Cryogenicrocketenginesreport 120907091108-phpapp01

Semester VII

Branch: Mechanical Engineering

Seminar Title: Cryogenic Rocket
Engines

INTRODUCTION
Cryogenics originated from two Greek words “kyros” which means cold or
freezing and “genes” which means born or produced. Cryogenics is the study of very
low temperatures or the production of the same. Liquefied gases like liquid nitrogen and
liquid oxygen are used in many cryogenic applications. Liquid nitrogen is the most
commonly used element in cryogenics and is legally purchasable around the world.
Liquid helium is also commonly used and allows for the lowest temperatures to be
reached. These gases can be stored on large tanks called Dewar tanks, named after
James Dewar, who first liquefied hydrogen, or in giant tanks used for commercial
applications.
The field of cryogenics advanced when during world war two, when metals were
frozen to low temperatures showed more wear resistance. In 1966, a company was
formed, called Cyro-Tech, which experimented with the possibility of using cryogenic
tempering instead of Heat Treating, for increasing the life of metal tools. The theory was
based on the existing theory of heat treating, which was lowering the temperatures to
room temperatures from high temperatures and supposing that further descent would
allow more strength for further strength increase. Unfortunately for the newly-born
industry the results were unstable as the components sometimes experienced thermal
shock when cooled too fast. Luckily with the use of applied research and the with the
arrival of the modern computer this field has improved significantly, creating more stable
results.
Another use of cryogenics is cryogenic fuels. Cryogenic fuels, mainly oxygen and
nitrogen have been used as rocket fuels. The Indian Space Research Organization
(ISRO) is set to flight-test the indigenously developed cryogenic engine by early 2006,
after the engine passed a 1000 second endurance test in 2003. It will form the final
stage of the GSLV for putting it into orbit 36,000 km from earth.
Cryogenic Engines are rocket motors designed for liquid fuels that have to be
held at very low "cryogenic" temperatures to be liquid - they would otherwise be gas at
normal temperatures.

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Toc H Institute Of Science And Technology, Arakkunnam, Ernakulam

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The engine components are also cooled so the fuel doesn't boil to a gas in the
lines that feed the engine. The thrust comes from the rapid expansion from liquid to gas
with the gas emerging from the motor at very high speed. The energy needed to heat
the fuels comes from burning them, once they are gasses. Cryogenic engines are the
highest performing rocket motors. One disadvantage is that the fuel tanks tend to be
bulky and require heavy insulation to store the propellant. Their high fuel efficiency,
however, outweighs this disadvantage.
The Space Shuttle's main engines used for liftoff are cryogenic engines. The
Shuttle's smaller thrusters for orbital maneuvering use non-cryogenic hypergolic fuels,
which are compact and are stored at warm temperatures. Currently, only the United
States, Russia, China, France, Japan and India have mastered cryogenic rocket
technology.
All the current Rockets run on Liquid-propellant rockets. The first operational
cryogenic rocket engine was the 1961 NASA design the RL-10 LOX LH2 rocket engine,
which was used in the Saturn 1 rocket employed in the early stages of the Apollo moon
landing program.
The major components of a cryogenic rocket engine are:


the thrust chamber or combustion chamber



pyrotechnic igniter



fuel injector



fuel turbo-pumps



gas turbine



cryo valves



Regulators



The fuel tanks



rocket engine



nozzle
Among them, the combustion chamber & the nozzle are the main components of

the rocket engine.

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HISTORY
The only known claim to liquid propellant rocket engine experiments in the
nineteenth century was made by a Peruvian scientist named Pedro Paulet. However, he
did not immediately publish his work. In 1927 he wrote a letter to a newspaper in Lima,
claiming he had experimented with a liquid rocket engine while he was a student in
Paris three decades earlier.
Historians of early rocketry experiments, among them Max Valier and Willy Ley,
have given differing amounts of credence to Paulet's report. Paulet described laboratory
tests of liquid rocket engines, but did not claim to have flown a liquid rocket.
The first flight of a vehicle powered by a liquid-rocket took place on March 16,
1926 at Auburn, Massachusetts, when American professor Robert H. Goddard
launched a rocket which used liquid oxygen and gasoline as propellants. The rocket,
which was dubbed "Nell", rose just 41 feet during a 2.5-second flight that ended in a
cabbage field, but it was an important demonstration that liquid rockets were possible.

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SPACE PROPULSION SYSTEM
Spacecraft propulsion is any method used to accelerate spacecraft and artificial
satellites. There are many different methods. Each method has drawbacks and
advantages, and spacecraft propulsion is an active area of research. However, most
spacecraft today are propelled by forcing a gas from the back/rear of the vehicle at very
high speed through a supersonic de Laval nozzle. This sort of engine is called a rocket
engine.
All current spacecraft use chemical rockets (bipropellant or solid-fuel) for launch,
though some have used air-breathing engines on their first stage. Most satellites have
simple reliable chemical thrusters. Soviet bloc satellites have used electric propulsion
for decades, and newer Western geo-orbiting spacecraft are starting to use them for
north-south station keeping. Interplanetary vehicles mostly use chemical rockets as
well, although a few have used ion thrusters to great success.

Classification of Space Propulsion System

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ROCKET ENGINE POWER CYCLES
Gas pressure feed system
A simple pressurized feed system is shown schematically below. It consists of a
high-pressure gas tank, a gas starting valve, a pressure regulator, propellant tanks,
propellant valves, and feed lines. Additional components, such as filling and draining
provisions, check valves, filters, flexible elastic bladders for separating the liquid from
the pressurizing gas, and pressure sensors or gauges, are also often incorporated. After
all tanks are filled, the high-pressure gas valve is remotely actuated and admits gas
through the pressure regulator at a constant pressure to the propellant tanks. The check
valves prevent mixing of the oxidizer with the fuel when the unit is not in an right
position. The propellants are fed to the thrust chamber by opening valves. When the
propellants are completely consumed, the pressurizing gas can also scavenge and
clean lines and valves of much of the liquid propellant residue. The variations in this
system, such as the combination of several valves into one or the elimination and
addition of certain components, depend to a large extent on the application. If a unit is
to be used over and over, such as space-maneuver rocket, it will include several
additional features such as, possibly, a thrust-regulating device and a tank level gauge.

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Gas-Generator Cycle
The gas-generator cycle taps off a small amount of fuel and oxidizer from the
main flow to feed a burner called a gas generator. The hot gas from this generator
passes through a turbine to generate power for the pumps that send propellants to the
combustion chamber. The hot gas is then either dumped overboard or sent into the
main nozzle downstream. Increasing the flow of propellants into the gas generator
increases the speed of the turbine, which increases the flow of propellants into the main
combustion chamber (and hence, the amount of thrust produced). The gas generator
must burn propellants at a less-than-optimal mixture ratio to keep the temperature low
for the turbine blades. Thus, the cycle is appropriate for moderate power requirements
but not high-power systems, which would have to divert a large portion of the main flow
to the less efficient gas-generator flow.

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Staged Combustion Cycle
In a staged combustion cycle, the propellants are burned in stages. Like the gasgenerator cycle, this cycle also has a burner, called a preburner, to generate gas for a
turbine. The pre-burner taps off and burn a small amount of one propellant and a large
amount of the other, producing an oxidizer-rich or fuel-rich hot gas mixture that is mostly
unburned vaporized propellant. This hot gas is then passed through the turbine, injected
into the main chamber, and burned again with the remaining propellants. The
advantage over the gas-generator cycle is that all of the propellants are burned at the
optimal mixture ratio in the main chamber and no flow is dumped overboard. The staged
combustion cycle is often used for high-power applications. The higher the chamber
pressure, the smaller and lighter the engine can be to produce the same thrust.
Development cost for this cycle is higher because the high pressures complicate the
development process.

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COMBUSTION IN THRUST CHAMBER
The thrust chamber is the key subassembly of a rocket engine. Here the liquid
propellants are metered, injected, atomized, vaporized, mixed, and burned to form hot
reaction gas products, which in turn are accelerated and ejected at high velocity. A
rocket thrust chamber assembly has an injector, a combustion chamber, a supersonic
nozzle, and mounting provisions. All have to withstand the extreme heat of combustion
and the various forces, including the transmission of the thrust force to the vehicle.
There also is an ignition system if non-spontaneously ignitable propellants are used.

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FUEL INJECTION
The functions of the injector are similar to those of a carburetor of an internal
combustion engine. The injector has to introduce and meter the flow of liquid propellants
to the combustion chamber, cause the liquids to be broken up into small droplets (a
process called atomization), and distribute and mix the propellants in such a manner
that a correctly proportioned mixture of fuel and oxidizer will result, with uniform
propellant mass flow and composition over the chamber cross section. This has been
accomplished with different types of injector designs and elements.
The injection hole pattern on the face of the injector is closely related to the
internal manifolds or feed passages within the injector. These provide for the distribution
of the propellant from the injector inlet to all the injection holes. A large complex
manifold volume allows low passage velocities and good distribution of flow over the
cross section of the chamber. A small manifold volume allows for a lighter weight
injector and reduces the amount of "dribble" flow after the main valves are shut. The
higher passage velocities cause a more uneven flow through different identical injection
holes and thus a poorer distribution and wider local gas composition variation.
Dribbling results in afterburning, which is an inefficient irregular combustion that
gives a little "cutoff" thrust after valve closing. For applications with very accurate
terminal vehicle velocity requirements, the cutoff impulse has to be very small and
reproducible and often valves are built into the injector to minimize passage volume.
Impinging-stream-type, multiple-hole injectors are commonly used with oxygenhydrocarbon and storable propellants. For unlike doublet patterns the propellants are
injected through a number of separate small holes in such a manner that the fuel and
oxidizer streams impinge upon each other. Impingement forms thin liquid fans and aids
atomization of the liquids into droplets, also aiding distribution. The two liquid streams
then form a fan which breaks up into droplets.

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Unlike doublets work best when the hole size (more exactly, the volume flow) of
the fuel is about equal to that of the oxidizer and the ignition delay is long enough to
allow the formation of fans. For uneven volume flow the triplet pattern seems to be more
effective.
The non-impinging or shower head injector employs non-impinging streams of
propellant usually emerging normal to the face of the injector. It relies on turbulence and
diffusion to achieve mixing. The German World War II V-2 rocket used this type of
injector. This type is now not used, because it requires a large chamber volume for
good combustion.
Sheet or spray-type injectors give cylindrical, conical, or other types of spray
sheets; these sprays generally intersect and thereby promote mixing and atomization.
By varying the width of the sheet (through an axially moveable sleeve) it is possible to
throttle the propellant flow over a wide range without excessive reduction in injector

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pressure drop. This type of variable area concentric tube injector was used on the
descent engine of the Lunar Excursion Module and throttled over a 10:1 range of flow
with only a very small change in mixture ratio.
The coaxial hollow post injector has been used for liquid oxygen and gaseous
hydrogen injectors by most domestic and foreign rocket designers. It works well when
the liquid hydrogen has absorbed heat from cooling jackets and has been gasified. This
gasified hydrogen flows at high speed (typically 330 m/sec or 1000 ft/sec); the liquid
oxygen flows far more slowly (usually at less than 33 m/sec or 100 ft/sec) and the
differential velocity causes a shear action, which helps to break up the oxygen stream
into small droplets. The injector has a multiplicity of these coaxial posts on its face.

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PHASES OF COMBUSTION IN THRUST CHAMBER
Rapid Combustion Zone
In this zone intensive and rapid chemical reactions occur at increasingly higher
temperature; any remaining liquid droplets are vaporized by convective heating and gas
pockets of fuel-rich and fuel-lean gases are mixed. The mixing is aided by local
turbulence and diffusion of the gas species. The further breakdown of the propellant
chemicals into intermediate fractions and smaller, simpler chemicals and the oxidation
of fuel fractions occur rapidly in this zone. The rate of heat release increases greatly
and this causes the specific volume of the gas mixture to increase and the local axial
velocity to increase by a factor of 100 or more.
The rapid expansion of the heated gases also forces a series of local transverse
gas flows from hot high-burning-rate sites to colder low-burning-rate sites. The liquid
droplets that may still persist in the upstream portion of this zone do not follow the gas
flow quickly and are difficult to move in a transverse direction. Therefore, zones of fuelrich or oxidizer-rich gases will persist according to the orifice spray pattern in the
upstream injection zone. The gas composition and mixture ratio across the chamber
section become more uniform as the gases move through this zone, but the mixture
never becomes truly uniform.
As the reaction product gases are accelerated, they become hotter (due to
further heat releases) and the lateral velocities become relatively small compared to the
increasing axial velocities. The combustion process is not a steady flow process. Some
people believe that the combustion is locally so intense that it approches localized
explosions that create a series of shock waves. When observing any one specific
location in the chamber, one finds that there are rapid fluctuations in pressure,
temperature, density, mixture ratio, and radiation emissions with time.

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Injection/Atomization Zone
Two different liquids are injected with storable propellants and with liquid
oxygen/hydrocarbon combinations. They are injected through orifices at velocities
typically between 7 and 60 m/sec or about 20 to 200 ft/sec. The injector design has a
profound influence on the combustion behavior and some seemingly minor design
changes can have a major effect on instability. The pattern, sizes, number, distribution,
and types of orifices influence the combustion behavior, as do the pressure drop,
manifold geometry, or surface roughness in the injection orifice walls.
The individual jets, streams, or sheets break up into droplets by impingement of
one jet with another (or with a surface), by the inherent instabilities of liquid sprays, or
by the interaction with gases at a different velocity and temperature. In this first zone the
liquids are atomized into a large number of small droplets. Heat is transferred to the
droplets by radiation from the very hot rapid combustion zone and by convection from
moderately hot gases in the first zone. The droplets evaporate and create local regions
rich either in fuel vapor or oxidizer vapor.
This first zone is heterogeneous; it contains liquids and vaporized propellant as
well as some burning hot gases. With the liquid being located at discrete sites, there are
large gradients in all directions with respect to fuel and oxidizer mass fluxes, mixture
ratio, size and dispersion of droplets, or properties of the gaseous medium. Chemical
reactions occur in this zone, but the rate of heat generation is relatively low, in part
because the liquids and the gases are still relatively cold and in part because
vaporization near the droplets causes fuel-rich and fuel-lean regions which do not burn
as quickly. Some hot gases from the combustion zone are re-circulated back from the
rapid combustion zone, and they can create local gas velocities that flow across the
injector face.
The hot gases, which can flow in unsteady vortexes or turbulence patterns, are
essential to the initial evaporation of the liquids. The injection, atomization and
vaporization processes are different if one of the propellants is a gas. For example, this
occurs in liquid oxygen with gaseous hydrogen propellant in thrust chambers or precombustion chambers, where liquid hydrogen has absorbed heat from cooling jackets

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and has been gasified. Hydrogen gas has no droplets and does not evaporate. The gas
usually has a much higher injection velocity (above 120 m/sec) than the liquid
propellant.
This cause shear forces to be imposed on the liquid jets, with more rapid droplet
formation and gasification. The preferred injector design for gaseous hydrogen and
liquid oxygen is different from the individual jet streams used with storable propellants.

Stream Tube Combustion Zone
In this zone oxidation reactions continue, but at a lower rate, and some additional
heat is released. However, chemical reactions continue because the mixture tends to be
driven toward an equilibrium composition. Since axial velocities are high (200 to 600
m/sec) the transverse convective flow velocities become relatively small. Streamlines
are formed and there is relatively little turbulent mixing across streamline boundaries.
Locally the flow velocity and the pressure fluctuate somewhat. The residence time in
this zone is very short compared to the residence time in the other two zones. The
streamline type, inviscid flow, and the chemical reactions toward achieving chemical
equilibrium persist not only throughout the remainder of the combustion chamber, but
are also extended into the nozzle. Actually, the major processes do not take place
strictly sequentially, but several seem to occur simultaneously in several parts of the
chamber. The flame front is not a simple plane surface across the combustion chamber
There is turbulence in the gas flow in all parts of the combustion chamber. The
residence time of the propellant material in the combustion chamber is very short,
usually less than 10 milliseconds. Combustion in a liquid rocket engine is very dynamic,
with the volumetric heat release being approximately 370 MJ/m3-sec, which is much
higher than in turbojets. Further, the higher temperature in a rocket causes chemical
reaction rates to be several times faster (increasing exponentially with temperature)
than in turbojet.
The four phases of combustion in the thrust chamber are
1. Primary Ignition

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2. Flame Propagation
3. Flame Lift off
4. Flame Anchoring

Primary Ignition
 begins at the time of deposition of the energy into the shear layer and ends when
the flame front has reached the outer limit of the shear layer
 starts interaction with the recirculation zone.
 phase typically lasts about half a millisecond
 it is characterised by a slight but distinct downstream movement of the flame .
 The flame velocity more or less depends on the pre-mixedness of the shear layer
only.

Flame Propagation
 This phase corresponds to the time span for the flame reaching the edge of the
shear layer, expands into in the recirculation zone and propagates until it has
consumed all the premixed propellants.
 This period lasts between 0.1 and 2 ms.
 It is characterised by an upstream movement of the upstream flame front until it
reaches a minimum distance from the injector face plate.
 It is accompanied by a strong rise of the flame intensity and by a peak in the
combustion chamber pressure.

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 The duration of this phase as well as the pressure and emission behaviour during
this phase depend strongly on the global characteristics of the stationary cold
flow before ignition.

Flame Lift Off
 phase starts when the upstream flame front begins to move downstream away
from the injector because all premixed propellants in the recirculation zone have
been consumed until it reaches a maximum distance.
 This period lasts between 1 and 5 ms.
 The emission of the flame is less intense showing that the chemical activity has
decreased.
 The position where the movement of the upstream flame front comes to an end,
the characteristic times of convection and flame propagation are balanced.

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Flame Anchoring.
 This period lasts from 20 ms to more than 50 ms, depending on the injection
condition.
 It begins when the flame starts to move a second time upstream to injector face
plate and ends when the flame has reached stationary conditions.
 During this phase the flame propagates upstream only in the shear layer .
 Same as flame lift-off phase the vaporisation is enhanced by the hot products
which are entrained into the shear layer through the recirculation zone.
 The flame is stabilised at a position where an equilibrium exists between the local
velocity of the flame front and the convective flow velocity.
 This local flame velocity is depending on the upstream LOX-evaporation rates,
i.e., the available gaseous O2, mixing of O2 and H2, hot products and radicals in
the shear layer.
 At the end of this phase, combustion chamber pressure and emission intensity
are constant.

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DIFFERENT TYPES OF CRYOGENIC ENGINES
HM-7B Rocket Engine
HM-7 cryogenic propellant rocket engine has been used as an upper stage
engine on all versions of the Ariane launcher. The more powerful HM-7B version was
used on Ariane's 2, 3 and 4 and is also used on the ESC-A cryogenic upper stage of
Ariane 5. Important principles used in the HM-7 combustion chamber were adopted by
NASA under license and it is this technology that formed the basis of today's US space
shuttle main engines - the first reusable rocket engine in the world.
The HM7 engine was built upon the development work of the 40kN HM-4 engine.
In 1973, the Ottobrunn team started development of the HM-7 thrust chamber for the
third stage of Ariane 1. Six years later, the HM-7 engine was successfully qualified with
the first launch of Ariane 1 in December 1979.

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With the introduction of Ariane 2 and Ariane 3, it became necessary to increase
the performance of the HM-7 engine. This was achieved by raising the combustion
chamber pressure from 30 to 35 bar and extending the nozzle, thereby raising the
specific impulse. The burn time was also increased from 570 to 735 seconds. The
upgraded engine was thus designated HM-7B and was qualified in 1983. When
subsequently used on Ariane 4, the burn time was increased to 780 seconds.
In February 2005, the HM-7B successfully powered the new cryogenic upper
stage of Ariane 5, designate ESC-A (Etage Superior Cryo-technique A). This flight was
a tribute to the performance and flight proven reliability of an engine first developed 30
years ago. With the ESC-A upper stage, the payload performance of Ariane 5 is
increased to 10 tonnes. In order to inherit the proven reliability of the HM-7B engine
from over one hundred Ariane 4 flights, engine changes were kept to a minimum. The
main change being a 20% increase in burn time from 780 seconds to 950 seconds on
Ariane 5 ESC-A.
Use of HM-7B on Ariane 5 is a first step toward increasing the payload
performance of Ariane 5. A second step will be the introduction of the new Vinci
expander cycle engine to an ESC-B cryogenic upper stage, increasing the payload
performance to 12 tonnes

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The HM7B engine is a gas generator liquid oxygen / liquid hydrogen engine that
powers the Ariane 4 third stage. The HM7 engine built upon the development work of
the 40 kN thrust HM4. The HM7 development program began in 1973 as part of
Europe's effort to develop an indigenous launch capability.

Final qualification of the

HM7 engine occurred in 1979 and the engine went on to power the third stage of the
Ariane 1. SEP continued to perfect and upgrade the engine, increasing the specific
impulse by 4 seconds by increasing chamber pressure and lengthening the nozzle. The
new engine, the HM7B, powered the third stage of the Ariane 2,3 and 4. As of June 1st,
1995, SEP had produced 111 HM7B engines, with a cumulated total of 171,700
seconds of operation, including 47,400 in flight.

300 N Cryogenic Engine:
This 300 N cryogenic propellant engine has a vacuum Isp of 415 seconds - the
highest value ever achieved in Europe for an engine of such small size.
Being pressure-fed, the engine assembly is relatively simple and avoids the need
for a turbo-pump. The thrust chamber and throat region of the engine are regenerative
cooled using hydrogen propellant. The nozzle extension is radiation cooled.
The engine incorporates a splash-plate injector having a star shaped
configuration. Ignition and subsequent re-ignition is achieved using Tri-ethyl aluminum
(TEA) - which is hypergolic with the oxygen propellant. The number of re-ignitions is a
function of the volume of Tri-ethyl aluminum accommodated. The engine nominally
provides for 1 ignition and 3 re-ignitions using just 1.5 cc of Tri-ethyl aluminum. The use
of a chemical ignition system enables a very compact design.
The engine needs no pre-cooling prior to ignition. Only the propellant feed lines
to the engine propellant valves need be pre-cooled.

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Engine construction materials are mainly stainless steel, Nimonic 75 (ChromiumNickel Alloy) and copper.

Applications
The 300 N cryogenic engines enable the simplicity of a pressure fed propulsion
system whilst offering the performance of a turbo-pump propulsion system.
Being pressure fed, the engine does not require an additional turbo-pump, with
its associated complexity.

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The 300 N cryogenic engines may be used as a main engine in dedicated stages
for orbital insertion, orbital transfer, orbital, and interplanetary applications, including:
Upper stages
Kick stages
Vernier stages
Transfer stages
The 300 N cryogenic engines may also be used as a thruster, or thruster cluster
with existing cryogenic turbo-pump propulsion systems and stages for such applications
as performance augmentation, upgrades, roll control.

Vulcain Rocket Engine
Vulcain (also known as HM-60) was the first main engine of the Ariane 5
cryogenic first stage (EPC). The development of Vulcain, assured by a European
collaboration, began in 1988 with the Ariane 5 rocket program. It first flew in 1996
powering the ill-fated flight 501 without being the cause of the disaster, and had its first
successful flight in 1997 (flight 502). In 2002 the upgraded Vulcain 2 with 20% more
thrust first flew on flight 517, although a problem with the engine turned the flight into a
failure. The cause was due to flight loads being much higher than expected, as the
inquiry board concluded.
Subsequently, the nozzle has been redesigned, reinforcing the structure and
improving the thermal situation of the tube wall, enhancing hydrogen coolant flow as
well as applying thermal barrier coating to the flame-facing side of the coolant tubes,
reducing heat load. The first successful flight of the (partially redesigned) Vulcain 2
occurred in 2005 on flight 521. The Vulcain engines are gas-generator cycle cryogenic
rocket engines fed with liquid oxygen and liquid hydrogen.
They feature regenerative cooling through a tube wall design, and the Vulcain 2
introduced a particular film cooling for the lower part of the nozzle, where exhaust gas
from the turbine is re-injected in the engine They power the first stage of the Ariane 5

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launcher, the EPC (Étage Principal Cryo technique, main cryogenic stage) and provide
8% of the total lift-off thrust (the rest being provided by the two solid rocket boosters).
The engine operating time is 600 s in both configurations.
The coaxial injector elements cause the LOX and LH2 propellants to be mixed
together. LOX is injected at the centre of the injector, around which the LH2 is injected.
These propellants are mainly atomized and mixed by shear forces generated by the
velocity differences between LOX and LH2. The final acceleration of hot gases, up to
supersonic velocities, is achieved by gas expansion in the nozzle extension, thereby
increasing the thrust.

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Applications:


main engine of the Ariane 5 cryogenic first stage (EPC)

VINCI Rocket Engine:
Vinci is a European Space Agency cryogenic rocket engine currently under
development. It is designed to power the new upper stage of Ariane 5, ESC-B, and will
be the first European re-ignitable cryogenic upper stage engine, raising the launcher's
GTO performances to 12 t. Vinci is an expander cycle rocket engine fed with liquid
hydrogen and liquid oxygen. Its biggest improvement from its predecessor, the HM-7 is
the capability of restarting up to five times. It is also the first European expander cycle

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engine, removing the need for a gas generator to drive the fuel and oxydizer pumps. It
features a carbon ceramic extendable nozzle in order to have a large, 2.15 m diameter
nozzle extension with minimum length: the retracted nozzle part is deployed only after
the upper stage separates from the rest of the rocket; after extension, the engine's
overall length increases from 2.3 m to 4.2 m.

Applications:


upper stage of Ariane 5

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CONCLUSION
The area of Cryogenics in Cryogenic Rocket Engines is a vast one and it cannot
be described in a few words. As the world progress new developments are being made
more and more new developments are being made in the field of Rocket Engineering.
Now a day cryo propelled rocket engines are having a great demand in the field of
space exploration. Due to the high specific impulse obtained during the ignition of fuels
they are of much demand.

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REFERENCES

 “Rocket propulsion elements” by G. P. Sutton, 7 th edition.
 “Advances in propulsion” by K. Ramamurthy.
 “Rocket and Spacecraft Propulsion” by M. J. Turner.
 “Ignition of cryogenic H2/LOX sprays” by O. Gurliat, V. Schmidt, O.J. Haidn, M.
Oschwald.
 National Aeronautics and Space Administration, United States Of America
 Vikram Sarabhai Space Centre, Thiruvananthapuram

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