resistance
shunt, which passed through a
hole drilled in the side of
the cap, in the ignition
circuit in parallel to the
control lamp installed on the
control panel or the automatic
shunt burned through, the lamp lit
to open the fuel valves for manual
start mechanism in operation. The
the end of the cap was selected so
in about 4 seconds, when the cap would

/184

The ORM-65 engines were operated repeatedly. For example,
ORM-65 No. 1 was started 49 times and operated 30.7 minutes on
the ground, including: 20 starts on the test stand (17 September -
S November 1936), 8 starts on the model 212 winged rocket designed
by S. P. Korolev (29 April-9 September 1937 and 2 October-8 Octo-
ber 1938), 21 starts on the RP-318 rocket plane designed by S. P.
Korolev (16 December 1937-11 January 1938).

/187

158

/18S

Exterior View of the ORM-65 Engine

During the first ground flame test on the RP-318 rocket
plane (.16 December 1937), ORM-65 engine No. 1 operated for 92.5
sec; during the next 26 days, 20 more test starts were conducted
The number of starts per day reached 5 (for example, 11 January
1938).

ORM-65 engine No. 2 was tested on the RP-318 and 212 16 times;
during its sixth start, it operated on the RP-318 rocket plane
during ground testing on 11 March 1938 for 230 seconds; after
adjustment operations on the rocket plane, the ground flame test-
ing of ORM-65 No. 2 continued. Between 3 February and 15
September 1938, 9 starts were conducted. This engine was started
twice during flying tests on the 212 winged rocket on 29 January

,/188

159

1959. According to the flight tests reports, the start-up and
operation of the ORM-65 engine were satisfactory.

The ORM-05 Hngine on the 111 Winged Rocket
with Powder- Fueled Rocket Accelerator

Continuing the traditions of the COL, the designers of the
ORM engines produced rocket engines distinguished by their
except ional rel iahi I i ty .

The oRM~05-A, a modification of the ORM-05, was smaller in
diameter.

The ORM- lib, an experimental engine with a thrust oi ISO kg,
was planned and manufactured in 1 9 3 1> ; stand tests were conducted
; n 1937-1938. The ORM- oo differed from the ORM-oS in that it was
lower in weight (0.9 kg) and smaller in sire, hut had increased
combustion chamber volume and a welded noz'le elongation
compensator.

The increase in combustion chamber volume and decrease in
head weight resulted in the fact that after IS seconds of opera-
tic) i at nominal mode, the head began to glow, and after 2S
sec nuls the engine had to be turned off. The head of the ORM-oO
engine was therefore improved by the addition oi fins and
ex t errs:! 1 fue 1 - fl ow coo 1 i ng .

The ORM- a? was an experimental engine with a thrust oi'
150 kg, developed and manufactured in 193". The engine used
a light-weight ignition device; the central electrode of the cur-
rent receiver had a channel used to measure the pressure in the

100

combustion chamber. In contrast to the ORM-66, the ORM-67
engine could be completely disassembled; the joints between the
head, combustion chamber-nozzle and their jackets were sealed
with asbestos strips. The head and chamber-nozzle were made of
EYa3A steel, the jackets of duralumin. The engine weighed about
S kg.

View A

/189

To

manometer
Oxidizer,

Fuel

The ORM-66 Engine

The ORM-68 (1937) differed from the ORM-67 in that the head,
chamber-nozzle and their jackets were made of duralumin, further
decreasing the weight of the device to 3.5 kg. The ORM-67 and
ORM-68 engines underwent only hydraulic testing and development
of a new ignition device in early 1938.

The ORM-69 engine was developed in 1938 and differed from
the ORM-68 in that larger, fuel -cooled ribs were used on the
head and an improved ignition device was fitted, following manu-
facture and refinement testing in early 1938.

In 1937, the ORM-70 design was developed. This was an
experimental engine with a thrust of 300 kg, burning nitric
acid-kerosene fuel. The design of the ORM-70 is similar to the
ORM-67. Eight sprayers are used. The maximum diameter of the
combustion chamber is 200 mm, the length is 500 mm. The material

/191

161

used is stainless and low-carbon steel, and duralumin. The
engine was manufactured in 1937-1938, but was never tested.

/190

of

Br

i

tw

CX3"'

II

/

503

The ORM-70 Engine

The ORM-101 - ORM-102 Engines

These experimental engines were planned in 1937 in or.ler to
study the possibility and expediency of using tetranitromethane
as an oxidizer. Corrosion testing of various metals in tetra-
nitromethane allowed structural materials stable in this
oxidizer to be selected. Experimental tests of the explosion
danger of tetranitromethane in operation were conducted. Kerosene
was selected as the fuel. The ORM-101, with a thrust of 80 kg,
was designed for brief operation. The ORM-102, with a thrust of
100 kg at the same combustion chamber pressure (28 atm) was
fully cooled. The engines were manufactured in 1937-1938, but
did not undergo flame testing due to the determination that the
determination that the use of tetranitromethane was dangerous.

The GG-1 and GG-2 Gas Generators

The gas generators (GG) developed were designed to feed the
working fluid for a piston engine or turbine. The zones of com-
bustion of the fuel components (nitric acid and kerosene) and the

162

zones of mixing with the cooling agent (water) were separate;
diaphragms were used to separate the liquid films from the walls
of the chamber. Due to the requirement for high purity of the gas
produced, preference was given to a two-chamber system, although
a one-chamber version was also constructed and underwent stand
flame testing.

In the GG-1 gas generator, the fuel components were
sprayed into the combustion chamber by 6 sprayers: three oxidizer
sprayers fed through the lower circular collector, three fuel
sprayers fed through the upper circular collector; water was
sprayed in through the two top sprayers. The gas generator was
designed for internal cooling of the walls by a protective
film of fuel components. Overheating of the combustion chamber
walls next to the sprayer belt and connecting collar between
the cambers (to 700° C) required that external flow cooling of
the spirally ribbed walls of the chamber with water, which was
then sprayed into the chamber, be used; the GG-1 p. ssed acceptance
testing in this form.

/193

/192

The ORM-102 Engine

Design and dimensions of the GG-1: material of combustion
chambers, mixing chambers and sprayer nipples -- EYa3S; of
jackets and collectors -- ST4; of sprayers and tubing -- dur-
alumin. Jackets sealed with asbestos cord soaked in liquid
glass. The GG-1 was started after a signal lamp or automatically,
with simultaneous injection of fuel components and water. In the
winter, antifreeze (75% water and 25% ethyl alcohol) was used in
place of water.

163

The GG-1 Gas Generator

generator was of high

analysis data, it con
and did not cause cor
The chemical composit
to results of analysi

at a * 0.88 with a pr

490° C, after condens
NO -- 20.21, C0 2 -- 2

nitrogen, acidity -- traces

punt

tained
rosion
ion oi
s of a

essure

ation

1.81,

The output of the GG-1 was
40-70 i/sec gas at 2U-25 atm and
4S0-580 C. Th <as generator
operated stabl) n nitric acid
and tractor kerosene with water /194
injection. The total consumption
of oxidizer and fuel was 0.15-
0.17 kg/sec, of water -- about
0.20 kg/sec. The fuel component
feed pressure was not over 30
atm. The weight of the gas
generator was 20 kg. After
1 hour 46 minutes operation, the
gas generator showed no essential
defects and was capable of further
operation. The time of contin-
uous operation was up to 15 min-
utes (determined by tank capacity).
The gas generator could operate
briefly (minutes) at gas output
temperatures of up to 700-800° C.
The gas produced bv the gas
y and was colorless; according to gas
no nitric acid or nitrogen di ^xide
of copper alloys during oper.it ion.
the gas produced by the GG-1, according
sample taken during acceptance testing
in the GG of 23 atm, gas temperature
of water, was as follows, in vol. %: /195

CO

15.2$, 0-

01, remainder

The GG-1 gas generator was developed during 1935 and 1936
and successfully passed official stand testing on 27 August 1937

In 1937, a design of a L-shaped two- chamber gas generator,

the GG-2, producing up to 100 £/sec gas ai a pressure of up to
30 atm and temperature of 450-600° C was developed. The GG-2 wa:
a further development of the GG-1 gas generator. The fuel
component feed pressure of the GG-2 was 36-40 atm; the weight of

the gas generators was less than 30 kg; the GG-2 was not con-
structed.

The RDA-1-150 Engine

The RDA-1-150 engine was developed unJer the leadership

of
fuel

L. S. Dushkin and A. V. Pallo for nitric acid plus kerosene
and was designed to develop 150 kg thrust. The centrifugal
sprayers, eight main fu°l sprayers and two start-up sprayers, were
placed on the uncooled rpherical head so that the streams of fuel

164

components were directed toward the center o ; the hemispnere, to
-he zone where the head was connected to the cylindrical portion
of the chamber. In th - ; upper portion of the head, on the axis
of the chamber, was 3 throat for the igniter device. The head
was fastened to the cylindrical portion of the chamber by means
of a thread. At the junction point there was a linear expansion
compensator gland.

Acid . rsm.m
Water —

Cross-Section of the GG-1 Gas Generator

The cylindrical combustion chamber had double spiral cooling
ribs. ihe nitric acid entered the cooling cavity at the point of
connection of the chamber to the nozzle, then passed through the
apertures in the head directly to the sprayers.

The removable nozzle was cooled by kerosene which entered
the jacket at the nozzle end and left at the point where it was
connected to the chamber. The outside surface of the nozzle
carried a doubxe spiral set of notches, the lands of which were
in tight contact with the nozzle jacket. In the lower portion of
the nozzle (at the exit plane) was a linear expansion compensator
gland.

The basic difference between the RDA-1 i>0 and the ORM-65
was the altered placement of the fuel sprayers in the engine
head. Whereas in the ORM-65 the fuel components were spra/ed
radially or at a slight angle away from the nozzle, in the

165

RDA- 1-150, all of the fuel was directed toward the center of the
chamber, toward the nozzle, and the sprayers were located around a
circle at identical angles to the chamber *xis. However, this
difference caused a significant reduction in the primary character-
istics of the engine.

/198

/196

Cross-Section of the GG-2 Gas Generator

Stand tests of the RDA- 1-150 engine began in the second
half of 1939. Two identical models were tested, and about 20
starts were made. In January of 1939, one of these models oper-
ated 200 s**c without damage. During March through September of
1939, combined tests were performed on a rocket plane, together
with the fuel feed system and the control system. During this
pe~i.d of time, the engine withstood 108 flame tests, showing
the following results: thrust 140 kg (compared to 175 kg for the
ORM-65), specific impulse with chamber pressure 18 atm reached
186 sec (as compared to 210-215 sec for the ORM-65).

As a result of the tests of the RDA-1-1S0, reliable operation
cf the engine was achieved, and the procedure for start-up, mode
control and shut-down of the engine from the cabin of the rocket
plane was developed. Exp^ience was gaine in the operation of the
engine, allowing the exp< menters to begin flight testing of the
engine following the g-tUiV. testing.

The first flight tests of the RDA-1-150 engine were conducted
by pilot V. i». Fedorov on 28 February 1940, using the RP-318
rocket plane.

166

/19:

The RDA-1-150 Engine

An ordinary aircraft with a piston engine towed the rocket
plane to an altitude of 2000 m, where the pilot disengaged the
rocket plane from the piston-engine plane, and began to glide.
After separating a sufficient distance from the tow plane, the
test pilot turned on the rocket engine, which continued to
operate until its fuel was fully expended. After shut-down of
the engine, the rocket plane continued to glide and landed at
the airfield.

This was the first manned flight of a flight vehicle with
LRE in the USSR.

The RDA-300 Engine

The RDA-300 engine was designed to develop a thrust of 300
kg, and was also intended for the RP-318 rocket plane, in order
to allow independent take-off, i.e., without requiring a tow
plane.

Th" 1 RDA-300 engine, developed in 1939 under the leadership
of L. S. Dushkin, differed from the ROA-1-150 only in its dimen-
sions. In order to increase the specific impulse to 200 sec,
the design pressure in the RDA-300 was ii creased. By the middle
of 1939, the planning and manufacture of the engine were com-
pleted. At the same time, another version of the RDA-300 was

/199

167

developed, with basic design changes based on the results of test-
ing of the RDA-1-150. The reliability of the cooling system was
increased by the use of both fuel components; the start-up condi-
tions and quality of mixture formation were improved.

The RDA-300 engine which passed
flame testing had a head which differed
significantly in design and operating
principle from the heads of all earlier
models. It had spiral sprayers, direct-
ing the stream toward the nozzle and
assuring fine atomization and good mixing
of the components. The nitric acid from
the cooling cavity of the cylindrical
portion of the chamber passed through
channels in the head to the needle-type
stop valves, which controlled two
sprayers each.

The RDA-300 Engine

Similarly, kerosene from the cooling
channel entered needle-type stop valves
for the kerosene sprayers.

The head carried the main sprayers
and the start-up sprayers, assuring
reliable start-up of the engine.

The fuel components were fed in
through the start valves directly from
the tanks independently of the main mass
of fuel which flowed through the cooling

cavities of the combustion chamber and
nozzle.

/200

The mixture sprayed by the start-up sprayers was ignited by
means of two ignition devices. In the central portion of the
head there was a glow plug, and at an angle of 20° to the axis
of the motor there were two electrodes on the head, between which
a spark jumped when the engine was started. The engine was
started in two stages. First, the start-up sprayers were used to
create a flame in the head, then the valves were opened and the
fuel components in the cooling cavities of the chamber and nozzle
were fed in under pressure through the main sprayers. The
quantity of fuel delivered and, consequently, the thrust of the
engine depended on the lift of the stop valve needles.

Tests in 1939 and 1940 showed that with a pressure of 19. S
atm, the engine developed a thrust of 280 kg and a specific
impulse of 202 sec. The duration of each test was 9 to 150 sec.
The fuel consumption in the starting mode was 0.12 kg/sec. The
engine never went through flight testing.

168

5.4. Oxygon i.Kl'

Oxygon engines pi anno J at UtKU developed thrusts of onlv up
to "(' kv: with specific impulses of up to l"S sec aiul did not
achieve extended rol tabic opor.it ion; .it RNll, the required thrust
oi oxygon engines was ISO 500 kg. In order to achieve tins, it
was necessary first of .ill to improve the mixture formation eon
ditions, to increase the pressure in the chamber an J to provide
re I Sab It coo line.

The l-k. Pngtncs

rea
jet
coo

ope

'"Cl-
Ot

spa
eng

ope
the

In the first version, the t:k engine used certain sol
li.od in the 02 ami 10 engines oi UIR1>: a prechamhor wi

ut ions
th
spiavers. a pear shaped combustion chamber and external flow
ting with liquid oxygon. In order to increase the reliable
rating time o\ the combustion chamber and uo.-le, shape
amic inserts were used, which also provided the require
the inner contours oi the g;ts path. Ignition was bv el
vk plug, introduced into the chamber through the no;, :1c
ine was tested in March o( I05S. On ring the "th second
ration, it burned through in the area ot the prechamhor
ceramic lining cracked.

d shape
ec t r ie
. The
oi
. and

'01

third Version of UK limiir

lourth Version o( I J* K l-ngine

1(0>

In its second version, the engine had a spherical combustion
chamber of stainless steel, allowing a reduction in the specific
weight of the structure. The upper hemisphere was lined with
ceramic made of roasted aluminum oxide. The lower portion of
the chamber and nozzle were made of steel, and were given
external flow cooling. The lower hemisphere of the chamber
burned through during the 19th second of a test conducted in May
of 1935.

In its third version, the engine did not have external flow /202
cooling, but the entire chamber was lined with an aluminum oxide
ceramic, the nozzle was lined with a magnesium oxide ceramic. The
streams of fuel components were directed against each other, which
achieved good mixing. The engine was tested in March of 1935.
The engine was shut down after 27 seconds for inspection, which
revealed small cracks in the ceramic lining.

The results of testing of all three versions of the engine
were used as a basis for analysis of the reasons why the required
specific impulse and stability of thermal mode of the combustion
chamber had not been achieved, leading to the conclusion that in
order to achieve a thrust of 300 kg with a pressure in the chamber
of about 12-16 atm, the chamber volume would have to be about
2 i. Furthermore, since all of the cooling and heat protection
systems tested had failed to assure extended reliable operation,
these engine versions were acknowledged to be suitable only for
brief experimental operation 1 .

In the fourth version, considering that the region of the
critical cross section had failed in earlier tests, the nozzle
was an all-metal copper part with external flow cooling but with-
out ribs. In order to avoid the thermal stresses which frequently
caused failure of the structure, the nozzle was cooled by the
alcohol fuel rather than by the liquid oxygen oxidizer. However,
during flame tests the nozzle failed after 30 seconds. This was
a result of the insufficient cooling intensity, a result of the
low velocity of movement of the alcohol through the cooling
channels.

The 205 Engines

Based on ballistic planning of wingless rockets and the
results of flame testing of the 12K engines, it was considered
necessary to assure constant thrust, decrease the amplitude of
fluctuations of chamber pressure, reduce the time required to
reach the nominal mode to 2.5 sec, increase the specific impulse
to at least 215 sec at a chamber pressure of 20 atm, provide

i Later, the 12K engine was tested on the Aviavnito rocket.

170

continual oprritt ion of si least 2$ see aiul reJuce engine weight

to n.'- we? •• i ,

. :in

First 'it of

the ; . ine

Third Vets toil of
i i '" 1 r • 'St-

ill accordance with these requirements, the ~0S engine was
made in several versions. In the first* all-metal version, jet*

type sprayers for the fuel and one central spiral sprayer for the
Milliter were used, sine? these sprayers had heen fully tested on
the ORN engines. The chamber and noirle were cooled by alcohol;
in order to lighten the structure, dura I i urn was Msetl for a number

of parts* Chambers were tested vith »ez^le< »»*Je of Ultra I um in and
of co i - , The ; erimeni • ;h -■-•■ ' ■■ - ith a ,•- »f co ig
alcohol of about ft m/see, an engine with a Jural turn tioszle failed

after 10 see, with « copper fiossle - - after 20 see.

A number of special studies were performed at RNU to deter- ;

mine the causes of failure of the ne:;,tes and find reliable
methods of cooling. It was established in particular that the
west vulnerable point was the region of transition of the spiral
path to a smooth path. Although the cooling een4itio«.s were
improve*! in the seecniJ vevsien, the ne".;le still failed during
flame testing during the first few seconds of operation of the

eflgllte-,- :::::

/■:i)4

171

In the third version, the velocity of the alcohol in the

cooling space was increased to -0 $»/see. in late December I93n,
a number of tests were performed with nickel -plated copper
nozzles. During the first test, fearing a failure, the experi-
menter shut down the engine aTfer SS seconds* during which time
the engine had developed a thrust of 94 kg and a specific impulse
of 20t» sec with si chamber pressure of 12.5-15 at*. Inspection
of tin « tg -,'*• even led si iht lasutgc '• econd t sf was less
successful «- the engine 'ailed after 27 seconds. The duration of
the third test was 44 seconds. During this test » with a chamber
pressure of 13. S atm» the engine developed ;i thrust of 9b kg, and
a specific impulse of ..* •• . sec.

As a result
and a sprayer
Mixing the fue

.it, it *u*s decided to use ribbed cooling channels
head which would provide better atomizat ion and

1 conpofients.

The RDK-1-1S0 Engine

The RDK-1-IS0 Engine f hurtling liquid oxygen and ethyl alcohol,
was intended to test experimentally the possibility and expediency
of using oxygen engines in manned fl ight" vehicles, ' The designers'
of the engine* l, S, Pushkin and V. A. Shtokolov> selected the
G-14 glider for the R0K-1-IS0, since this glider had a greater
load carrying capacity than the M*M$» so that larger fuel tanks
coal 1 '-: -■ ca rriei.

Whereas the IZK and Z0$ engines were based on the C.IRP i*Z and

10 engines, the RdK- 1-150 engine wade extensive use of the

experience gained at RMi! in the development of nitric acid

engines.

The head of the RDK-1-150 Has a hemisphere with 12 centri-
fugal sprayers for oxygen and a sprayer unit with ft alcohol
sprayers. Two aviation spark plugs were placed between the /.JOT
sprayers for ignition purposes, l»ue to the low power of this
ignition source* in Infer models pyrotechnicnl and chewiest
, gn i I i m were used,

The combustion chamber consisted of the inner wall* made of
copper, and the jacket, made of duralumin. The copper chamber

carried ,i quadruple spiral rib pattern on the outside ;o guide

the flow of the liquid oxygen. The nozzle and chamber were

joined by means of flanges. The iwztle consisted of an inner

copper wall, a duralumin jacket and a sleeve, The quadruple

spiral ribbing on the outer surface of the wall formed channels f ZOb

for the cooiing fluid. The fluid used to cool the nozzle was

alcohol* which passed through the nozzle cooiing channels, then

through tubing into the upper cavity of the chamber head and thus

into tbi iprtyers.

it:

The RPK-1-lSO en^itu* passed a
series of flame tests in May of

1938, During the first tests, the
chamber burned through in areas
where the $trea»s ef oxygen struck
the «alls. After several change*
in design of the head* these
fiii lures were stopped. During
1 1 I i Septei b . •;••' ' -.18 » a
chamber pressure of 10 atm was
reached, providing a thrust of 150

kg and a specific
sec

impulse of 200

The RM- 1-150 Engine

In January of i940» the engine
passed flame testing on a model of
the ti-14 glider, operating at the
design thrust level for 72 sec with
a chamber pressure of 11 at». €e»-
pj s *< — ' i ■■• ''.'■ • • 1 50 w 1 RDK-i-
150 engines, designed for atanned
flight' vehicles, as to their opera-
tional, ecetto - i ■ ■ • els irac-
teris U* •• ins c it ;..' i sa I a :h c< .
find its own area of effective
appl „\ it i in.

The KDK- 1 • ! St> Kngine

The f-Ingtne
• • lev

P. I

In i

, I

developed at Mil

Shatilov was

The plan for this engine
included several promising
ideas. Far e x amp U% the f we I
a ! pen »nt ei ' . i '• iter

>C the ci minis m\
chamher through tiny afer-
t tires distributed over the
Mir face of the conduction
chamber. The fuel components
were fed to these apertures
through longitudin.nl channels
ferwed by the wall of the
vomhust to/, chnmher and its
■ ket 111 fuel empof -nts*

mixed next to the chamher walls* upon leaving the tiny channels,.
Th«* fuel component > flowed out of the tiny channels in a tangen-
tial directum, creating a boundary layer n**xt to the wall to

;:® m

r3

improve the protection of the walls from high heat fluxes.

The liquid oxygen evaporated in the cooling system was sent
to the turbine which drove the fuel pump. The oxygen from the
turbine was then sent to the combustion chamber in gaseous form.

Thus, the engine of P. I. Shatilov featured the progressive
ideas of the porous combustion chamber and a feed system which
allowed the spent turbine gas to be burned.

RNII encountered technical and technological difficulties
which were insurmountable at the time in its attempts to construct
the P. I. Shatilov engine. Therefore, work on this engine was
halted in 1936.

3.5. Developments by Design Bureau No. 7 (KB-7)

In August of 1935, a design bureau (KB-7) was set up in the
Main Artillery Administration of the Red Army. This bureau
included some of the workers from the oxygen team of RNII plus
a number of specialists from various general machine building
enterprises. KB-7 was headed by L. K. Korneyev; other workers
included A. I. Folyarnyy, E. P. Sheptitskiy, P. I. Ivanov,
M. G. Vorob'yev, A. S. Rayetskiy and other. KB-7 had a small
production base, two laboratories and a testing station. The
laboratories and testing station were equipped with modern (for
the time) measurement apparatus, since the flame test stand of
KB-7 was considered a very significant installation.

Together with the development of LRE, KB-7 performed flight /209
testing of rockets burning liquid oxygen-ethyl alcohol fuel.
First, the planning of the LRE was based on the experience of
the work with the 02 and 10 engines (GIRD), then on certain
achievements of the teams at RNII.

The first models of engines in the M family were designed
for the R-03 and R-06 rockets. They were most reminescent of the
10 engine of GIRD, with its pear-shaped combustion chamber with
ceramic lining and prechamber with jet-type sprayers. One such
engine, with a design thrust of 100 kg, was installed on the
R-03 rocket.

The length of the rocket was 2.18 m; diameter 0.2 m; launch
weight 30-33.5 kg, including 8 kg of oxygen and 4.5 kg of alcohol;
extractive fuel component feed was used.

The first launch of the rocket was conducted in April of /210
1937. After modification of the rocket (it was now called the
R-03-02), it wus tested with the same engine in flight 6 times.

174

The R-06 rocket, the first version
of which was planned and manufactured
in 1934 by A. I. Polyarn'iy for Osoavia-
khim, utilized an engine with a cylindri-
cal combustion chamber and a peaked head.
The nozzle had a shaped ceramic insert.
After modernization at KB-7, the rocket,
with a launch weight of 39 kg, was tested
a number of times in 1937-1938 with an
engine with a design thrust of 100 kg.

A number of versions of the M-29
engine were planned for the R-05 rocket,
made in 1938-1939 on order from the
Geophysics Institute of the Academy of
Sciences USSR. In this engine, both
fuel components entered the combustion
chamber through spiral sprayers with
ball back valves. The fuel was extracted
from the tanks by means of a powder
pressure accumulator developed by A. B.
Ionov. The combustion chamber was
conical in shape; the head had a ceramic
liner. In other versions, the chamber
was cylindrical. The nozzle of the
engines had external flow cooling by
alcohol flowing through a spiral ribbed
channel space. The M-29s engine, which
passed stand testing, was designed for
the R-05 rocket; the design thrust was
175 kg.

KB-7 developed and tested a combined
engine which was transferred from RNII.
As a result, the M-17 combined engine,
an LRE which carried a charge of solid
fuel in its combustion chamber, was
planned under the leadership of V. S.
Zuyev. The solid fuel burned first,
providing high thrust for several seconds
(first stage). At the end of the burning of the solid charge,
liquid fuel components were fed to the combustion chamber and the
engine went over to its main operating mode (second stage) . The
solid fuel charge of the M-17 engine consisted of two one-channel
caps and was held in place at the nozzle end by an easily burned
oak plug. Black powder igniters were placed at both the nozzle
and head ends. The head of the chamber carried spiral sprayers
with ball back valves. The exit apertures of the sprayers were
plugged with powder on the chamber end. The nozzle was flow
cooled by alcohol, which began moving when the engine shifteU to
the LRE mode. During combustion of the solid fuel, the powder

/208

The R-03 Rocket

/211

175

plugs and the oak plug were fully consumed,
testing in 1938.

The engine passed

Since the volume of the combustion chambers of modern LRE
would allow the placement of a solid fuel charge incomparably
small in comparison to the quantity necessary for first stage
operation, combined engines have not been further developed.

The activity of KB-7 did not
yield the expected results, and it
was disbanded in 1939, its test stand
and equipment transferred to RNII.

One Version of the M-29
Engine

176

The first great step of mankind will
be when he flies beyond his atmos-
phere and orbits the Earth.

K. E. Tsiolkovskiy

Chapter 4. Liquid-Fueled Rocket Engines for Aviation /212

The Great Patriotic War, from its very beginning, required
increases in the speed, altitude and maneuverability of all
types of combat aircraft.

One solution to this problem was the use of rocket engines
as primary or supplementary (accelerator) engines. Therefore,
the suggestion of leading specialists in the area of rocket
technology that LRE be used in this manner was actively
supported.

Engines intended to be the main engines for combat aircraft
must develop rather high thrust -- about 1000 kg, while reaction
accelerators must develop 300 kg or more. These engines were to
provide high specific impulse, long-term (totalling several hours)
reliable operation, multiple restart capability, plus the capa-
bility of being refueled rapidly. Therefore, these engires were
only designed to utilize nonvolatile oxidizers.

Liquid-fueled rocket engines for combat fighters (inter-
ceptors) were developed at 0KB under the leadership of V. P.
Glushko, at RNII under L. S. Dushkin and at the Design Bureau (i
the Peoples Commissariat for the Aviation Industry (NKAP) by a
team headed by A. M. Isayev.

4.1. The Liquid-Fueled Rocket Engines of 0KB NKAP

In 1934-1938, V. P. Glushko continued to develop LRE (0RM-53-
ORM-102) and gas generators (GG-1, GG-2) in the subdivision of /213
RNII which he headed, which had been transferred from GDL and
reinforced with additional engineers and technicians -- F. L.
Yakaytis, S. S. Ravinskiy, D. P. Shitov, V. N. Galkovskiy and
others.

Beginning in 1939, according to a task assigne.1 by the
Peoples Commissariat for the Aviation Industry, the team of
designers headed by V. P. Glushko began to specialize primarily
in the creation of aircraft LRE -- accelerators. By this time,
some experience had been gained with such engines, sine t as
early as 1932 GDL had begun development of experimental LRE for
aircraft. Plans called for the installation of two LRE with
thrust of 300 kg each beneath the wings of an 1-4 aircraft.

177

In 1940-1946, a series of LRE were produced with pump
fuel feed: RD-1, RD-lKhZ, RD-2 and RD-3. Some of these engines
passed flight and state testing and were put in series production.
The planning and development of these engines were preceded by
the development of individual LRE plans and plans for subunits.
For example, in 1940 the Design Bureau developed a plan for a
two-chamber LRE-accelerator with a thrust of 2 x 300 kg for instal-
lation on the S-100 aircraft. This engine was to burn nitric
acid and kerosene. The fuel components were to be fed by a pump
unit driven by one of the main (piston) engines of the aircraft.

During this same year, a single-chamber nitric-acid LRE with
a thrust of 300 kg was planned. The turbine pump unit of this
engine had a single-stage turbine, a speed reducer, oxygen, kero-
sene and oil pumps.

In 1940, development was begun on a four chamber nitric acid-
kerosene LRE with a thrust of 1000-1200 kg with a sivgle turbine
pump unit.

In 1942-1945, this trubine pump unit was constructed, but it
was never fully developed, since by this time the testing of gear
pumps driven by the main (piston) engine was completed.

A two-chamber L-shaped gas generator, the GG-3, delivering
2 kg/sec gas at 450° C and 25 atm pressure, was planned in 1939-
1940 for planned turbines of marine torpedos. The generator
burned nitric acid and kerosene, but water was sprayed into the /214
combustion products in order to reduce the gas temperature. All
three components were supplied to the generator by me'ns of a
supplementary turbine -pump unit.

The nitric acid and kerosene were sprayed into the combustion
chamber of the generator through spiral sprayers. The combustion
chamber of the generator was cooled with water flowing over the
spiral ribs in the space between the chamber wall and jacket,
then was sprayed into the gas stream in the area where the com-
bustion chamber and mixing chamber were connected. The mixing
chamber was also cooled by water flowing through a spiral channel.
The water was then fed through centrifugal sprayers into the
combustion chamber; here it evaporated, additionally reducing
the temperature of the generator gas and cooling the walls of
the mixing chamber.

The supplementary turbine-pump unit, designed to feed the
generator, consisted of a turbine which drove the working
wheels of three rotating blade pumps through a reducing gear unit.
The turbine was to be started by a pyrotechnical starter with a
cap of trotyl pyroxylin powder. The consumption of gas and
vapor for the turbine pump unit amounted to 31 of the delivery
of the gis generator; the power of the supplementary turbine was

178

15 hp at 28,600 rpm; the GG-3 including turbine pump unit weighed
54 kg. The plan was never brought to life.

In 1942, the combustion chamber of an RD-1 engine with pumv
feed operated for 1 hour and 10 minutes without being removed
from the test stand, and servea as a prototype for the combus-
tion chamber for the RD-l-RD-3 engines. In 1947, the design of
the RD-4 engine, supplied by a turbine-pump unit, was developed.

The RD-1 Engine

The single-chamber RD-1 reaction engine was designed as a
supplementary engine --an accelerator for aircraft in order to
briefly improve their flying, speed and tititvde characteristics.

The calculated data of the RD-1 are as follows: fuel -
nitric acid (OST-701-41) and tractor kerosene (OST-6460); maximum
thrust at ground level -- 300 kg; fuel consumption in maximum
thrust mode -- 1.5 kg/sec; pressure in combustion chamber --
22.5 atm; time of continuous operation at maximum thrust --30
min: nump shaft rotating speed -- 2000 rpm; operating time until
first disassembly --45 minutes.

The RD-1 engine consisted of the following units, separately
installed on the aircraft: the engine itself (combustion chamber /215
with starting and control units) , located in the tail portion of
the fuselage or motor gondola or in the wings of the aircraft;
the pump unit, driven by the main engine of the aircraft either
directly or through a transmission shaft; the choke valve unit
and nitric acid and kerosene lines. The choke valve unit was
controlled by the pilot from his instrument panel, which ?lso
carried a display and the testing and control instruments.

The engine mode control system was supplied by the electric
batteries and compressed air cylinders of the aircraft. The
engine could be started as many as five times in one flight
(limited by capacity of starting tank).

The combustion chamber was mounted on the frame of the rocket
engine together with the following units: starting unit, st • fn?
carburetor, acid and kerosene filters, acid and kerosene v,. ;e>
and electromagnetic pneumatic control valve. The combustion
chamber of the engine consisted of the ignition chamber and .he
combustion chamber itself plus its nozzle. The ignition chamber
was split; its forward half was finned and air cooled, it!" rear
half was cooled by kerosene. The combustion chambtv consisted of
the kerosene-cooled head, and the chamber-nozzle c?e!ed by nitric
acid. A liquid flow gap was maintained between tiio ;_•< ket sur-
rounding the head and chamber-no2-le and the split, »naped
inserts. The middle portion cf the chamber head carried the

179

nitric acid arid kerosene sprayers, centrifugal, closed type, with
a hydraulically controlled needle valve. The bodies of the
sprayers were made barically of stellite. The material of the
acid sprayers was stellite, of the kerosene sprayers -- EZh-2.
The fuel which leaked through the sprayer seals was drained
through the blocking valves.

The bolts of the chamber had spring washers to allow tem-
perature expansion of chamber parts without disrupting the seal
of the joints. The exit portion of the nozzle was equipped with
a gland seal which allowed the nozzle to move relative to the

jacket as the temperature changed.

The ether-air starting mixture was fed into the ignition
chamber through the starting valve, the kerosene -- through a
nipple in the throat of the head jacket, the nitric acid --
through a nipple in the throat of the combustion chamber-nozzle

jacket. The combustion chamber carried a glow plug, starting

valve, pressure relay, filling and blocking valves.

/218

/216

-w«a*>

Overall View of RD-1 Rocket Engine

The application of LRE for aviation required that combustion
c iambers be produced with long operating lives. I here fore,
particular attention was i,iven to intensification of cooling.
Combustion c' iraber walls were made using metals with low modulus
of elastici ;-, coefficient of linear expansion, Poisson's coeffi-
cient and h.jj.'t values of heat conductivity and strength at the
operating temperature. This, in combination with low wall thick-
ness and effective f ii. ling on the side wet by the cooling fluid,

180

was designed to increase the life of combustion chamber walls.

Starting Unit

Pressure v Relay

Combustion
Chamber/

/217

Starter
Valve

Pressure
Relay

Sector

Blocking- 4

Relay

Control
Sector

ne manometer
To acid manometer

Switch

Schematic Diagram of RD-1 Engine
Installation

Beginning in 1941, methods were developed for intensifica-
tion of heat exchange by decreasing the thickness of the boundary
layer and elimination of the products of vapor formation and
gasification from this layer.

Turbulization of the boundary layer in the most highly
stressed sections of the chamber -- the area of inflow and the
critical cross section of the nozzle -- was achieved by drilling
a system of apertures in the aluminum liner of the nozzle, allow-
ing components to be withdrawn from areas with elevated pressure.

The pump unit was attached to the plate of the trout flange.
Two stainless steel shafts made in one piece with the gears
delivering the nitric acid were placed in the split aluminum body
of the pump unit. Splines on these shafts carried the driving
gears which delivered the kerosene, and a guaranteed minimum gap
was maintained between the teeth of the acid gears, to prevent
them from contacting and wearing. Each shaft had three middle-
type journal bearings and two ball thrust bearings on one end.
A guaranteed minimum clearance was also provided between the
body and the "nds of tne gears of the oxidizer pump. The seal was
provided by graphitized asbestos glands. The fluid which soaked
through the glands was carried away through internal drilled
apertures to the intake cavity of the pump. The pump unit carried

iei

reducing valves, which also acted as safety valves protecting the
lines from hydraulic shock.

S. P. Korolev worked in the special design bureay headed by
V. P. Glushko (1942-1946 as Deputy Chief Designer for Flight
Testing), as did Deputy Chief Designers G. S. Zhiritskiy and /219
D. D. Sevruk, subunit leaders V. A. Vitka, N. N. Artamonov, A. S.
Nazarov, G. N. List, N. L. Umanskiy, N. S. Shnyakin, A. A.
Meyyerov N. A. Zhelfikhin, M. A. Kolosov and other highly quali-
fied specialists.

In 1944, the special design bureau of A. A. Meyerov developed
nitro oils and lubricants which did not react with the nitric
acid. They were successfully used in the seals and ball bearings
of the RD-1, RD-lKhZ, RD-2 and RD-3 engines.

In order to continue development of the RD-1 under flying
conditions and accumulate operating experience, S. P. Korolev in
1943 developed an installation for this engine for the Pe-2 series-
produced aircraft. The engine was installed in the tail portion
of the fuselage. The pump unit, compensation and drainage tanks
were carried in the left motor gondola behind the forward
longeron. The engine had dual controls, carried in the pilots
cabin and the radio operator-gunner's cabin.

The Pe-2 aircraft conducted 24 flight tests at altitudes of
up to 7000 m to develop the ignition system. After ground flame
tests were conducted, in 1943 this same aircraft performed 18
start-ups of the RD-1 engine on the ground and 11 in flight.
The longest time of continuous operation of the RD-1 engine at
full thrust in flight was 10 minutes, determined by the capacity
of the fuel tanks.

Flight testing was performed by test pilots A. G. Vasil'chenko
and A. S. Pal'chikov, with S. P. Korolev and D. D. Sevruk flying
as experimental engineer.

The tests of the Pe-2 aircraft continued in 1944-1943 in
order to increase the reliability and altitude capability of
the ignition system, with 49 flame tests on the ground and 38 in
flight. Preference was given to the system of repeated chemical
ignition, which was well-developed by that time, rather than the
ether-air ignition system with glow plug and oxygen feed used
earlier.

In 1944-1945, the RD-1 engines passed ground and flight
tests on fighter aircraft designed by S. A. Lavochkin (La-7),
A. S. uirovlev (Yak-3), P. 0. Sukhoy (Su-6) and the aircraft
designs by V. M. Petlyakov (Pe-2).

182

The RD-lKhZ Engine

/220

An improved version of the
RD-1 engine, with chemical
ignition and a number of design
innovations, came to be called
the RD-1 KhZ.

The two internal parts
of the combustion chamber of
the RD-lKhZ -- the chamber-
nozzle, made of EZh-2 stainless
steel, and the head, made of
heat resistant DPS aluminum
alloy -- were connected by means
of steel jackets of EZh-2.
Between the jackets and the
internal parts of the chamber
there was a passage for nitric
acid in the chamber-nozzle and
kerosene in the head. Longi-
tudinal and spiral fins were
made on the outer surfaces of
the chamber-nozzle and head of
the combustion chamber in order
to improve cooling conditions.
Split aluminum sleeves with an
interior profile corresponding
to the profile of the chamber
parts were placed around the
throat of the head and the
nozzle.

The kerosene entered the
jacket of the combustion
chamber head and moved, cooling
the chamber, to its middle por-
tion, toward the belt of
sprayers. The nitric acid was
fed into the jacket around
the chamber-nozzle through a
nipple at the critical cross
section, then flowed first
toward the exit plane of the

nozzle, then through the spaces between fins between the insert

and chamber-nozzle to the sprayers.

/221

One Version of the Combustion
Chamber of the RD-1 Engine

The sprayers were located at the head of the combustion
chamber, inclined to its axis and directed away from the nozzle.
The sprayers were of the same design as those u:vd in the RD-1
engine.

183

The starting sprayer
was located on the axis
of the chamber. The
starting fuel was fed in
through the central por-
tion of this sprayer, with
nitric acid fed in through
the annular space around
this valve.

The starting fuel
used in the RD-lKhZ
engine was product B23-75,
hypergolic in combination
with nitric acid, developed
at OKB in 1945 by A. A.
Meyerov. This product consisted of 75% (by weight) carbonal and
25% type B-70 gasoline. Chemical ignition of the RD-lKhZ engine
was first tested on the stand, then on the PE-2 aircraft.

/223

Overall View of the RD-lKhZ Engine

/222

One Version of the Combustion Chamber of
the RD-lKhZ Engine

The pump unit of the RD-lKhZ engine consisted of two sections:
the nitric acid and kerosene pumps. A gear- type pump was used,
the kerosene gears serving as the driving gears, allowing a
guaranteed minimum clearance between the teeth and gear ends in
acid pump.

184

Let us study the pneumatic-hydraulic system of the RD-lKhZ
engine. The nitric acid and kerosene were fed into the combus-
tion chamber by the pump unit, driven by the main aircraft
engine. This drive was by means of a friction clutch, switched
on by feeding oil under pressure through an electrohydraulic
valve. This valve was opened by the end switch on the engine
control sector.

The acid and kerosene delivery lines were connected through
the choke valve unit to the intake lines of the pump. With the
valves closed, the pump unit developed the maximum feed pressure,
corresponding to the maximum engine thrust. When the valves
were opened, the acid and kerosene pressure dropped, reducing
the thrust. This allowed the thrust of the LRE to be regulated
without changing the operating speed of the main aircraft
engine.

The safety valves of the pump unit opened when the feed
pressure rose above the maximum value and allowed the excess
fluid to return to the pump intake line.

The nitric acid and kerosene from the pump unit were fed
through filters to the fuel valves, which were opened by com-
pressed air passing through an electromagnetic pneumatic valve,
and were closed by springs. As the engine operated, the fuel
components were fed through the valves into the combustion
chamber, the nitric acid cooling the combustion chamber and
nozzle before entering the sprayer, while the kerosene cooled
the head.

The engine
was star.eJ by
simultaneously
feeding nitric acid
and the starting
fuel to the start-
ing sprayer. The
nitric acid was
supplied by the
pump unit, the
fuel -- from its
tank. The starting
fuel ignited spon-
taneously upon
contact with the
nitric acid, form-
ing the ignition
flame. The slight
pressure arising in
the chamber was

used to open the fuel valves and make the switch to the main

operating mode.

/22S

Testing of the RD-l-KhZ Engine

18S

/224

r ,i Kerosene
; ' from Tank

The Pump Unit of the RD-l-KhZ Engine

In order to eliminate hydraulic shocks and explosions in the
chamber during start-up (as in the RD-1 engine), the fuel com-
ponents were continually fed through the cooling cavity, provid-
ing a staged start-up mode. The c uel components i ere drained from
the hydraulic lines in the chambc • when the engine was shut down.
The actual operating life of the RD-lKhZ engine was increased to
several hours.

During the deve-opment of the RD-lKhZ engine, 2200 start-ups
were performed, 228 of these on the Pe-2 aircraft. At the same
time, RD-lKhZ engines were developed for the aircraft of A. R.
Yakovlev (Yak-3), S. A. Lavochkin (La-7R and 120R) and P. 0.
Sukhoy (Su-7). The Yak-3 aircraft underwent plant flight testing
in 1945, showing an increase in speed of 182 km/hr at an alti-
tude of 7800 m. The test with the La-7R aircraft achieved a
maximum speed of 795 km/hr at an altitude of 6300 m. In 1946,
ground tests (58 start-ups) and flight tests (5 start-ups) of
the RD-lKhZ engine were conducted on an La-120R aircraft. On
18 August 1946, on Aviation Day, 120R aircraft No. ASh-83 par-
ticipated in an air parade, flying
its ID-lKhZ engine in operation.

/226

over Tushino airfield with

The RD-1 and RD-1KHZ engines were series produced during
the war. These engines were stand and flight tested, and the
RD-lKhZ underwent state testing in 1946.

186

The RD-2 Engine

In order to
double the thrust of
the RD-1 engine, the
length of the cylin-
drical portion of
the chamber-nozzle
s.nd number of fuel
sprayers were
increased in the
RD-2 engine, and a
number of design

changes were made, reflecting the experience gained in earlier

investigations.

The RD-2 Rocket Engine

The Combustion Chamber of the RD-2 Engine

The RD-2 engine, like the earlier engines in its family,
u ilized a gei ype pumping unit, differing from the pumping
unit of the R .<hZ engine in its increased operating speed.

The pneumatic-hydraulic systems of the RD-2 and RD-lKhZ
engines were similar, but improvements *ere made to the system
of the RD-2, allowing a softer start-up.

The pneumatic-hydraulic and electrical systems of the RD-2
engine, due to improvements in certain individual elements, were
utilized on the RD-lKhZ engine beginning in the second half of
1946.

/227

/228

187

The? RD-2 engine passed state testing in 1947 and had an
operating life of several hours (the life was limited by puiv
gear wear) .

The basic data of the engine are: thrust at ground level
600 kg; fuel consumption 3 kg/sec; time of continuous operation
at nominal thrust 6 min (limited by capacity of fuel tanks);
guaranteed operating life before first disassembly 1 hour; pres-
sure in combustion chamber 21 a tin. Operating speed of pump
unit drive shaft 2500 rpm.

The RD-3 Engine

The gas generator included three chambers: the ignition,
combustion and mixing chambers. The turbine pump unit consisted
of an active single-stage turbine, a reduction gear, oil unit,
acid, kerosene and water pumps. The turbine used friction
bearings; one of these was water cooled. The maximum turbine
shaft speed was 26000 rpm.

This series of engines was completed in the three -chamber
RD-3 liquid fueled rocket engine, which was stand tested in
1944-1945. It was an autonomous engine, since for the first
time the nitric acid and kerosene were supplied by a turbine
pump unit driven by a gas turbine. The working fluid of the
turbine consisted of the combustion products of the fuel of the
LRE (nitric acid and kerosene), produced in a special unit -- a
gas generator. The RD-3 engine installation included three RD-1
combustion chambers, each of which included a ?et of service
devices -- carburetor, gas pressure relay, filters, fuel valve
and electromagnetic-pneumatic control valves, plus filler valves.
The thrust of this engine at ground level was 900 kg, in a
vacuum -- 1000 kg; the RD-3 could be regulated in thrust from
100 to 1000 kg. In the maximum operating mode (take-off, forced
vertical climb and horizontal acceleration), all three chambers
were used, with thrust varying in the range of 300 to 900 kg;
during horizontal flight, taxiing and landing, only 1 chamber was
used, providing a thrust range of 100 to 250 kg. The pressure
in the combustion chamber reached 22.5 atm.

Control of the engine was fully automated, and automatic
blocking was used to prevent improper starting of the engine.
Start-up of the chambers and control of the engine (start-up,
thrust regulation, shut-down) were performed by means of a single
lever, equipped with an end switch and connected to the choke /230
valve unit of tie gas generator. Choking was used to set the
proper value of pressure in the gas generator and the correspond-
ing operating speed of the turbine pump unit and, consequently,
the thrust of the engine. The design of the engine included
remote control of start-up and shut-down.

188

/229

The RD-3 Rocket Engine (Tor View)

The first version of the turbine pump unit used a high-
speed three- stage centrifugal acid pump, the rotor of which
turned in ball bearings. The blade- type kerosene and water pumps
of this version were identical in design. Their rotors were
balanced, the body was profiled. In order to assure normal oper-
ation of the kerosene and water pumps, they were equipped with
safety valves. In the second version of the turbine pump unit,
all pumps were centrifugal. The fuel components entered the
gas generator from the tanks under compressed air pressure.

Thus, between 1940 and 1946, the Design Bureau headed by

V. P. Glushko created a series of RD engines, distinguished by a
number of advantages. Designed for aircraft, they allowed thrust
variation over a broad range and were auit-3 reliable. The engines
could be repeatedly restarted. In spite cf many hundreds of
restarts of an engine without removal from the test stand, the
limit of the operating life was never reached. Therefore, the
instructions for operation of these engines stated that the
number of permissible restarts, within tie total operating time
of the engine, was not limited. These engines first used grouping
of several chambers, which was later widely developed in domestic
rocket engine construction, and utilized turbine pump units and
gas generators. Finally, the processes of start-up, control and
shut-down of the engines were fully automated. The road leading
to the development of e 'gine^ with this degree of sophistication
was not an easy one. During development of the electrical and

189

pneumatic-hydraulic systems of these engines, repeated accidents /231
occu-red, fortunately resulting only in material damage.

The RD-4 Engine

Exterior View of RD-3 Engine Turbine
Pump Unit

In 1946, the
design of the autono-
mous RD-4 engine, with
1000 kg thrust, was
developed. The turbine
pump unit of this
engine was driven by
the products of decom-
position of hydrogen
peroxide, and the
reducing gear was dis-
tinguished by its low
weight and small size,
thanks to the use of
high-speed centrifugal
pumps for all fuel
components. However,
this design was not
further developed,
since OKB then special-
ized in the development
of powerful LRE.

4.2. The Liquid-Fueled Engines of RNII and the NKAP
Design Bureau

At RNII, a team of designers headed by L. S. Dushkin devel
oped the D-l-A-1100 liquid- fueled ro-ket engine, intended for
use on an interceptor designed by V. F. Bolkhovitinov, A. I.
Bereznyak and A. I. Isayev.

The data of the engine are: nominal thrust 1100 kg; pressure
in chamber 19 atra; specific impulse 204 sec; fuel -- nitric acid
and kerosene; ignition -- glow plug; weight 48 kg.

Due to the difficulty of adjustment of the fuel component
feed system pump, A. I. Isayev, on the suggestion of V. F.
Bolkhovitinov, developed an extractive feed system for the
D-l-A-1100 engine n cooperation with M. V. Mel'nikov. The use
of this feed system required a redesign of the aircraft.

The first flight of an interceptor with the D-1-A-M00 engine
was held on 15 May 1942, by test pilot G. Ya. Bakhchr uidzhi.

/233

190

/ W *J &K

The D-l-A-1100 Liquid-Fueled
Rocket Engine

Installation of the D-l-A-1100 Engine
in an Aircraft

After 1943, the D-l-A-1100 engine was modernized by A. M.
Isayev, This version of the engine retained the basic dimensions
of the chamber and nozzle of the D-l-A-1100 engine, in which the
nozzle had spiral fins with constant spacing, the fins oeing

perpendicular to the axis of the nozzle, so that at the exit plane
they approach the wall at an angle of 30°. The new nozzle had

sextuple fins of variable spacing and variable slant. This

allowed a decrease in wall thickness w A th a simultaneous increase /254

in rigidity of the structure.

Like the D-l-A-1100, the nozzle was
cooled with kerosene, the cylindrical
portion of the chamber with the ox.dizer,
passing through a multiple spiral channel
system. The head, as before, was spheri-
cal in shape; the spiral sprayers were
located in a circle, at the center of
the head was the starting unit, the
sprayers of which were equipped with
ball valves to prevent leakage and fur-
ther combustion of the components when
the engine was shut down. The engine
operated in three modes; starting raode,
then, depending on the position of the
control lever, developing a thrust of
400 or 1100 kg.

State stand testing of the engine
was conducted in October of 1944, after
which it was installed on an aircraft,

which performed the planned program of flight testing successfully
with the LRE in operation.

/235

Aleksey Mikhaylovich
Isayev

Exterior View of the RD-2M-3 Two-Chamber Engine

192

Aleksey Mikhaylovich Isayev ;i 308-1971) , was born 24 October
1908 in Peterburg. After graduating from the Moscow Mining
Institute in 1932, he first worked in construction, then in
planning organizations, and beginning in 1934 at enterprises of
the aviation industry, Together with ¥. F. Bolkhovitinov and
A. Ya. Bereznyak, A. M. Isayev participated in the creation of the
first Soviet aircraft with LRE, in which test pilot C>. Ya.
Bakhchivandzhi flew on IS May 1942.

In 1944, he headed one of the design organizations involved
in rocket engine building and was among the creators of many

engines for rockets and spacecraft. Engines developed under the /23 i
leadership of Aleksey Mikhaylov *ch were carried on the Vostok-
Vaskhod and Soyuz manned spacec.aft and on automate interplane-
tary stations.

A, M. Isayev was a member of the CPSU, a Hero cf Socialist
Labor, a Lenin and State Prise Laureate, a Pocket of Technical
Sciences. A. M. Isayev was awarded four Orders of Lenrn, the
Order of the October Revolution and many meda* of the USSR.

The team of L. S. Dushkin. of which we spoke earlier, devel-
oped aircraft LRE with turbine pump units, designed as main
engines for the aircraft and eliminating the need to use a
propeller motor installation. This family of engines included
the RD-2M, RD-2M-3, RD-2M-3V, RD-KS-1 and others.

The RD-2M engine burned nitric acid and kerosene, its gas
generator operated on hydrogen peroxide. Its maximum thrust was
1400 kg, minimum *hrust 350 kg; the duration of concinuous oper-
ation was 40-60 sec. After 40-45 operational cycles, th» com-
bustion chamber was replaced with a new one. The operating lTfe
of the turbine pump unit (with two-stage turb? »e) and the \apor-
gas generator was 1.5 hr. *■ •ie combustion chamber carried single-
component spiral-type sprayers; ignition was by an electric spark
plug.

The RD-2M-3 engine was developed in 1944. In contrast tc
the RD-2M. it had an additional combustion chamber, developing a
maximum thrust of 300 kj; and a minimum thrust of 100 kg.

The next engine, the RD-2M-3V, with a thrust of 2000 kg, wa r
designed tor an experimental aircraft; its development was begun
in 1944. In 1947-19^8, the engine underwent further testing.
As in earlier models, the fuel was supplied by a turbine-pump
unit; the unit had thre, centrifugal pumps: for nitric acid,
kerosene and 801 hydrogen peroxi ' . A >oiid catalyst was used to
break down the hydrogen peroxide.

The RDKS-1 regulated engine, designed for multiple starts,
utilised liquid oxygen and ethyl alcohol. The thr >~t of the

engine in the maximum mode was 1500 ,cg, in the minimum mode --

193

300 kg; the specific impulse at the nominal mode was 205-210 sec.
The cooling was combined -- external flow cooling in combination
with internal film cooling. The component for creation of the
film entered the cylindrical portion of the chamber and the
expanding portion of the nozzle. The sprayer head-precharaber /237
was made in the form of a cone expanding toward the chamber. The
oxygen sprayers were jet type; the alcohol sprayers were centri-
fugal. The walls of the chamber and the nozzle were spirally
ribbed. Fuel was supplied by a turbine-pump unit. The
turbine gas was produced by decomposition of 801 hydrogen peroxide
ir a gas iterator. Testing of the RDKS-1 was completed in 1947.

L. S. Dushkin began developing the RDD-203 rocket and its
Klw-600 combined engine in 1939.

The KRD-600 had two stages of thrust - 500!) and 1100 kg.
During operation in the first stage - with 5000 kg thrust -- the
fuel used was powder, which filled the combustion chamber; the
pressure in the chamber was 220 atm; the operating time of the

engine was 0.5-0.6 sec, depending on the initial temperature of

the charge.

Operation in the second stage -- with 1100 kg thrust --
utilized liquid fuel -- nitre acid and kerosene; the feed system
was extractive, using a powder- type pressure accumulator; the
pressure in the combustion chamber was 42 atm; specific impulse
220 sec; operating time 9 sec.

The combustion chamber was made of steel and was not cooled,
the nozzle was made of cepper, cooled by a copper heat-accumulat-
ing insert. The sprayers were centrifugal (spiral type) with a
plug which burned out as the engine operated in the first stage.

Test firing of the rocket was conducted in 1939-1940,
initially from a nonmoving support, then from a mechanized 10-
charge launcher.

The basic data on the RDD-203 are: diameter 200 mm; length
3000 mm; launch weight 220 kg; payload 50 kg; design range 23 km.

In addition to these engines, L. S. Dushkin directed the
development of LRE designed for various purposes.

Beginning in 1954, the Design Bureau headed by S. A.
Kosberg worked on LRE for aircraft using one-component fuel, then
after 1956 -- two-component fuel. This office soon developed a
number of medium-thrust LRE designs which were widely used in
rocket and space technology.

194

Tsiolkovskiy pushed back the
boundaries of human knowledge and
his ideas on rocket flight in
space have only today begun to be
realized in their full grandiosity.
S. P. Korolev

Conclusions

After the victorious conclusion of the Great Patriotic War, /238
we could only expect to achieve success in the study of space by
means of the use of powerful LRE with high characteristics,
including reliability. Developaent of theoretical problems of
rocket dynamics, the creation of rocket designs with high payload
efficiency a, id the study of systems for stabilization and control
of the flight of these rockets were also required, as well as
the development of the ground equipment for spaceports.

The forces of all workers in the area of space technology
were devoted to the solution of this group of problems in our
country.

As a result, the Soviet Union continued along the path to
space and opened the space era on 4 October 1957 with the launch
of the world's first artificial Earth satellite.

Since 1949, high altitude rockets had been launched system-
atically in the USSR. One ol the first rockets -- the V-2-A --
*>as a geophysical rocket, designed to study the upper layers of
the atmosphere, photograph tae spectrum of the sun, perform
medical and biological investigations, etc.

Tho V-5-V rockets were designed for astrophysical , geophysi-
cal, medical-biological, ionospheric and other studies. On these
rockets, experiments were continued with animals, including their
return to Earth.

The engines of the V-2-A and V-5-V rockets, designed by
the GDL special design bureau, were single-chamber engines, burn-
ing liquid oxygen and alcohol fuel. The fuel was carried in load-
bearing tanks (the walls of tr • tank formed the skin of the
rocket) by a turbine pump unit driven by the products of decomposi-
tion of hydrogen peroxide. Operation of the V-2-A and V-5-V and
similar models allowed the designers to go on to the creation of /240
morepowerful, improved models, making basic changes in the design
of the engines.

On 12 April 1961, the world's first manned space flight

occurred. A multistage rocket designed by Academician S. P.

195

Korolev carried the Vostok spacecraft and pilot-cosmonaut Yuriv
Alekseyevich Gagarin into orbit.

/239

The V-5-V and V-2-A Geophysical Rockets

The three- stage Vostok booster rocket consists of four
side units (first stage) located around the central unit (second
stage). Above the central unit is the third stage of the rocket,
bach of the first stage units carried a type RD-107 four-chamber
LRH, while the second stage carried a four-chamber type RD-108
engine. These engines, created by GDL-OKB, have been in use
since 195? and are still used.

Burning liquid oxygen and kerosene, the RD-107 engine devel-
ops a thrust of 102 t in a vacuum with a specific impulse of
314 sec, while the RD-108 develops 9b t with a specific impulse
of 315 sec.

196

The main combustion, chambers of each engine, like the
guidance chambers, are supplied by a common turbine pump unit;
the RD-107 includes two, the RD-108 -- four guidance chambers.

*»3£3¥-"

»ff5f*S*-'

/241

0\*€ • aJ ; : ' :■■ ••: >J •* '■-" <■•-, to] \ H ket

The use of several chambers in a single engine allows the
length of the engine and the weight of the rocket to be reduced.
Furt! • 10 it is a e: a : hi • •/• a stable - ml >t Lo! process
in a < h imh< ■ o I >maJ I : r ■■: 3 /me.

The turbine pump unit (TPU) consists of a gas turbine, two
centrifugal pumps supplying the main fuel components, and two
supplementary pumps, driven through an rpm multiplier and designed
to feed liquid nitrogen and hydrogen peroxide to the TPIL The
liqu. I itrogen, •• - I to blow j *'■• • inks, is e- ipoi ited in a
tubu • bine.

11. : combu it ion

cylindrical soldered
The fire wall of the
those areas most hea
fire wall is finned,
outer supporting jac
a va /•'/ furm c - , 1
is s • , .• ;ti •'-• I* -he j

he o
system of the struct

hi ml :•, the ID- 107 ■ RD LOi engines is a
-welded structure with a flat sprayer head.

chamber is made of heat-resistant bronze in
vily thermally loaded. The outside of the

The ips ' the - • ar< "... ;cted to :he
ket by a hi^h- temperature solder applied in
n less thermally loaded parts, the fire wall
acket by the same solder by weans of a corru-
uter, cold wall is a part of the load-bearing
ure» allowing strong, light combustion

/244

19?

chambers to be made. The head of the chamber carries two-com-
ponent bronze sprayers, assurin.: good mixing of th? components
and, consequently, complete combustion. The combustion chamber
has not only external flow cooling by the fuel, but also internal
film cooling.

General View of RD-107 Engine

The third stage /242
carries a single-chamber
LRE with four guidance
nozzles.

During the powered
section of the flight,
the engines of the cen-
tral and side units at
first operate simultan-
eously. After the fuel
of the side units is
exhausted, their engines
are shut down and the
side units are separated
from the central unit,
the rocket engine of
which continues to oper-
ate at full thrust.
After the fuel of the
central unit is exhausted,
the third stage engine
starts up and the third
stage is separated from
the central unit. The
third stage is shut down
and the spacecraft
separated from the boos-
ter by a control system
when the design velocity,
corresponding to injec-
tion of the spacecraft
into the desired orbit,
is reached.

Another Soviet booster rocket which has been widely and
successfully used for many years for comprehensive study and
the performance of practical tasks in near-Earth orbit is the
Kosmos rocket. The two-stages of this rocket are located one
above the other.

The first stage of the Kosmos
engine, which develops a thrust of
fie impulse of 264 sec.

rocket utilizes an RD-214

74 t in a vacuum with a speci

/248

198

the

greatest thrust and
specific impulse ,v< all
know engj ->- s a ■ this
type, bu ".> . tig n t r i :.
acid- •• : ?carl >i lei.
The c ',-; i e i s a
chamber engine, with a
coramc n turb e mrop
unit. ' I e o •; us 1 on
chamber had external
flow cot ] ing - Fun her-
more, t i< p< i , ph ;-ral
sprayers f.-<- m ; p >-
tective fu I Layer along
the walls • he '• ' - -'t ing
fuel , h) pe - goi i> in com-
binat ii wi h h - ' jsic
oxid 3 i into
the fuel J ? the
pump. The engine has
thrus t an i ; . € ! ir.sump-
tion regul tc • a] tow-
ing f 1 -• ■. ib ' Lty :>i its
Flight progi am. The
RD-2! ; . .; ine was
design d i : e >2
and 195? a--, i ha • been
flying since 1957. It
is or i :he t .. *ly
devel'. - me > ' s . . ' : :
GD ,-OK)

/243

General View of the RD-108 Engine The second stage of

the (Cosmos rocket
carries an RD-119 engine, developed by GDL-OKB between 1958 and
1962. The engine burns liquid oxygen and unbalanced dimethyl
hydrazine, developing a thrust of 11 t in a vacuum. It has the
highest specific impulse of all oxygen engines using non-volatile
fuel. The specific impulse of the RD-119 engine in a vacuum is
- ' se . , if t he comb is ton :i ai be i I ^ , c.

Theengine has a high altitude nozzle profile. The gas for /249
the TPU is produced in a single-component gas generator, utilizing
the basic fuel. The design of the engine makes wide use if the
latest structural materials, basically titanium. The steering
system of the engine is designed for control and orientation of
the second stage of the rocket in flight, Control is achieved by
redistribution of the spent turbine gas among the guidance nozzles.
Ignition is pyrotechnical. Preliminary spinning of the turbine
pump unit is by a powder charge in the gas generator.

139

m

jssss^

Since 1965, the Soviet Union has performed
deep studies of high and super-high energy cosmic
rays utilizing apparatus carried on the heavy proton
space stations.

The engines of the proton booster rocket are
made according to a new, highly perfected design.
The power of the proton engine installation is
three times that of the Vostok booster rocket.
The high pressure in the combustion chamber, the
high quality achieved in the processes of mixture
formation and combustion and the care given to
development of the processes of exhaust of the com-
bustion products from the nozzle and design of the
feed system have allowed these powerful engines to
be made quite small with exceptionally high
characteristics.

Low power rocket engines are used in manned

spacecraft and unmanned spacecraft for various
purposes. One such engine is the correction device
designed by A. M. Isayev, used to correct the orbits
of the Molniya-1 communications satellites and the
flight trajectories of automatic interplanetary
space probes such as the Zand spacecraft.

This engine operates on liquid fuel, develop-
ing a thrust of 200 kg in a vacuum for a period of

65 sec.

The ac

technology
"Space" Pav
at the Exhi
Economy in
History of
in Kaluga,
number of o

hievements of Soviet rocket and space
have been extensively exhibited in the
ilion of the Academy of Sciences USSR
bition of Achievements of the National
Moscow, in the State Museum of the
Astronautics imeni K. E. Tsiolkovskiy
in the GDL Museum in Leningrad and a
ther museums and exhibitions.

Almost a half century ago, K. E. Tsiolkovsky
wrote: "Man will not always remain on the Earth,
Th v n c mr ,<. but in his pursuit of light and space will first
Booster penetrate timidly beyond the limits of the atmo-
Rocket sphere, then master all of solar space."

This prediction of the great genius is today
being confirmed as a scientifically well-founded prediction.

The first steps on the path to space were made using a
liquid- fueled rocket engine. The solution of many current prob-
lems requires improved LRE as well as engines of basically new
design.

/252

200

The use of fluorine fuel can increase the specific impulse
of LRE to approximately 500 sec; the specific impulse of a
nuclear rocket engine with a solid-phase reactor can reach 1000
sec, with a gas-phase reactor -- 2500 sec.

The use of thermonuclear energy can be expected to be rtill
more effective.

The electric rocket engines now being designed will signifi-
cantly expand our capabilities in the area of space technology.

It is possible that in time the achievements of science and
technology will show us means and methods of penetrating outward
into the universe so effective that progress in the area of
astronautics will exceed our most optimistic dreams.

/24S

The Vostok Booster Rocket

201

/246

General View of the RD-214 Engine

202

/247

General View of the RD-119 Engine

20 3

General View of the RD-119 Engine
Turbine Pump Unit

/2 50

The Cc i -• . .v \ ■.. • Ses igned
by A, M. I save v

204

/2S1

The "Space" Pavilion of the Academy cf Sciences
USSR at the Exhibition of Achievements
of the National Flccnomy

:05

STANDARD TITLE PAGE

I. R

SRsf'

TT F-15,408

J. G»*«rnm*Ai Aec«iniof» No.

4. T..u^v, k ,»,. MVELOPMMTOF RUSSIAN
ROCKET ENGINE TECHNOLOGY

Ye. K. Moshkin

«. Performing 0r,«,i««tjen Htm* m4 AtfeVa**

Leo Kanner Associates
P. 0. Box 5187
,M$M99d City California.

-24036-

NASA, Code KSS-1
Washington, D. C. 20546

3. fUcipieni't Ca»al»g Ne.

5. Report Dele

May 1974

6. P»tf*t»»ng Ofg«nii«»iOft C«<f«

8. Performing Orgenitetisn Reftert No.

10. Wo* Unit No.

i I . Cen«o«» o> Orent Ho.

13. Tyee »i «•?«,! on.-s Peried

Translation

IS. %upt>l*<**nfatf No«e»

Translation of Razvitiye Otechestvennogo

Raketnogo Dvigatelestroyeniya , Moscow, Mashinostroyeniye
., 1973, 256 pp.

>ress,

t*. *»»»««» This book covers the history of the creation of Soviet
liquid-fueled rocket engines. The works of K. E. Tsiolkov-
skiy and Yu. V. Kondratyuk on the selection of engine and
rocket designs, including multi-stage designs, are described.
The properties of fuel components for liquid-fueled engines
and the work of F. A. Tsander in the area of the creation of
original plans for spacecraft and rocket engines are studied.
The use f elements of the structure of the rockets as
additional fuel is analyzed; the creative path of a number
of leading Soviet scientists is traced. The activity of the
first Soviet rocket organizations is discussed and the
liquid- fueled rocket engines and aircraft rocket engines
which they created are analyzed. Some modern, high-power
liquid-fueled rocket engines are described.

!?, K#?Wer«j,

©y Aytf!<*f{0|

?«. 0if«rlW«t«»

Unclassified. Unlimited.

»«. Security Clattil. (el thii ra^wt)

None

». Security Cl»»§il. {•( «.*« »««•)

None

«. Ne. «f P«

205

22. Price

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