Engines
for first space rocket
B.I.Katorgin, V.F.Rakhmanin
Launch of the first artificial
satellite to the Earth orbit which took place on October 4, 1957 became a historical
event and resulted from natural scientific-technical progress, firstly in the
area of engines capable of overcoming the Earth gravity.
Man from ancient times has turned his eyes towards
the Moon and stars striving for space flights. It was reflected in tales,
legends, and science fiction books. All of these works described the space
transport that corresponded to the level of science and technology of those
times. These were powerful vortices, goose caravans, sailing vessels, cannon
balls and many others.
This continued till K.E.Tsiolkovskiy's
paper "Study of outer space by means of jet apparatus" was published in 1903.
(His 150th anniversary was celebrated the last September). In this paper
the author scientifically substantiated that the only means for transportation
in space was a rocket. Unlike powder monofuel used
earlier K.E.Tsiolkovsky proposed to fill the tanks with liquid chemical
substances of higher fuel heat capacity which allowed overcoming the Earth
gravity and moving in outer space.
Significance of
K.E.Tsiolkovsky's services performed by him for the world science and society
is not only in explanation of scientifically proved technical principles of
space flights but also in the fact that his works attracted a group of talented
enthusiasts seeking for possibilities of efficient application of their
creative forces.
Practical works on liquid-propellant rocket
engine (LPRE) creation in our country began in Leningrad Gas-Dynamic Laboratory
(GDL) in 1930th headed by V.P.Glushko. Similar works were conducted
simultaneously in Moscow group for jet propulsion research
(GJPR).
Design of the first LPRE in GDL showed the
necessity of scientific research for successful development of engines for jet
aircrafts. V.P.Glushko started his own research at the beginning of 1930th
accounting for the fact that there were only preliminary theoretical works of
K.E.Tsiolkovsky in the area of liquid-propellant jet engine. During 1930-1933 studying model chambers he found
the optimal profile of supersonic nozzle, chose effective design of combustion
chamber cooling circuit and reliable ceramic heat protective coating made of
zirconium dioxide, used centrifugal injectors that provided higher mixing
quality, tested several variants of fuel ignition at the start including
starting fuel, studied efficiency of different chemical substances as the
components of rocket propellant and so on.
In 1934-38 V.P.Glushko continued his studies at
Rocket Research Institute simultaneously with development of LPRE for winged
rocket and the rocket glider designed by S.P.Korolev.
In the first half of the 1940th a group
of designers supervised by V.P.Glushko used their knowledge and experience for
development of LPRE for propeller airplanes boosters. This stage of engine
development was characterized by LPRE design with pump fuel feed. Endurance of
those engines at multiple starts equaled 40 minutes. LPREs
"RD-1", "RD-1H3" were manufactured in
small series and mounted on experimental fighters "La-7", "Yak-3", "Su-7".
The period from July 1945 till November 1946
V.P.Glushko spent being on business trip in Germany where he studied design and
production technology of German long-range rocket "A-4". The rocket operated using liquid
oxygen and 75% aqueous solution of ethyl alcohol as a fuel. In July 1946 he was appointed Chief Designer
of Experimental Design Bureau at the plant in Khimki
and charged with the works on reproduction of "A-4" rocket engine using domestic
materials. In May 1948 such engine intended for the first Soviet rocket "R-1" successfully passed its first test
at the new firing test bed. Subsequent modernization of the engine base design
allowed increasing the flight range of "R-2" rocket up to 600 km and then the flight range of "R-5" rocket - up to 1200 km.
Studying the design of "A-4" rocket engine the Soviet engineers did
not see any essentially unknown for them design decisions. Moreover, some engine
elements of their own design (in 1930-1940) were even more progressive. They
were impressed only with overall dimensions of the engine and its thrust level
of 25 tons. The most powerful domestic engine at that time had thrust of 1.5
tons. German engineers who created "A-4" rocket engine removed psychological
barrier thus accelerating the progress of our rocket engineering.
After development of "R-5" rocket the issue of military
intercontinental rocket design arose. It could not be solved by further modernization
of "A-4" rocket as engineering solutions
built into it exhausted themselves and a new approach was required. At the same
time positive results of reproduction and modernization of "A-4" rocket and its engine showed
scientific and technical maturity of Soviet specialists and their ability to
create new engineering devices.
During development of intercontinental rocket the
liquid oxygen and kerosene having higher fuel heat capacity were chosen. It
required solving the problem of combustion chamber cooling as at transition
from alcohol to kerosene the combustion temperature increased up to 1000°C and cooling possibilities of
kerosene were inferior to alcohol cooling capacity both in thermophysical
properties and in relative quantity of coolant. A new design was required to
provide combustion chamber cooling: internal wall was made of copper alloy with
high heat conductivity; the ribs were made on the cooled wall surface to enhance
heat transfer and strengthen the structure. The ribs were brazed to the outer
steel wall which was subject to pressure load in combustion chamber. This
design concept and its technological implementation made it possible to create
liquid-propellant rocket engines almost of any thrust in the range of its
engineering expediency and energetic linkage of engine parameters. In 1948 the
first combustion chamber of such design was produced. It was designated as "KS-50" or "Lilliput"
and had an internal diameter of 60 mm and thrust of 100 kg.
In 1950 the combustion chamber "ED-140" with thrust of 7 tons was designed.
This chamber had cooling passage similar to "KS-50", a slot ring of internal cooling
with tangential swirl of the cooling liquid and a plane injector head with
diameter of 200 mm. Injectors were brazed to inner copper
bottom and middle steel bottom. Outer steel bottom had a shape of spherial calotte.
Experience gained during the fire tests of different
"KS-50" and "ED-140" versions as well as results of
investigations made by other Design Bureaus and Research Institutes under
supervision of A.P.Vanichev, A.M.Isaev,
M.V.Melnikov and other authoritative specialists
allowed proving some fundamental design decisions for determination of
combustion chamber design for intercontinental rocket.
Project of two-stage rocket provided cluster
configuration of engine package. Such configuration was chosen according to
several reasons including the absence of experience in the LPRE ignition in
space. The first stage consisted of 4 side engines mounted on the central engine
of the second stage. All engines had one chamber with thrust of 60 tons at the
pressure of 60 atmospheres in combustion chamber. The engine had a turbo pump
feed and the turbine was driven by products of catalytical
decomposition of hydrogen peroxide. Thrust vector control was provided by gas-jet
rudders on the first stage and by either gas-jet rudders or deflecting
(swinging) chambers on the second stage. The project assumed that design
development and autonomous tests of these control units had to be done by the
rocket designer.
The first firing tests of experimental chambers
with thrust of 60 tons and cylindrical section diameter of 600 mm showed that combustion process
tended to instability which could not be eliminated by known methods. Stability
could be increased by decreasing combustion chamber diameter but that required
transition to multi-chamber engine.
New four-chamber engine was developed which
possessed improved thrust performance, and its height was 1.5 m less due to short chamber. Accordingly
the rocket itself was also shorter. Rudder chambers (engines) were used instead
of gas-jet rudders due to increased temperature and velocity in exhaust gas
jet. S.P.Korolev's DB remained responsible for
development and autonomous tests of rudder chambers. They already had an
experimental chamber with the thrust of 3 tons which successfully operated at
the test bed. Later the rudder chamber was redeveloped in V.P.Glushko's DB by
analogy with the main combustion chamber. This allowed reducing its mass and increasing
specific propulsive burn by ~15 sec. Such chambers were put into operation in
1959. Y.A. Gagarin's rocket was equipped with rudder chambers made according to
V.P.Glushko's DB.
Introduction of four-chamber cruise engines and
rudder chambers required providing simultaneous ignition in 32 chambers at the
rocket start: 20 main chambers and 12 rudder chambers. In order to work this
process out a special test bed was constructed in Khimki.
About 1000 tests of fuel ignition and transition to preliminary stage were
conducted at this bed.
During test bed development of "RD-107" and "RD-108" engines two zones of unstable
operation were revealed: "lower" zone (at 40-70% of nominal pressure in
combustion chamber) and "upper" zone (where pressure value exceeded the nominal
pressure by 5-7%). "Lower" zone was dangerous at the engine start. Stable
operation at the start was provided by accelerated pressure increase by means
of specially designed start oxidizer
valve.
Dangerous proximity of the "upper" unstable zone
to nominal engine regime could not be eliminated, however it was determined
that the lower limit of the "upper" unstable zone moved up with increase of
fuel rate through the injectors which formed a "screen" (unfortunately, this resulted
in significant specific propulsive burn decrease). In order to avoid this
danger a selective technology was chosen to be applied when producing injectors
and assembling the mixing head. Each engine was subjected to
control-technological test for rejection of chambers disposed to unstable
operation.
"R-7" rocket engines were the first ones
equipped with the systems of thrust control and fuel components ratio control. The
thrust control system was designated for providing the rocket movement
according to required trajectory at the active path to increase firing accuracy
or accuracy of putting the payload into orbit. The system of fuel components
ratio control provided simultaneous depletion of fuel components from tanks of
all engines and considerably reduced mass of guaranteed residuals which
resulted in rocket flight range increase. Newly designed systems for engine
operation regime control provided requirements specified to them. Their design
became a basic one for development of such units for next engines.
Final engine adjustments resulted in various
and sometimes unexpected problems which were nevertheless successfully solved.
Successful firing tests of packages of five engines which took place in
February and March, 1957 became the final stage of ground development.
Further development of "R-7" rocket continued during flight
development tests. The first "R-7" launch took place on May
15, 1957. It
finished in 98 seconds because of the fire in one of the side tail modules due
to fuel line leakage. The second launch was cancelled due to flaws in operation
of oxidizer line control units. During the third launch the rocket lost its
stability in 33 seconds due to control system failure and collapsed at the altitude
of ~ 4.5 km. The forth launch on August
21, 1957
was successful and it was announced to the whole world by the "Telegraph Agency
of the Soviet
Union". On
October 4, 1957 "R-7" rocket put into orbit the first
artificial Earth satellite. Due to high reliability of all the systems since
April 1961 the "R-7" rocket in its space version started
delivering manned spacecrafts to outer space.
High reliability of engines was supported during
the years of their flights. The critical remarks were of production type and introduced
changes were only cosmetic.
After more than 30 years of engine operation
the designers dared to introduce significant changes concerning the mixing
heads in the chambers of main engines. New mixing head was designed on basis of
"R-9" rocket engine chamber. Similar to
"RD-170" engine combustion chambers it had antipulsating injector baffles. Spiral ribbing of cooling
passage was implemented at the most heat-beat areas of the chamber. These measures
allowed raising the lower limit of the upper unstable zone, increasing specific
propulsive burn by ~ 5 sec and the payload mass by 200 kg. "Soyuz-FG" rockets with such
engines have been operating since 2002.
Over ten versions of space rockets based on "R-7" rocket design operated for 50
years. Over 1725 launches have been fulfilled; over 8600 engines have been
utilized during the flights; over 10000 engines were produced including engines
for test bed development. This liquid-propellant engine is the only one in the
whole world possessing such a statistics.
