Report on cryogenic engine

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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 common-
ly 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 hydrog-
en, 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.
Cryogenic Engines are rocket engine that uses a cryogenic fuel or oxidizer, that is, its fuel
or oxidizer (or both) are gases liquefied and stored at very low temperatures, otherwise
they would 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.

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The Space Shuttle's main engines used for lift-off 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.
1.1 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.
After World War II the American government and military finally seriously considered
liquid-propellant rockets as weapons and began to fund work on them. The Soviet Union
did likewise.
The RL10 was the first liquid hydrogen rocket engine to be built in the United States, and
development of the engine by Marshall Space Flight Center and Pratt & Whitney began in
the 1950s, with the first flight occurring in 1961. The RL10 was first tested on the ground
in 1959, at Pratt & Whitney's Florida Research and Development Center in West Palm
Beach, Florida. It was first flown in 1962 in an unsuccessful suborbital test, the first

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successful flight took place on November 27, 1963. For that launch, two RL10A-3
engines powered the Centaur upper stage of an Atlas launch vehicle.
Figure 1:An RL10 engine at U.S Space and Rocket centre
The engine produces a specific impulse (Isp) of 373 to 470 s (3.66–4.61 km/s) in a
vacuum and has a mass ranging from 131 to 317 kg (289–699 lb) (depending on version).
Six RL10A-3 engines were used in the S-IV second stage of the Saturn I rocket, one or
two RL10 engines are used in the Centaur upper stages of Atlas and Titan rockets and one
RL10B-2 is used in the upper stage of Delta IV rockets.

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CHAPTER-2
PRINCIPLE
Figure-2:Rocket Thrust Equation
In a rocket engine, stored fuel and stored oxidizer are ignited in a combustion chamber.
The combustion produces great amounts of exhaust gas at high temperature and pressure.
The hot exhaust is passed through a nozzle which accelerates the flow. Thrust is produced
according to Newton's third law of motion.
The amount of thrust produced by the rocket depends on the mass flow rate through the
engine, the exit velocity of the exhaust, and the pressure at the nozzle exit. All of these
variables depend on the design of the nozzle. The smallest cross-sectional area of the
nozzle is called the throat of the nozzle. The hot exhaust flow is choked at the throat,
which means that the Mach number is equal to 1.0 in the throat and the mass flow rate m
dot is determined by the throat area. The area ratio from the throat to the exit Ae sets
the exit velocity Ve and the exit pressure pe. You can explore the design and operation of
a rocket nozzle with our interactive thrust simulator program which runs on your browser.
The exit pressure is only equal to free stream pressure at some design condition. We
must, therefore, use the longer version of the generalized thrust equation to describe the

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thrust of the system. If the free stream pressure is given by p0, the thrust F equation
becomes:
F = m dot * Ve + (pe - p0) * Ae
Notice that there is no free stream mass times free stream velocity term in the thrust
equation because no external air is brought on board. Since the oxidizer is carried on
board the rocket, rockets can generate thrust in a vacuum where there is no other source
of oxygen. That's why a rocket will work in space, where there is no surrounding air, and
a gas turbine or propeller will not work. Turbine engines and propellers rely on the
atmosphere to provide air as the working fluid for propulsion and oxygen in the air as
oxidizer for combustion. The thrust equation shown above works for both liquid
rocket and solid rocket engines.

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CHAPTER-3
CONSTRUCTION
Figure-3:Parts of Cryogenic Engine
Major component of cryogenic engine are-
 Combustion chamber (thrust chamber)
 Pyrotechnic ignite
 Fuel injector
 Fuel cryopumps
 Oxidizer cryopumps
 Gas turbine
 Cryo valves
 Regulators
 Fuel tanks, and
 Rocket engine nozzle
Among them, the combustion chamber & the rocket engine nozzle are the main
components of the cryogenic engine.

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CHAPTER-4
WORKING
4.1 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.
Figure-4:Sectional View of Cryogenic Engine

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4.2 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 oxygen
hydrocarbon 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

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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.
Figure-5:Different Type of Fuel Injecter
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.

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4.3 PHASES OF COMBUSTION IN THRUST CHAMBER
4.3.1 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.
4.3.2 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,

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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 pre-combustion
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.

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4.3.3 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
4.3.4 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.

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 The flame velocity more or less depends on the pre-mixedness of the shear layer
only.
Figure-6: Flame Ignition
4.3.5 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.
Figure-7:Flame Propagation

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4.3.6 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.
Figure-8 Flame Lift Off
4.3.7 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.

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 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.
Figure-9:Flame Anchoring

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4.4 ADVANTAGES OF CRYOGENIC ENGINE
1.High energy per unit mass
Propellants like oxygen and hydrogen in liquid form give very high amounts of energy per
unit mass due to which the amount of fuel to be carried aboard the rockets decreases.
2.Clean fuel
Hydrogen and oxygen are extremely clean fuels. When they combine, they give out only
water. This water is thrown out of the nozzle in form of very hot vapor. Thus the rocket is
nothing but a high burning steam engine.
3.Economical
Use of oxygen and hydrogen as fuels is very economical, as liquid oxygen costs less than
gasoline.
4.5 DISADVANTAGE OF CRYOGENIC ENGINE
1)Fuel tanks tend to be bulky and require heavy insulation for storage and supply line
to store and transport the propellant. Their high fuel efficiency, however, outweighs
this disadvantage.
2)Propellant may leak as propellant are stored and transported at high pressure to
combustion chamber.

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CHAPTER-5
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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CHAPTER-6
REFERENCES
[1] “Rocket propulsion elements” by G. P. Sutton, 7th edition.
[2] “Advances in propulsion” by K. Ramamurthy.
[3] “Rocket and Spacecraft Propulsion” by M. J. Turner.
[4] “Ignition of cryogenic H2/LOX sprays” by O. Gurliat, V. Schmidt, O.J. Haidn,
M.Oschwald.
[5] National Aeronautics and Space Administration, United States Of America
[6] Vikram Sarabhai Space Centre, Thiruvananthapuram
[7] “An Experimental Study of Cryogenic Engine” from International Journal of
Innovative Research in Science, Engineering and Technology.