Rocket propulsion report (2)

INTRODUCTION
TO
ROCKET
PROPULSION

1-INTRODUCTION-
A rocket engine, or simply “rocket”, is a jet engine that uses only stored
propellant mass for forming its high speed propulsive jet. Rocket engines are
reaction engines and obtain thrustin accordancewith Newton’s third law.
Most rocketengines are internal combustion engines.
Rocket engines as a group havethe highest thrust, areby far the lightest, but
are the least propellant efficient (havethe lowest specific impulse) of all types
of jet engines. Rocket engines become more efficient at high velocities .Since
they do not benefit fromair but only suffer fromair resistance, they are best
suited for uses in spaceand the high atmosphere. Only very few rocketplanes
have been built for special uses in the low atmosphere.
Rocket technology can combine very high thrust (mega Newton), very high
exhaust speeds (around 10 times the speed of sound in air at sea level) and
very high thrust/weightratios (>100).
2-HISTORY OF ROCKET ENGINES-
According to the writings of the Roman Aulus Gellius, in 400 BC, a Greek
Pythagorean named Archytas, propelled a wooden bird along wires using
steam. However,itwould not appear to havebeen powerfulenough to take off
under its own thrust.Theaeolipile described in the firstcentury BC (often
known as Hero’sengine) essentially consists of a steam rocketon a bearing. It
was created almost two millennia beforethe IndustrialRevolution but the
principles behind it were not well understood, and its full potential was not
realized for a millennium.The availability of black powder to propel projectile
was a precursor to the development of the first solid rocket. Ninth Century
Chinese Taoist alchemists discovered black powder in a search for the Elixir of
life; this accidental discovery led to fire arrows which werethe firstrocket
engines to leave the ground.
Rocket engines were also brought in useby Tipu Sultan, the king of Mysore.
These rockets could be of various sizes, butusually consisted of a tube of soft
hammered iron about 8 in (20 cm) long and 1 1⁄2–3 in (3.8–7.6 cm) diameter,

closed at one end and strapped to a shaftof bamboo about 4 ft (120 cm) long.
The iron tube acted as a combustion chamber and contained well packed black
powder propellant. A rocketcarrying about one pound of powder could travel
almost 1,000 yards (910 m). Theserockets, fitted with swords used to travel
long distance, several meters abovein air before coming down with swords
edges facing the enemy. These rockets wereused againstBritish empire very
effectively.
Slow development of this technology continued up to the later 19th century,
when the writings of the Russian Konstantin Tsiolkovsky firsttalked about
liquid fueled rocketengines. He was the firstto develop the Tsiolkovsky rocket
equation, though it was not published widely for someyears . The modern
solid- and liquid-fueled engines became realities early in the 20th century,
thanks to the American physicistRobert Goddard. Goddard was the firstto
use a De Lavalnozzle on a solid-propellant(gunpowder) rocketengine,
doubling the thrustand increasing the efficiency by a factor of about twenty-
five. This was the birth of the modern rocket engine. He calculated fromhis
independently-derived rocket equation that a reasonably sized rocket, using
solid fuel, could place a one-pound payload on the Moon. He began to use
liquid propellants in 1921 and was the firstto launch, in 1926, a liquid
propellant rocket. Goddard pioneered the useof the De Lavalnozzle,
lightweight propellant tanks, thrustvectoring, the smoothly-throttled liquid
fuel engine, regenerative cooling, and curtain cooling.
During the late 1930s, German scientists, such as Wernher von Braun and
Hellmuth Walter, investigated installing liquid-fueled rockets in military aircraft
The turbopump was firstemployed by German scientists In WWII. Untilthen
cooling the nozzlehad been problematic, and the A4 ballistic missile used
dilute alcohol for the fuel, which reduced the combustion temperature
sufficiently.Staged combustion was firstproposed by Alexey Isaev in 1949. The
firststaged combustion engine was the S1.5400 used in the Soviet planetary
rocket, designed by Melnikov, a former assistantto Isaev. Aboutthe sametime
(1959), NikolaiKuznetsov began work on the closed cycle engine NK-9 for
Korolev’s orbitalICBM, GR-1. Kuznetsov later evolved that design into the NK-
15 and NK-33 engines for the unsuccessfulLunar N1 rocket.In the West, the
firstlaboratory staged-combustion testengine was built in Germany in 1963,
by Ludwig Boelkow.
Hydrogen peroxide / kerosenefueled engines such as the British Gamma of the
1950s used a closed-cycleprocess (arguably not staged combustion,) by

catalytically decomposing the peroxide to drive turbines before combustion
with the kerosenein the combustion chamber proper. This gave the efficiency
advantages of staged combustion, whilstavoiding the major engineering
problems. Liquid hydrogen engines were firstsuccessfully developed in
America, the RL-10 engine first flew in 1962.Hydrogen engines wereused as
part of the ProjectApollo; the liquid hydrogen fuel giving a rather lower stage
mass and thus reducing the overall sizeand cost of the vehicle. The Space
Shuttle Main Engine has the highest specific impulse of ground-launched
rocket engines to fly. On the 15th of July 2013, engineers at NASA successfully
tested a fully 3D-printed injector, demonstrating that even critical rocket
components can be safely produced and used with a 3D printer.
3- PRINCIPAL OF OPERATION-

Rocket thrustis caused by pressures acting in the combustion chamber and
nozzle. FromNewton’s third law, equal and opposite pressures acton the
exhaust, and this accelerates it to high speeds. Rocket engines produces the
part of their thrustdue to unopposed pressureon the combustion chamber.
Rocket engines producethrustby the expulsion of exhaust which has been
accelerated to a high-speed. The exhaust is usually a fluid and nearly always a
gas which is created by high pressure(10-200 bar) combustion of solid or liquid
propellants, consisting of fuel and oxidiser components, within a combustion
chamber. The fluid exhaustis then passed through a supersonic propelling
nozzlewhich uses heat energy of the gas to accelerate the exhaust to very high
speed, and the reaction to this pushes the engine in the oppositedirection. In
rocket engines, high temperatures and pressures arehighly desirable for good
performanceas this permits a longer nozzleto be fitted to the engine, which
gives higher exhaust speeds, as well as giving better thermodynamic efficiency.
4-POWER CYCLE OF ROCKET ENGINE-

Some of the cycle used in rocket propulsion are-
1-pressurefed cycle
2-gas generator cycle
3-staged combustion cycle
4-expander cycle
5- WORKING MECHANISM IN STAGES-
5.1-INTRODUCINGPROPELLENT INTO COMBUSTION
CHAMBER-
Rocket propellant is mass that is stored, usually in some formof propellant
tank, prior to being ejected froma rocket engine in the formof a fluid jet to
producethrust. Chemical rocketpropellants are mostcommonly used, which
undergo exothermic chemical reactions which producehot gas which is used
by a rocketfor propulsivepurposes. Alternatively, a chemically inert reaction
mass can be heated using a high-energy power sourcevia a heat exchanger,
and then no combustion chamber is used
5.2-COMBUSTIONCHAMBER-
For chemical rockets the combustion chamber is typically justa cylinder .The
dimensions of the cylinder are such that the propellant is able to combust
thoroughly; different propellants require different combustion chamber sizes
for this to occur. The combination of temperatures and pressures typically
reached in a combustion chamber is usually extreme by any standards. Unlike
in air-breathing jet engines, no atmospheric nitrogen is presentto dilute and
cool the combustion, and the temperature can reach true stoichiometric
ratios. This, in combination with the high pressures, means thatthe rate of
heat conduction through the walls is very high.

5.3- ROCKET NOZZEL-
Rocket nozzledoes the function of converting enthalpy ( h+pv)
into kinetic energy. The large bell or cone shaped expansion nozzlegives a
rocket engine its characteristic shape. In rockets the hot gas produced in the
combustion chamber is permitted to escape fromthe combustion chamber
through an opening (the “throat”), within a high expansion-ratio 'de Laval'
nozzle. When sufficientpressureis provided to the nozzle(about 2.5-3xabove
ambient pressure) thenozzle chokesand a supersonic jetis formed,
dramatically accelerating the gas, converting mostof the thermal energy into
kinetic energy. The exhaust speeds vary, depending on the expansion ratio the
nozzleis designed to give, but exhaustspeeds as high as ten times the speed of
sound at sea level air are not uncommon. Abouthalf of the rocketengine’s
thrustcomes fromthe unbalanced pressures insidethe combustion chamber
and the rest comes from the pressures acting against the inside of the nozzle
(see diagram). As the gas expands (adiabatically) the pressureagainstthe
nozzle’s walls forces the rocketengine in one direction while accelerating the
gas in the other.
6-PROPELLANTEFFICIENCY-
For a rocket engine to be propellant efficient, it is important
that the maximum pressures possiblebecreated on the walls of the chamber
and nozzleby a specific amount of propellant; as this is the sourceof the
thrust. This can be achieved by all of the following:-
1- Heating the propellant to as high a temperature as possible(using a high
energy fuel, containing hydrogen and carbon and sometimes metals such as
aluminum, or even using nuclear energy)
2- Using a low specific density gas (as hydrogen rich as possible)
3- Using propellants which are, or decomposeto, simple molecules with few
degrees of freedom to maximize translationalvelocity
Since all of these things minimize the mass of the propellant used, and since
pressureis proportionalto the mass of propellant present to be accelerated as
it pushes on the engine, and since from Newton’s third law the pressurethat
acts on the engine also reciprocally acts on the propellant, it turns out that for
any given engine the speed that the propellant leaves the chamber is

unaffected by the chamber pressure(although the thrustis proportional).
However speed is significantly affected by all three of the abovefactors and
the exhaustspeed is an excellent measureof the engine propellant efficiency.
This is termed exhaustvelocity, and after allowance is made for factors
that can reduce it, the effective exhaust velocity is oneof the mostimportant
parameters of a rocketengine (although weight, cost, ease of manufactureetc.
are usually also very important).
For aerodynamic reasons the flow goes sonic ("chokes") atthe narrowestpart
of the nozzle, the 'throat'. Since the speed of sound in gases increases with the
squareroot of temperature, the useof hot exhaust gas greatly improves
performance. By comparison, at roomtemperature the speed of sound in air is
about 340 m/s while the speed of sound in the hot gas of a rocketengine can
be over 1700m/s; much of this performanceis due to the higher temperature,
but additionally rocket propellants are chosen to be of low molecular mass,
and this also gives a higher velocity compared to air. Expansion in the rocket
nozzlethen further multiplies the speed, typically between 1.5 and 2 times,
giving a highly collimated hypersonic exhaustjet. The speed increaseof a
rocket nozzleis mostly determined by its area expansion ratio—the ratio of the
area of the throat to the area at the exit, but detailed properties of the gas are
also important. Larger ratio nozzles are more massivebutare able to extract
more heat from the combustion gases, increasing the exhaust velocity. Nozzle
efficiency is affected by operation in the atmosphere because atmospheric
pressurechanges with altitude; but due to the supersonic speeds of the gas
exiting froma rocket engine, the pressureof the jet may be either below or
above ambient, and equilibrium between the two is not reached at all altitudes
.
7-BACK PRESSURE AND OPTIMAL
EXPANSION-
For optimal performancethe pressureof the gas at the end of the nozzle
should justequal the ambient pressure: if the exhaust’s pressureis lower than
the ambient pressure, then the vehicle will be slowed by the difference in
pressurebetween the top of the engine and the exit; on the other hand, if the
exhaust’s pressureis higher, then exhaust pressurethat could havebeen
converted into thrustis not converted, and energy is wasted.

To maintain this ideal of equality between the exhaust’s exit pressureand the
ambient pressure, thediameter of the nozzlewould need to increase with
altitude, giving the pressurea longer nozzleto act on (and reducing the exit
pressureand temperature). This increase is difficult to arrangein a lightweight
fashion, although is routinely done with other forms of jet engines. In rocketry
a lightweight compromisenozzleis generally used and somereduction in
atmospheric performanceoccurs when used at other than the 'design altitude'
or when throttled. To improveon this, various exotic nozzledesigns such as
the plug nozzle, stepped nozzles, the expanding nozzleand the aerospikehave
been proposed, each providing someway to adapt to changing ambient air
pressureand each allowing the gas to expand further against the nozzle, giving
extra thrustat higher altitudes.
When exhausting into a sufficiently low ambient pressure(vacuum) several
issues arise. Oneis the sheer weight of the nozzle—beyond a certain point, for
a particular vehicle,the extra weight of the nozzleoutweighs any performance
gained. Secondly, as the exhaust gases adiabatically expand within the nozzle
they cool, and eventually someof the chemicals can freeze, producing 'snow'
within the jet. This causes instabilities in the jet and mustbe avoided.
8-DESIGN CONSIDERATIONS-
8.1-SPECIFIC IMPULSE-
The most important metric for the efficiency of a rocketengine is impulse per
unit of propellant, this is called specific impulse (usually written Isp ). This is
either measured as a speed (the effective exhaust velocity ve in metres/second
or ft/s) . An engine that gives a large specific impulse is normally highly
desirable. The specific impulse that can be achieved is primarily a function of
the propellant mix (and ultimately would limit the specific impulse), but
practical limits on chamber pressures and the nozzleexpansion ratios reduce
the performancethat can be achieved.
8.2-NET THRUST-
Below is an approximate equation for calculating the net
thrustof a rocket engine-

Fn = m_ ve = m_ ve�act + Ae(pe -pamb)
Since, unlike a jet engine, a conventionalrocket motor lacks an air intake, there
is no 'ramdrag' to deduct fromthe gross thrust. Consequently the net thrustof
a rocket motor is equal to the gross thrust(apartfromstatic back pressure).
The m_ ve�act term represents the momentum thrust, which remains
constantat a given throttle setting, whereas the Ae(pe - pamb) term
represents the pressurethrust term. At full throttle, the net thrust of a rocket
motor improves slightly with increasing altitude, becauseas atmospheric
pressuredecreases with altitude, the pressurethrustterm increases. At the
surfaceof the Earth the pressurethrustmay be reduced by up to 30%
depending on the engine design. This reduction drops roughly exponentially to
zero with increasing altitude. Maximum thrust for a rocket engine is achieved
by maximizing the momentum contribution of the equation without incurring
penalties fromover expanding the exhaust. This occurs when pe = pamb . Since
ambient pressurechanges with altitude, most rocket engines spend very little
time operating at peak efficiency.
8.3-THROTTLING-
Rockets can be throttled by controlling the propellant combustion rate m_
(usually measured in kg/s or lb/s). In liquid and hybrid rockets, the propellant
flow entering the chamber is controlled using valves, in solid rockets it is
controlled by changing the area of propellant that is burning and this can be
designed into the propellant grain (and hence cannot be controlled in real-
time).
Rockets can usually be throttled down to an exit pressureof about one-third of
ambient pressure(often limited by flow separation in nozzles) and up to a
maximum limit determined only by the mechanical strength of the engine. In
practice, the degree to which rockets can be throttled varies greatly, but most
rockets can be throttled by a factor of 2 without great difficulty; the typical
limitation is combustion stability, as for example, injectors need a minimum
pressureto avoid triggering damaging oscillations but injectors can often be
optimised and tested for wider ranges. Solid rockets can be throttled by using
shaped grains that will vary their surfacearea over the courseof the burn.

8.4-THRUSTTO WEIGHT RATIO-
Rockets, of all the jet engines, indeed of essentially all engines, have the
highest thrustto weight ratio. This is especially true for liquid rocketengines.
This high performanceis due to the small volume of pressurevessels that
make up the engine—the pumps,pipes and combustion chambers involved.
The lack of inlet duct and the useof dense liquid propellant allows the
pressurisation systemto be small and lightweight,whereas ductengines have
to deal with air which has a density about one thousand times lower.Of the
liquid propellants used, density is worstfor liquid hydrogen. Although this
propellant is marvellous in many ways, it has a very low density, about one
fourteenth that of water. This makes the turbopumps and pipework larger and
heavier, and this is reflected in the thrust-toweightratio of engines that use it
(for example the SSME) compared to those that do not (NK-33).
8.5-ENERGY EFFICIENCY-
Rocket engine nozzles aresurprisingly efficient heat engines for generating a
high speed jet, as a consequenceof the high combustion temperature and high
compression ratio. Rocket nozzles give an excellent approximation to adiabatic
expansion which is a reversibleprocess, and hence they give efficiencies which
are very close to that of the Carnotcycle. Given the temperatures reached,
over 60% efficiency can be achieved with chemical rockets. For a vehicle
employing a rocketengine the energetic efficiency is very good if the vehicle
speed approaches or somewhatexceeds the exhaust velocity (relative to
launch); but at low speeds the energy efficiency goes to 0% at zero speed (as
with all jet propulsion

9-COOLING-
For efficiency reasons, and becausethey physically can, rockets run with
combustion temperatures that can reach ~3500 K (~3227 °Cor ~5840 °F).
Most other jet engines havegas turbines in the hot exhaust. Due to their larger
surfacearea, they are harder to cool and hence there is a need to run the
combustion processes atmuch lower temperatures, losing efficiency. In
addition, duct engines useair as an oxidant, which contains 78% largely
unreactive nitrogen, which dilutes the reaction and lowers the temperatures.
Rockets have none of these inherent disadvantages. Thereforetemperatures
used in rocketsare very often far higher than the melting point of the nozzle
and combustion chamber materials. Two exceptions are graphite and
tungsten (~1200 K for copper), although both are subject to oxidation if not
protected. Indeed many construction materials can make perfectly acceptable
propellants in their own right. Itis important that these materials be prevented
fromcombusting, melting or vaporising to the point of failure. This is
sometimes somewhatfacetiously termed an 'engine rich exhaust'. Materials
technology could potentially place an upper limit on the exhaust
temperature of chemical rockets. Alternatively, rockets may use more common
construction materials such as aluminium, steel, nickel or copper alloys and
employ cooling systems thatprevent the construction material itself becoming
too hot. Regenerative cooling, where the propellant is passed through tubes
around the combustion chamber or nozzle, and othertechniques, such as
curtain cooling or film cooling, are employed to give longer nozzle and
chamber life. These techniques ensurethat a gaseous thermal boundary layer
touching the material is kept below the temperature which would cause the
material to catastrophically fail. In rockets, the heat fluxes that can pass
through the wall are among the highest in engineering. The strongestheat
fluxes are found at the throat, which often sees twice that found in the
associated chamber and nozzle. This is due to the combination of high speeds
(which gives a very thin boundary layer).
9.1-COOLING METHODS-
In rockets the cooling methods include:-

1. Uncooled (used for shortruns mainly during testing)
2. Ablative walls (walls are lined with a material that is continuously vaporised
and carried away).
3. Radiative cooling (the chamber becomes almost white hot and radiates the
heat away).
4. Dump cooling (a propellant, usually hydrogen, is passed around the chamber
and dumped)
5. Regenerative cooling (liquid rockets use the fuel, or occasionally the
oxidiser, to cool the chamber via a cooling jacketbefore being injected).
6. Curtain cooling (propellant injection is arranged so the temperature of the
gases is cooler at the walls)
7. Film cooling (surfaces arewetted with liquid propellant, which cools as it
evaporates).
In all cases the cooling effect that prevents the wall from
being destroyed is caused by a thin layer of insulating fluid (a boundary layer)
that is in contact with the walls that is far cooler than the combustion
temperature. Provided this boundary layer is intact the wall will not be
damaged. Disruption of the boundary layer may occur during cooling failures
or combustion instabilities, and wall failure typically occurs soon after.
With regenerative cooling a second boundary layer is found in the coolant
channels around the chamber. This boundary layer thickness needs to be as
small as possible, since the boundary layer acts as an insulator between the
wall and the coolant. This may be achieved by making the coolant velocity in
the channels as high as possible. In practice, regenerative cooling is nearly
always used in conjunction with curtain cooling and/or film cooling. Liquid-
fueled engines are often run fuel rich, which results in lower temperature
combustion. Cooler exhaust reduces heat loads on the engine allowing lower
cost materials, a simplified cooling system, and a lower performanceengine.

10-ENGINE DIAGRAM-
This labeled diagram showsvariouspartsof space shuttle main engine (SSME)
NOTE-Details of partsand their respective functionsis skipped/avoided asthese
are beyond the scope of seminar

11-ENGINEERING PROBLEMS
ENCOUNTERED -
11.1-MECHANICAL ISSUES-
Rocket combustion chambers are normally operated at fairly high pressure,
typically 10-200 bar (1 to 20 MPa, 150-3000 psi). When operated within
significant atmospheric pressure, higher combustion chamber pressures
give better performanceby permitting a larger and more efficient nozzleto be
fitted without it being grossly overexpanded. However, thesehigh pressures
causethe outermostpart of the chamber to be under very large hoop stresses
– rocketengines are pressurevessels. Worse, dueto the high temperatures
created in rocketengines the materials used tend to havea significantly
lowered working tensile strength. In addition, significant temperature
gradients are set up in the walls of the chamber and nozzle, these cause
differentialexpansion of the inner liner that create internal stresses.
11.2-ACOUSTICS ISSUES-
The extreme vibration and acoustic environmentinside a rocketmotor
commonly result in peak stresses wellabove mean values, especially in the
presenceof organ pipe-like resonances and gas turbulence.
11.3-COMBUSTION INSTABILITIES-
The combustion may display undesired instabilities, of sudden or periodic
nature. The pressurein the injection chamber may increase until the
propellant flow through the injector plate decreases; a moment later the
pressuredrops and the flow increases, injecting more propellant in the
combustion chamber which burns a moment later, and again increases the
chamber pressure, repeating the cycle. This may lead to high-amplitude
pressureoscillations, often in ultrasonic range, which may damage the motor.

Oscillations of ±200 psiat 25 kHz werethe causeof failures of early versions of
the Titan II missilesecond stage engines. The other failure mode is a
deflagration to detonation transition; the supersonic pressurewaveformed in
the combustion chamber may destroy the engine .The combustion instabilities
can be provoked by remains of cleaning solvents in the engine, reflected shock
wave, initial instability after ignition, explosion near the nozzle that reflects
into the combustion chamber, and many more factors. In stableengine designs
the oscillations are quickly suppressed; in unstable designs they persist for
prolonged periods. Oscillation suppressors arecommonly used.
Periodic variations of thrust, caused by combustion instability or longitudinal
vibrations of structures between the tanks and the engines which modulate
the propellant flow, are known as "pogo oscillations" or “pogo”, named
after the pogo stick.
Three different type of combustion instabilities which
occur are-
11.3.1-CHUGGING-
This is a low frequency oscillation at a few Hertz in chamber pressureusually
caused by pressurevariations in feed lines due to variations in acceleration of
the vehicle. This can cause cyclic variation in thrust, and the effects can vary
frommerely annoying to actually damaging the payload or vehicle. Chugging
can be minimised by using gasfilled damping tubes on feed lines of high density
propellants.
11.3.2-BUZZING-
This can be caused due to insufficientpressuredrop across theinjectors. It
generally is mostly annoying, rather than being damaging. However, in
extreme cases combustion can end up being forced backwards through the
injectors–this can causeexplosions with monopropellants.
11.3.3-SCREECHING-
This is the most immediately damaging, and the hardest to control. Itis due to
acoustics within the combustion chamber that often couples to the chemical

combustion processes thatare the primary drivers of the energy release, and
can lead to unstable resonant“screeching” that commonly leads to
catastrophic failure due to thinning of the insulating thermal boundary layer.
Acoustic oscillations can be excited by thermal processes, such as the flow of
hot air through a pipe or combustion in a chamber. Specifically, standing
acoustic waves inside a chamber can be intensified if combustion occurs more
intensely in regions where the pressureof the acoustic wave is maximal. Such
effects are very difficult to predict analytically during the design process, and
have usually been addressed by expensive, time consuming and extensive
testing, combined with trial and error remedial
correction measures.
Screeching is often dealt with by detailed changes to injectors,
or changes in the propellant chemistry, or vaporizing
the propellant beforeinjection, or useof Helmholtz
dampers within the combustion chambers to change the
resonantmodes of the chamber.
11.4-EXHAUST NOISE-
For all but the very smallestsizes, rocketexhaust compared to other engines is
generally very noisy. As the hypersonic exhaustmixes with the ambient air,
shock waves areformed. The SpaceShuttle generates over 200 dB(A) of noise
around its base.The Saturn V launch was detectable on seismometers a
considerabledistance fromthe launch site. The sound intensity fromthe shock
waves generated depends on the sizeof the rocket and on the exhaust
velocity. Such shock waves seemto account for the characteristic crackling and
popping sounds produced by large rocket engines when heard live. These noise
peaks typically overload microphones and audio electronics, and so are
generally weakened or entirely absent in recorded or broadcastaudio
reproductions. For largerockets at close range, the acoustic
effects could actually kill. More worryingly for spaceagencies, such sound
levels can also damage the launch structure, or worse, be reflected
back at the comparatively delicate rocketabove.This is why so much water is
typically used at launches. The water spray changes the acoustic qualities of
the air and reduces or deflects the sound energy away fromthe rocket.
Generally speaking, noiseis most intense when a rocket is close to the ground,
since the noise fromthe engines radiates up away fromthe plume, as well as
reflecting off the ground. Also, when the vehicle is moving slowly, little of the

chemical energy input to the engine can go into increasing the kinetic energy
of the rocket (sinceuseful power P transmitted to the vehicle is P = F _ V for
thrustF and speed V). Then the largestportion of the energy is dissipated in
the exhaust’s interaction with the ambient air, producing noise. This noise can
be reduced somewhatby flame trenches with roofs, by water injection around
the plume and by deflecting the plume at an angle.
12-SAFETY-
Rockets have a reputation for unreliability and danger; especially catastrophic
failures. Contrary to this reputation, carefully designed rockets can be made
arbitrarily reliable. In military use, rockets are not unreliable. However,
one of the main non-military uses of rockets is for orbital launch. In this
application, the premiumhas typically been placed on minimum weight, and it
is difficult to achieve high reliability and low weight simultaneously. In
addition, if the number of flights launched is low, there is a very high chance of
a design, operations or manufacturing error causing destruction of the vehicle.
Essentially all launch vehicles are test vehicles by normalaerospace
standards (as of 2006). TheX-15 rocketplane achieved a 0.5% failure rate,
with a single catastrophic failure during ground test, and the SpaceShuttle
Main Engine managed to avoid catastrophic failures in over 350 engine-flights.
13-CHEMISTRY OF ROCKETS-
Rocket propellants requirea high specific energy (energy per unit mass),
because ideally all the reaction energy appears as kinetic energy of the exhaust
gases, and exhaust velocity is the single most important performance
parameter of an engine, on which vehicle performancedepends. Aside from
inevitable losses and imperfections in the engine, incomplete combustion, etc.,
after specific reaction energy, the main theoretical limit reducing the exhaust
velocity obtained is that, according to the laws of thermodynamics, a fraction
of the chemical energy may go into rotation of the exhaust molecules, whereit
is unavailable for producing thrust. Monatomic gases like helium
have only three degrees of freedom, corresponding to the three dimensions of
space, {x,y,z}, and only such spherically symmetric molecules escapethis kind
of loss. A diatomic molecule like H2 can rotate about either of the two axes

perpendicular to the one joining the two atoms, and as the equipartition law of
statistical mechanics demands that the available thermal energy be divided
equally among the degrees of freedom, for such a gas in thermal equilibrium
3/5 of the energy can go into unidirectional motion, and 2/5 into rotation. A
triatomic molecule like water has six degrees of freedom, so the energy is
divided equally among rotational and translational degrees of freedom. For
most chemical reactions the latter situation is the case. This issue is
traditionally described in terms of the ratio, gamma, of the specific heat of the
gas at constantvolume to that at constantpressure.Therotational energy loss
is largely recovered in practice if the expansion nozzle is large enough to allow
the gases to expand and cool sufficiently, the function of the nozzlebeing to
convertthe randomthermal motions of the molecules in the combustion
chamber into the unidirectional translation that produces thrust. As long as
the exhaustgas remains in equilibrium as it expands, the initial rotational
energy will be largely returned to translation in the nozzle. Although the
specific reaction energy per unit mass of reactants is key, low mean molecular
weight in the reaction products is also important in practice in determining
exhaust velocity. This is because the high gas temperatures in rocket engines
poseserious problems for the engineering of survivablemotors. Because
temperature is proportionalto the mean energy per molecule, a given
amount of energy distributed among more molecules of lower mass permits a
higher exhaustvelocity at a given temperature. This means low atomic mass
elements are favoured. Liquid hydrogen (LH2) and oxygen (LOX, or LO2), are
the mosteffective propellants in terms of exhaust velocity that havebeen
widely used to date, though a few exotic combinations involving boron or
liquid ozoneare potentially somewhatbetter in theory if various practical
problems could be solved. Itis important to note in computing the specific
reaction energy, that the entire mass of the propellants, including both fuel and
oxidizer, mustbe included. The fact that airbreathing engines are typically able
to obtain oxygen “for free” without having to carry it along, accounts for one
factor of why air-breathing engines are very much more propellant-mass
efficient, and one reason that rocketengines are far less suitable for most
ordinary terrestrialapplications.Fuels for automobile or turbojet engines,
utilize atmospheric oxygen and so havea much better effective energy output
per unit mass of propellant that must be carried, but are similar per unit mass
of fuel. Computer programs that predict the performanceof propellants
in rocket engines are available.

14-FUTUREOF ROCKET PROPULSION-
In futurefollowing methods will be used for rocketpropulsion-
1-IONTHRUSTERS
2-SOLARSAIL
3-NUCLEARPOWER
ION THRUSTER

Rocket propulsion report (2)

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