Cyogenic rocket engine report

CONTENTS

1. Introduction

2. history

3. Space propulsion system

4. Classification of space propulsion system

5. Rocket engine power cycle

6. Combustion in thrust chamber

7. Fuel injection

8. Phase of combustion in thrust chamber

9. Different type of cryogenic engine

10. Conclusion

11. reference

DEPT MECHANICAL ,SEACET 1

CRYOGENIC ROCKET ENGINE

CHAPTER-1

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.

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.

CHAPTER-2HISTORY

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.

CHAPTER-3

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.

CHAPTER-4

Classification of Space Propulsion System



CHAPTER-5ROCKET ENGINE POWER CYCLE

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.



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.



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.



CHAPTER-6

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.



Chapter 7



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.



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

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.

CHAPTER-8DEPT MECHANICAL ,SEACET 13


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.



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 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

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.

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.



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.



CHAPTER-9

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.



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



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.



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.

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

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.



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

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

CHAPTER-10

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.



CHAPTER-11

REFERENCES

“Rocket propulsion elements” by G. P. Sutton, 7th 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



Cyogenic rocket engine report

Engineering

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