seminar report

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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.
1.1 HISTORY
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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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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2. CRYOGENIC LIQUIDS
Cryogenic fuels, mainly liquid hydrogen, have been used as rocket fuels. Liquid oxygen
is used as an oxidizer of hydrogen, but oxygen is not, strictly speaking, a fuel. For example,
NASA's workhorse space shuttle uses cryogenic hydrogen fuel as its primary means of getting
into orbit, as did all of the rockets built for the Soviet space program by Sergei Korolev. (This
was a bone of contention between him and rival engine designer ValentinGlushko, who felt
that cryogenic fuels were impractical for large-scale rockets such as the ill-fated N-
1rocketspacecraft.)
Russian aircraft manufacturer Tupolev developed a version of its popular design Tu-154
with a cryogenic fuel system, known as the Tu-155. The plane uses a fuel referred to as
liquefied natural gas or LNG, and made its first flight in 1989.

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Fig 2. Characteristic tempereratures of cryogens
2.1 PROPELLANT:LOX / LIQUID HYDROGEN
The engines burn liquid hydrogen and liquid oxygen from the Space Shuttle external
tank. They are used for propulsion during its ascent, in addition to the two more powerful
solid rocket boosters and partly the Orbital Maneuvering System. Each engine can generate
almost 1.8 meganewtons (MN) or 400,000 lbf of thrust at liftoff. The engines are capable of
generating a specific impulse (Isp) of 453 seconds in a vacuum, or 363 seconds at sea level
(exhaust velocities of 4440 m/s and 3560 m/s respectively). Overall, a space shuttle main
engine weighs approximately 3.2 t (7,000 lb). The engines are removed after every flight and
taken to the Space Shuttle Main Engine Processing Facility (SSMEPF) for inspection and
replacement of any necessary components.

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Liquid propellant rocket engines are mostly widely used rocket engines because of many
advantages that liquid propellants have. The first rockets used solid propellants because of the
simplicity of their construction (just a barrel with gunpowder), but such engines were difficult to
control. Chemistry and physics of combustion were undeveloped, combustion was unpredictable
and it was nearly impossible to control it. Liquid rocket engines (LRE) were very promising:
their thrust could be controlled by dosing propellant flow ratio with valves. Although nowadays
the techniques of solid propellants have enormously advanced, liquid propellants retain their
importance for the rocketry. The working principle of all liquid rocket engines is transformation
of the potential chemical energy of liquid propellants to kinetic energy of the exhaust gases. It is
important to mention that LRE is only a part of the propulsion system; other parts are tanks,
plumbing, hydraulics, framework etc.
3. TYPES OF PROPELLANTS
There are two basic types of LRE propellants: monopropellants and bipropellants.
3.1 MONOPROPELLANTS
Monopropellants are liquids which may be stored in a single tank (and remain stable).
They are decomposed releasing energy in presence of a catalyst. Among such monopropellants
are hydrogen peroxide H2O2 and hydrazine N2H4. Hydrogen peroxide is used, for example, as
propellant for the turbopumps in RD-107/108 engines in the first and the second stages of the
Soyuz launch vehicles. It was also used in the Mercury manned spacecraft, in the Centaur upper
stage (in the ullage and attitude control motors) etc. However, this propellant slowly decomposes
by itself, so it cannot be hold for years and thus cannot be used in spacecraft with long lifetimes.
Hydrazine is widely used in maneuvering thrusters or main engines of spacecraft, also in descent
engines. For example, hydrazine was used in the thrusters of the Voyager spacecraft, in the
descent engines of the Viking and the Phoenix Martian landing probes etc. Hydrazine is stable
and may be hold for years.
The main advantages of monopropellants is that they need only one tank and that the
ignition system is not needed, they react by themselves in presence of catalyst. Due to low

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temperatures in the thrust chamber (combustion chamber) they may work for a long time (for
hours) and be restarted (may give thousands of very short pulses). That makes them ideal for
thrusters. Their thrust may vary from tens of grams to several kilograms. The main disadvantage
is their low specific impulse (Isp), for hydrazine mostly Isp < 250 s.
3.2 BIPROPELLANTS
Bipropellants consist of fuel and oxidizer that should be hold separately and react in the
thrust chamber when ignited.
Bipropellant mixtures divide in practice by their properties into hypergolic and cryogenic
propellants (although the first does not exclude the second). Hypergolic mixtures ignite
spontaneously when brought in contact. Thus, no complex ignition system and starting procedure
is needed (thus, multiple restarts are simply feasible). The process of combustion of hypergolic
propellants is also more stable, so engines are more simple to develop and less likely to be
destroyed at work. Hard starts are less likely with hypergolic propellants (hard start occurs when
the ignition in the thrust chamber takes place at presence of excessive concentration of
propellants, thus instantaneous overpressure is established and it may lead to explosive
destruction of the engine).
The most widely used hypergolic mixture is hydrazine or its variants
(monomethylhydrazine /MMH/, unsymmetrical dimethilhydrazine /UDMH/, aerozine /50%
hydrazine + 50% UDMH/) as fuel and dinitrogen tetroxide (N2O4) (or nitric axid /NO/ in earlier
applications) as oxidizer. The great advantage of this propellant is that the fuel and the oxidizer
both are liquids at the normal conditions, they are not cryogenic, so they may be stored at the
common temperatures. This makes them ideal for military applications (ICBM may stay for
arbitrary time in its silo or stored) as well as for spacecraft with a long lifetime (like
interplanetary probes). This is the reason why they have been used on many ICBMs which later
have been converted into launch vehicles: the Titan, the Strela and the Rokot (ICBM UR-
100/100H), the Dnepr (R-36M). These mixtures have quite high Isp (vacuum values 300 – 320 s
and even more), so they are most common on spacecraft, although need more complex engines
when that based on the hydrazine as monopropellant. The Cassini spacecraft uses hydrazine
monopropellant thrusters for small attitude maneuvers, but burns N2O4/MMH in its main engine.

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Orbital engines of the Space Shuttle use N2O4/MMH, Apollo LM and CSM used N2O4/aerozine,
Luna E8 series (Lunokhod, soil sample missions) used N2O4/UDMH. The highest disadvantage
of this mixture is that it is highly toxic (both fuel and oxidizer), so it should be handled very
carefully. The flame of hypergol is nearly colorless, slightly blue. Plumes of even powerful
engines burning this propellant is nearly invisible.
Cryogenic propellants include mixtures at least one component of which (fuel or
oxidizer or both) need low temperatures to liquefy. One of the most common such mixtures is
LOX/kerosene (some special technological processes are used to produce kerosene for rocket
industry, the resulting fuels have different names; in US RP-1 is a popular type of kerosene fuels,
in USSR sintin was used). LOX/UDMH is also an option. LOX, one of the most widespread
oxidizers, boils at –1830
C, thus being moderately cryogenic: this temperature is higher than the
boiling point of air (– 1940
C), which is produced in industrial quantities. So, air does not liquefy
on cold walls of LOX tanks, and extensive termostating is not needed. For LOX/kerosene
propellants, in vacuum Isp  340 s, and it is very widespread propellant for launch vehicles,
specially for the first stages. It is used on all stages of the Soyuz and the Zenit, on the first stage
of the Atlas V, it was used on the first stages of the Saturn I/IB, the Saturn V and the Energia.
Due to its better energetic characteristics than that of the hypergols, LOX/kerosene is also used
on the Blok DM, the 4th
stage of the Proton (3-stage N2O4/UDMH heavy rocket). Cryogenic
propellants cannot be held for a long time, and the tanks should be insulated. In space, some
amount of the propellant boils out, and if some time passes between multiple burns, the loss may
be significant, so instulation is needed. Plumes of LOX/kerosene engines contain many carbon
particles and thus are bright yellowish.
The most energetic propellant that is used in practice is LOX/LH2, it may have Isp > 450
s (vacuum value). Liquid hydrogen is highly cryogenic fuel, having the boiling point at –2530
C.
So it is difficult to use, since continuous thermostating is unavoidable. Tanks containing LH2
should have extensive thermal protection (which leads to increase of their weight), the fuel
cannot be hold in space for a long time since it boils out. Extremely low temperatures of LH2
lead to changes of physical properties of metals which contact with this fuel, metals saturate with
H2; these factors should be taken into account at building of a LOX/LH2 rocket engine,
plumbing and tanks. Because of very low density of LOX/LH2 (~280 kg/m3
), it needs large

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tanks, which also increase the mass of stages with this propellant. In addition, Isp of such engines
significantly drops in the atmosphere. So, LOX/LH2 is used primarily on the upper stages of the
launch vehicles. The first stage where this propellant was applied was the Centaur upper stage
for the Atlas launch vehicle. Later it appeared on the upper stage S-IV (Saturn I) and S-IVB
(Saturn IB and Saturn V), as well as on the second stage S-II of the Saturn V. The Space Shuttle
has become the first spacecraft where LOX/LH2 is used on the stage working from the sea level.
Later it appeared on the second stage (also working from the sea level) of the Energia. Today it
is a common propellant on upper stages of launch vehicles: Delta, Atlas, Ariane, GSLV (India),
CZ-3 (China). It is also used on the first stage started on the sea level, but is aided by strap-on
solid rocket boosters: Ariane, H-II (Japan). There is only one full-cryogenic launch vehicle
which burns this propellant on all stages, that is the Delta IV. The plumes of LOX/LH2 engines
are nearly invisible
4. ENGINE DESIGNS
To introduce propellants into the thrust chamber, two principle designs of LRE are used:
pressure-fed and pump-fed. The first one is the most simple and reliable, the second one enables
to get higher specific impulse.
4.1 PRESSURE-FED ENGINES
In a pressure-fed LRE, the propellants are forced to the thrust chamber by pressure of gas
which pressurizes the tanks. For pressurization, a separate gas supply is provided. So, there is a
special tank with pressurizing gas onboard (helium is commonly used for this purpose).

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Fig 4.1 Pressure-fed engine
The greatest advantage of pressure-fed engines is simplicity and thus reliability of this
design: contrary to pump-fed engines, no complex turbopumps are needed, no gas generators etc.
Such engines contain much less parts and much less moving parts, so there are much less things
that might fail. The procedure of engine cut-off and restart is also very simple: there is no need to
stop and restart the turbopumps, its enough to close or open the valves, and the propellant flow to
the thrust chamber ceases or recommence. To avoid the pressurizing gas to cool down due to
expansion inside the tanks, it is often warmed up in the heat exchanger.

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The advantages of this solution make pressure-fed engines ideal for applications where
reliability and simplicity are important, as well as capability for multiple restarts. This is the
reason why all engines of the Apollo CSM, as well as all engines of the Apollo LM were
pressure-fed. Shuttle orbital maneuvering and control engines are pressure-fed as well.
Maneuvering and attitude control thrusters of satellites and space probes are mostly pressure-fed
since they are restarted thousands of times.
However, this design have two principle disadvantages (mutually related). Specific
characteristics of rocket engines depend on the pressure in the thrust chamber. But the pressure
in the thrust chamber cannot exceed the pressure in the the tanks (actually, the pressure in the
tanks should be higher). To withstang high pressure, large tanks should have more robust and
heavier construction. The larger is the tank, thicker should be its walls to bear the same pressure.
Thus, pressure-fed systems are generally limited by chamber pressures of ~10 bar (on the Apollo
SM it was 7 bar, on the LM Ascent Stage 8.4 bar). They are rarely applied on first stages due to
large size of the tanks (however, the engine on the 2nd
stage of the Delta II rocket is pressure-
fed). To avoid additional pressure in the tanks, regenerative cooling jacket is often avoided, that
obliges to use ablative and radiative cooling.
4.2 PUMP-FED ENGINES
Pump-fed systems do not have the limitations of the pressure-fed systems. In this design,
the propellant is forced into the thrust chambers with dedicated pumps. The required efficiency
may be provided only with centrifugal pumps, herewith the pump should rotate at tens of
thousands rpm. Only a turbine is capable to ensure such speeds, so the natural solution is a
turbopump. A turbopump consists of one or more pumps often mounted on the same shaft with a
driving turbine. The turbine is driven by gas flow, the gas may be produced in a gas generator by
preburning some amount of the propellant, by burning a separate propellant (like hydrogen
peroxide in the RD-107/108 engines on the Soyuz) or by gasification of some propellant in the
cooling jacket of the thrust chamber and the nozzle. The pumps may be multistage. The
turbopump assembly may include also booster pumps, which are added to unload principle
pumps and to increase the pressure in the gas chamber (these pumps may be driven by a
hydraulic turbine powered by liquid from a high-pressure line, but also by the main turbine).
Turbopump assembly is the most complex part of the engine, since it should have enormous

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productivity and work in harsh conditions (the turbine is driven by very hot gases and rotates
very quickly). For example, the turbopump of RD-170/171 (the most powerful LRE ever
produced, Energia & Zenit launch vehicles, LOX/kerosene) has a mass flow rate of ~2.4 tons/s,
it provides the pressure in the thrust chamber of ~250 bar, the power of the turbine is ~200 MW,
it rotates at ~14 000 rpm. The pressure of gases driving the turbine is ~500 bar, their temperature
is ~5000
C. At the same time the turbopump should be compact and lightweight (the mass of
whole RD-170 is about 10 tons). So high characteristics are possible only because the lifetime of
such assemblies is only tens or hundreds of seconds. However, there exist pump-fed engines of
multiple use which may work for hours, be restarted and continue to work after revision. An
example of such engines is the Space Shuttle Main Engine (SSME).
The obvious advantage of pump-fed engines is that they may provide very high pressures
inside the thrust chamber and so their specific impulse is high. In spite of their complexity, they
may be compact enough and be lighter than pressure-fed engines with their pressurizing gas
vessels and thick propellant tanks; thanks to their efficiency, they make it possible to spend less
propellant. Their complexity is the highest disadvantage, since complex turbopump assemblies
tend to be more expensive and less reliable than pressure-fed designs. However, if efficiency is
critical, pump-fed design is a natural solution. Turbopumps are used on all stages of launch
vehicles, but also on spacecraft. The Soviet lunar probes E8 (Lunokhods, soil sample missions)
used pump-fed design, and the Soviet lunar module for the manned expeditions as well (since
weight was critical). The space stations Salyut, the Soyuz manned spacecraft have been provided
with pump-fed engines.
4.21 GAS GENERATOR CYCLE
There are several designs of pump-fed engines. The most spread is the gas generator
cycle. In these engines the turbine of the turbopump is powered by gas resulting from burning
some of propellant in the gas generator (also called preburner sometimes) – a special small
combustion chamber. In some designs there may be two gas generators (like the RD-170/171),
sometimes each gas generator provides gas for separate turbines (of fuel and oxidizer). The
mixture in the gas generator is ordinarily very fuel-rich or oxidizer-rich in order to keep the
temperature reasonably low and not to damage the blades of the turbine (actually, only small
amount of the propellant burns, the rest is only gasified). After the turbine, the gas is ejected,

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either through the main nozzle either through a special nozzle. Due to its low temperature, its
contribution to the engine thrust is quite low, so it is nearly “wasted” for the thrust. Several
percent of the propellant are lost. However, sometimes this gas is used in steering nozzles or may
participate in film cooling of the main nozzle (like in the F-1 engine of the Saturn V).
Fig 4.21 Gas generator cycle

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4.22 STAGED COMBUSTION CYCLE
To improve efficiency of the engine, another version of this cycle is used, that is the so-
called staged combustion cycle (or closed cycle). The main difference of this cycle is that the gas
after turbine is not dumped, but is returned to the thrust chamber. So, all propellant and all heat
pass through the thrust chamber and nothing is wasted. The disadvantage of this solution is that
the turbine have to do work against the pressure of the gases which it should press into the thrust
chamber. So the efficiency of the turbine drops, and it needs more power to work. Thus, it works
in worse and more harsh conditions, the plumbing of hot gases ducts is much more complex, as
well as the control. So such engines are generally more complex, more expensive and less
reliable. They are very sensitive to productional quality and to external particles that may
occasionally get into the ducts, turbines and pumps. But the gain of Isp may be so high that this
design makes sense. It first appeared in the USSR, and they have a long tradition of building
engines of the closed cycle. For example, the RD-170/171 applies the closed cycle (contrary to
the F-1), as well as the SSME of the Shuttle.

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Fig 4.22 Staged combustion cycle
In most cases only small amount of the propellant is gasified in the gas generator (and in
the gas generator cycle it cannot be else to avoid excessive loss of propellant). But in some
developments the full amount of the fuel and the oxidizer passes through the turbine (the so-
called full flow staged combustion cycle). It enables to reduce the temperature of the gas and the
rotation velocity of the turbine, since the it is driven by larger mass. The lifetime and reliability
grow. Of course, two separate gas generators and turbines are needed for the fuel and the
oxidizer. However, separate systems for both components are usual for LOX/LH2 engines, since

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the components have very different physical properties (density on the first place), so it is
difficult to provide optimal characteristics for them in a single assembly.
Sometimes it is possible to get rid of the gas generator assembly at all (gas generator is a
small combustion chamber by itself, with its own nozzle ejecting gas into the turbine, so it is a
complex unit). This is the expander cycle design. In this cycle the gas for driving the turbine is
produced from the fuel vaporized in the cooling jacket of the thrust chamber and the nozzle. A
gas generator is sometimes used to start the engine. This cycle may be opened or closed. In the
opened cycle, only a small portion of the fuel is used to drive the turbine and thereafter it is
dumped. In the closed cycle the fuel is redirected into the thrust chamber after leaving the
turbine. Although the close cycle saves fuel, the open cycle enables higher pressure drop on the
turbine which increases its efficiency and enables to raise the pressure in the chamber. This leads
to higher Isp (this is the case of LE-5A/B on the second stage of the Japanese H-II rocket, LE-5
used the gas generator cycle). The famous RL-10 and its modifications on the Centaur upper
stage use the expander cycle. Generally, the expander cycle is mostly applied in LOX/LH2
engine since fuel is ordinarily used for regenerative cooling (oxidizer is too reactive) and LH2
has low boiling point and is very effective as reaction mass.

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Fig 5. Thrust chamber and nozzle
5. THRUST CHAMBER AND NOZZLE
The thrust chamber is the principle component of the rocket engine, the propellant is
injected into it and burns, transforming into hot gases that escape through the nozzle (the bell).
The thrust chamber assembly consists of the following main components: the thrust chamber
body, the nozzle (the chamber narrows to its end, the narrowest part of it is called throat, and
behind the throat it expands again, forming the nozzle; the end part of the nozzle is called
extension), the injector. In the case of regenerative cooling, the body and the nozzle may be
combined with a cooling jacket.
6. COOLING OF THE THRUST CHAMBER AND THE NOZZLE
Since gases in the thrust chamber have very high temperatures, its walls should be
cooled, as well as the walls of the nozzle. Without cooling, the walls cannot withstand such
temperatures for a long time. There are two principles of cooling: passive and active cooling.

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Different methods may be applied simultaneously in different parts of the chamber and the
nozzle.
6.1 PASSIVE COOLING
Passive cooling includes ablative and radiative methods. Ablative cooling means that the
walls are covered with substance called ablation, which have high heat capacity and absorbs heat
by transforming itself chemically and/or physically. The ablation burns slowly and removes heat
with the gases created in this process. This method is limited by timespan and by temperature:
ablative materials cannot withstand very high temperatures and since they are gradually
removed, ablation works only for a limited time. However, the highest advantage of this method
is simplicity and reliability, so it is sometimes applied even on large engines, like RS-68, the
most powerful LOX/LH2 engine in use (Delta IV).
Radiative cooling is the process when the hot wall loses heat by radiation. Being very
simple, this method is limited to thin surfaces with relatively moderate incident heat fluxes. If the
heat flux is very intensive, the equilibrium temperature of such wall with only radiative cooling
is too high and the wall may be damaged. If the wall is thick, its hot side is damaged before the
heat diffuses to the cold side. So, radiative cooling is mostly applied to cool nozzle extensions, as
well as in small maneuvering engines, where the heat of short burns is absorbed by a massive
conductive wall of the chamber (made from cooper alloy, for instance) and is irradiated between
the burns.
6.2 ACTIVE COOLING
Active cooling methods include regenerative cooling and film cooling. Regenerative
cooling means that a flow of cold propellant is organized along the the hot wall and the
propellant carries away excessive heat. Thus, the walls of the thrust chamber and the nozzle have
a cooling jacket with a propellant flow inside (fuel is commonly used). The propellant is directed
into the jacket from the tank before it is injected into the thrust chamber. There are different
ways to build the cooling jacket. The simplest way is two walls, inner and outer, separated by a
folded metal sheet, with propellant flowing along the folds. This design have been preferred in
Russia and now is also applied in US. The liquid may also flow along rectangular channels
machined or formed into a liner fabricated from high-conductivity material (like cooper alloys).

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The example of such design is the SSME. In US the traditional design have been the thrust
chamber and nozzle built from thin rectangular tubes strengthened by outer bracing (tubular
wall). The tubes are directed downwards and upwards the walls of the thrust chamber and the
nozzle. Since the diameter of the nozzle changes along its axis, the form of the tubes also change
(but their cross-section remains constant). The tubes may bifurcate. The hot wall is very thin, so
heat exchange with the liquid is very effective. The example of this design is the F-1.
Efficiency of the regenerative cooling is very high, but there are also some limitations. In
the throat, the diameter and thus the surface of the chamber wall is small, and it is impossible to
pump enough liquid along the limited surface to provide cooling. It is not always possible to
increase the velocity of the flux since it would require raise of pressure that forces the propellant
through the cooling jacket. In this case, film cooling is additionally applied. The main idea of
film cooling is that some fuel in injected through additional injectors into the hotest parts of the
thrust chamber right against the wall. The liquid fuel absorbs heat by boiling and evaporating,
thus a protective cold boundary film is created and protects the wall from contact with the hot
gas. This film is spread along the wall by gas moving along the chamber. A variation of this
method is transpiration: the coolant gets into the chamber from the jacket through a porous
chamber wall. Film cooling may be realized also with cold gas from turbine directed along the
wall of the nozzle to protect it from hot gases from the thrust chamber. The protection of the
nozzle extension of the F-1 engine was performed in this way.
7. THE INJECTOR AND COMBUSTION STABILITY
The propellant is introduced inside the thrust chamber through the injector. The injector
forms a spray of the components to provide their effective mixing an burning. A typical injector
head consists of a plate with holes for the propellant components organized in a special pattern.
Some of the injector elements may represent sleeves sticking out from the plate, they may have
multiple holes. The injector head may also be divided into sections by partitions.
Stability of the combustion process in the chamber depends highly on the effectiveness of
mixing and thus, on the injector. The size of propellant drops and the parameters of the spray
define the lifetime of the drops, intensity of their evaporation and the quality of the burning
mixture. There is a number of reasons why burning process may become unstable, and the

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combustion may become resonant. For example, pressure pulsation may influence the injection
system: raise of pressure inside the chamber slows down the injection rate, and the following
drop of pressure (when the excess of the propellant leaves the chamber) leads to a new increase
of the injection rate. Self-oscillating process thus establishes with the frequency from tens to
hundreds cycles per second. This instability is in strong dependence on the lifetime of propellant
drops, i.e. on the delay between the propellant injection and its combustion. This instability is
often eliminated by changing pressure drop on the injector; an injector pressure drop usually
makes ~1/4 of the chamber pressure.
Another type of instabilities is high-frequency combustion instability (frequencies > 500
cycles per second), it is the most dangerous and is specially pronounced in engines having high
thrust. These instabilities arise because the time of drops vaporization is not constant and
depends on the pressure near the injector head. The higher is this pressure, more intensively
vaporize the drops, the combustion process accelerates, and a shockwave spreads along the
chamber, reflects from the opposite wall and returns, raising the pressure even more. The period
of the oscillations depends on the time in which the shock front returns. These oscillations
usually destroy a thin-wall chamber within seconds. This kind of oscillations is suppressed by
changing the chamber length and width, by installing additional partitions inside the chamber
which divide it into smaller volumes, and by matching the injector head (number and position of
the holes, the sleeves etc.) The designers of the F-1 engine, the biggest one-chamber engine ever
used, faced the high-frequency instabilities and had many problems with them. The problems
were solved by matching the injector head: it was divided into sectors by partitions. Vaporization
of the components before introducing them into the chamber also may be applied. A radical way
to solve the problem is to replace one large thrust chamber by several smaller ones: due to
smaller dimensions of the single chambers, the danger of appearance of high-frequency
instabilities is smaller. This way was chosen by the designers of RD-170/171: the engine has 4
chambers with the thrust of ~200 tons per chamber. Smaller size of the chamber also permits to
decrease the length of the engine. High-frequency instabilities make creation of high-thrust LREs
quite problematic and expensive.
There is another instability, low-frequency combustion instability (typically 10 – 100
cycles per second), not directly related to the injector. It appears due to resonance between the

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thrust caused by changes of the fuel flow rate, and and proper frequencies of the tanks and
structure of the rocket. The result is so-called pogo oscillations of the rocket. This kind of
oscillations is often eliminated by dampers in propellant lines, like in the Space Shuttle and
Saturn V. To damp oscillations, a small amount of helium is introduced into the propellant line to
shift the natural frequency of the line and to destroy the resonance.
8. THROTTLING OF THE ENGINE. START AND CUT-OFF
8.1 THROTTLING
In general, LREs are throttled by adjustment of the amount of the propellant delivered
into the thrust chamber, and this idea may seem to be quite simple. However, in practice it is not
so simple to realize. If the engine is pump-fed, we should take into account the fact that the
turbopumb is driven by the propellant, and if the flow rate is decreased, there may be lack of
reaction mass to power the turbine. The pumps are also designed for a certain propellant flow,
and significant change of the flow rate may lead to significant drop of efficiency, and the
turbopump assembly would work in unbalanced conditions: the power of the turbine may be
insufficient to drive the pumps. Of course, these problems may be overcome technically, but that
would mean impossibility to provide optimum conditions for the turbopump: as everywhere in
technical sciences, an universal unit is usually not so effective through the whole wide range of
working conditions as an specialized unit could be.
When the problems of the turbopump are solved (or in the case of a pressure-fed engine),
problems with cooling jacket may arise (if cooling is regenerative). Drop of propellant flow rate
may slow down and liquid may begin to vaporize, thus the engine would stall. The amount of the
coolant available also would drop. The temperature in the thrust chamber will remain nearly
unchanged if the mixture ratio is intact, and it may become impossible to cool down the walls. A
possible solution is to change mixture ratio and thus to decrease the temperature in the thrust
chamber, but this would lead to decrease of Isp (which would drop anywhere since the pressure in
the chamber will drop). Another solution is to use only passive cooling methods (ablation), but
that would mean lower temperatures and a shorter lifetime of the engine.
But the turbopump and cooling are not the only issues to take into account. The mayor
problem is combustion stability. When the propellant flow rate decreases, the injector pressure

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drop falls quicker than the pressure inside the chamber, and, as we have seen, the flow rate
becomes dependent on small variations of pressure in the chamber. A feedback between the
chamber pressure and the propellant supply appears and combustion becomes unstable. To avoid
that problem, variable-geometry injectors may be used: the area of the injector head is decreased
when the flow rate drops (this was the case of the Apollo LM Descent Stage).
Generally, most of LREs may work with moderate throttling by several percent without
stalling nor serious loss of efficiency. But deep throttling requires special designs and is a
difficult problem to solve. The highest throttling range among the human-rated engines was the
TRWS for the Descent Stage of the Apollo LM. It could be throttled down to 10%. However, the
range of ~65%  95% was unusable due to stability issues. SSME is throttled in-flight in the
range of 65%  105% (and a little bit wider range is available). RD-180 (a two-chamber version
of RD-170) is throttled in the range of ~40%  100%.
8.2 ENGINE START
To start a LRE, several operations should be performed in a right sequence. First of all,
the tanks should be pressurized. This is done with separate gas (like helium stored in special
vessels) or vaporized propellant components. In the last case the vaporized components are
available only after the engine is started, so temporarily some other gas is used. If cryogenic
components are used, specially LH2, the plumbing should be chilled with a small initial flow of
the component before the full flow is opened. This is done to prevent boiling of the cryogenic
component inside the plumbing.
If turbopump is applied, the turbines should be started to begin delivery of the propellant
into the chamber. This may be performed by burning main components in the gas generator: the
initial amount of propellant is directed into the gas generator to gasify. Turbines may also be
started by initial flow of a separate gas like helium stored in a starting vessels. When turbines are
able to pump components, the flow of the propellant into the chamber is initiated by opening
main vents.
If propellants are not hypergolic, they should be ignited. For a single-burn engines, like
that of launch vehicles, chemical ignitors are often used : a small amount of hypergolic
propellant is introduced into the gas generator and the thrust chamber before the main

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propellants, and they ignite the propellant mixture (often this hypergolic component self-ignites
when mixed with the main oxidizer). Ignition with a pyrotechnic charge or even a torch
introduced into the thrust chamber also may be applied. For multiple-burn engines electric
ignition (with spark) is often used.
To avoid hard start (when the pressure in the thrust chamber rises too quickly, damaging
the chamber), the start of a large engine should be performed carefully. Sometimes the start is
performed in two stages: at first the engine is started at a fraction of the full thrust, and when the
thrust is raised to the nominal value. The engine may also be started at a mixture ratio different
from the nominal, and then the ratio is set to the nominal.
8.3 ENGINE CUT-OFF
Large engines cannot be cut-off by only closing main propellant valves. Since turbines
continue rotating for some time, the pressure in the pump tract will not drop immediately. To
avoid raise of pressure at the main valves, the flow may be redirected to a low pressure line. If
the engine is pump-feed, at first the propellant flow to the gas generator is decreased, and the
turbines rotation slows down. Large engines are often cut-off in two stages to avoid rapid
transient processes that might damage the engine and the rocket. Pressure-fed engines may be
cut-off by propellant depletion: the flow of one of the components stops, and the engine shuts
down itself.

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9. 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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10. REFERENCES
1. G. P. Sutton, Rocket propulsion elements, 7th edition.
2. K. Ramamurthy ,Advances in propulsion.
3. M. J. Turner, Rocket and Spacecraft Propulsion.
4. O. Gurliat, V. Schmidt, O.J. Haidn,
M.Oschwald, Ignition of cryogenic H2/LOX sprays .
5. Indian Cryogenics Council (2010), Indian
Journal of Cryogenics, Vol. 35A, ISSN 0379-0479.
6.Richard Cohn (2012), Developments in
Liquid Rocket Engine Technology, Air Force Research Laboratory.