it teaches u what is liquid propulsion and the propellants used in it as well as how these propellants are fed to thrust chamber and so on.

1. Tallinn University of Technology
Introduction to Astronautics
Sissejuhatus kosmonautikasse
Vladislav Pustõnski
2009

2. Liquid propellant rocket engines
General characteristics
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 main advantages of liquid propellants:
• high specific impulse;
• high thrust;
• high thrust to weight ratio of the rockets;
• easy to control.
However, they also have certain disadvantages:
• complexity of the construction;
• it is impossible to achieve very high specific impulse;
• they are difficult to scale, complexity grows quickly with growing the thrust;
• some propellants are highly toxic (hydrazine and its variants) or cryogenic (hydrogen).
2

3. 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.
Types of propellants
There are two basic types of LRE propellants: monopropellants and bipropellants.
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
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

4. 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 4

5. N2O4/MMH in its main engine. 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 –1830C, thus being moderately cryogenic: this temperature is
higher than the boiling point of air (– 1940C), 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 –253 0C.
5

6. 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
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.
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.
6

7. 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).
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 (if the propellant
is hypergolic, even no ignitor is needed). To avoid the pressurizing gas to cool down due to
expansion inside the tanks, it is often warmed up in the heat exchanger.
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
7

8. 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 2 nd
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.
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
8

9. have enormous 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 ~5000C. 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.
9

10. 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, 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).
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 10

11. closed cycle (contrary to the F-1), as well as the SSME of the Shuttle.
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 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.
11

12. 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.
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. Different
methods may be applied simultaneously in different parts of the chamber and the nozzle.
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
12

13. 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.
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). 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
13

14. 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.
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
combustion may become resonant. For example, pressure pulsation may influence the
14

15. 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 15

16. 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
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.
Throttling of the engine. Start and cut-off
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.
16

17. 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 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%.
17

18. 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
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
18

19. 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.
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.
19

20. Physics of a rocket engine and nozzle
In the previous lecture we have seen that the thrust T of a rocket
T = qVmean + S a ( pa − p )
engine is defined by the formula
Here q = dm/dt – mass flow rate, S – area of the nozzle, pa – static pressure of the exhaust
gases, p – ambient air pressure. If pa > p, the nozzle works with underexpansion, if pa < p,
the nozzle works with overexpansion, pa = p is the case of ideal expansion. Increase of the
expansion ratio ε (the ratio of the pressure inside the chamber pch and the pressure at the
nozzle exit pa, ε = pch/pa) leads to an increase of the first summand and to a decrease of the
second summand. It may be proved that the case of ideal expansion provides the highest
thrust possible for the corresponding external pressure. As the rocket ascends, the ambient
pressure changes, so it is impossible to provide optimal conditions for all heights. A nozzle
that has overexpansion at lower heights (after the lift-off) will work in underexpansion
conditions after the liftoff. In vacuum, every nozzle works in underexpansion conditions,
since it is impossible to provide zero pressure at the nozzle exit (for that, the nozzle expansion
should have infinite length and with). But larger is the nozzle, less is the underexpansion.
However, it makes no sense to increase very much the size of the nozzle, since the gain of
efficiency would be cancelled by grow of the nozzle size and weight. For the first stage, the
expansion ratio of the nozzle cannot be very large, since too high overexpansion would lead
to separation of the jet of the exhaust gases from the nozzle. That may lead to catastrophic
drop of efficiency and even damaging the nozzle, if measures are not taken. Generally, the
nozzle of the first stages work in overexpansion conditions in the lower atmosphere. Nozzles
of upper stages work at significantly lower ambient pressures, so they 20 much larger
have
expansion ratios than the

21. nozzles of the first stages. The expansion ratio of a nozzle is a compromise between thrust
efficiency at different heights, weight and size. In the recent time an original design have
been implemented in several engines: the nozzle extension changes inflight to improve
efficiency while the rocket ascends from the denser atmosphere to more rarefied layers. This
solution is applied on the second stage of the Delta IV rocket (LOX/LH2 engine RL-10B-2).
The shape of the rocket nozzle is of de Laval type. It is convergent-divergent, having a
throat. In the convergent section of the nozzle the gas flow is subsonic, it velocity increases
towards the throttle and achieves the sound speed at the throttle. In the divergent section the
gas jet expands and achieves supersonic speeds.
The gas escapes in slightly different directions in different parts of the nozzle. This fact
may be taken into account with the coefficient ϕ that relates the Vmean 1 + cos α
escape may velocity to the actual escape velocity Va and the half-angle of Vmean = Va
2
conicity of the nozzle α:
2λ RTch  
λ −1
From gas dynamics, the gas escape velocity may be found as
Va = k 1 − ε λ 
R – gas constant, ε – expansion ratio, T – temperature inside the λ −1 µ  


ch
thrust chamber, λ – adiabatic index, µ – molar weight of the exhaust gases, k – imperfection
coefficient that takes into account incomplete combustion, film cooling etc.
It is seen that higher escape velocity are provided with higher temperatures in the chamber
and lower molar weights of the exhaust. The dependence of the expansion ratio has a
21

22. maximum at ε = 0, i.e. the pressure inside the thrust chamber should be possibly high and the
outer pressure should be possibly low.
The term in the brackets is η – expansion efficiency, Vc = (λRTch /µ )1/2 V = kV 2
a c η
is the sound speed. So formula the escape velocity may be given a form λ −1
For typical propellants λ = 1.25, and the escape velocity is ~2.8 Vc (if k ~ 1 and η ~ 1).
22

23. End of the Lecture 9
23

24. Launches of Proton & Titan II
Launch of Proton. 6 engines, Launch of Titan. 2 engines
N4H4/UDMH. The plume is N4H4/aerozine. (By source)
nearly invisible (By source)

25. Launches of Soyuz & Saturn IB
Launch of a Soyuz. 5 engines, Saturn IB launch. 8
LOX/kerosene. The plume is engines, LOX/RP-1
yellowish bright (By source) (By source)

26. Launches of Ariane V & Delta IV
Launch of Ariane V. 2 Delta IV Medium launch. 1
SRB + 1 LOX/LH2. The engine, LOX/LH2. The
plume of the LOX/LH2 is color is caused by nozzle
nearly invisible (By source) ablative covering (By source)

27. Scheme of the pressure-fed engine
(By source)

28. MA-5 turbopump assembly
Turbopump of MA-5 engine for the Atlas. (By source)

29. Scheme of the gas generator cycle
(By source)

30. Scheme of the staged combustion cycle
(By source)

31. Scheme of the expander cycle
(By source)

32. Thrust chamber
1) injector; 2) thrust chamber body; 3) nozzle part; 4)
throat; 5) holes; A, B) propellant components. (By
V.I.Feodosjev, Osnovy tehniki raketnogo poleta.)

33. Cooling jacket
1) outer jacket; 2) inner cooled jacket; 3) soldering; 4) cooling
liquid flow; 5) hot gases; (a) typical size several mm. (By
V.I.Feodosjev, Osnovy tehniki raketnogo poleta.)
Outer wall of F-1 engine for
the Saturn V. (By source)

34. Cooling jacket of F-1
Inside view of the F-1 engine (Saturn V). Bifurcating
tubes are clearely visible (By source)

35. Injector heads
Injector head of the HM-7
engine (LOX/LH2, Ariane
upper stages). Sleeves for one
of the components are inside
the holes for the another
component (By source)
Injector head of the RD-170/171
engine (LOX/kerosene,
Energia/Zenit first stage).
Sleeves for one of the
components form partitions (By
source)

36. F-1 and RD-170
F-1 (vacuum thrust ~800 tons) 4-nozzle 4-chamber RD-170
(By source) (vacuum thrust ~810 tons) (By
source)

37. Nozzle expansion at different heights
(By source)

38. de Laval nozzle
Diagram of the velocity V,
pressure P and temperature
T in a de Laval nozzle. M is
the Mach number
(By source)