PROPULSION II

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Propulsion - II Notes DepartmentofAeronautical Engineering
AE6504 PROPULSION - II
SEMESTER : V

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SYLLABUS
AE6504 PROPULSION - II L T P C
3 0 0 3
OBJECTIVES:
 To impart knowledge in non air-breathing and hypersonic propulsion methods to students so that
they are familiar with various propulsion technologies associated with space launch vehicles,
missiles and space probes.
UNIT I HYPERSONIC AIRBREATHING PROPULSION 8
Introduction to hypersonic air breathing propulsion, hypersonic vehicles and supersonic combustion need
for supersonic combustion for hypersonic propulsion – salient features of scramjet engine and its
applications for hypersonic vehicles – problems associated with supersonic combustion – engine/airframe
integration aspects of hypersonic vehicles – various types scramjet combustors – fuel injection schemes in
scramjet combustors – one dimensional models for supersonic combustion using method of influence
coefficients.
UNIT II FUNDAMENTALS OF CHEMICALROCKET PROPULSION 9
Operating principle – specific impulse of a rocket – internal ballistics – performance considerations of
rockets – types of igniters- preliminary concepts in nozzle-less propulsion – air augmented rockets – pulse
rocket motors – static testing of rockets & instrumentation –safety considerations
UNIT III SOLID ROCKET PROPULSION 10
Salient features of solid propellant rockets – selection criteria of solid propellants – estimation of solid
propellant adiabatic flame temperature - propellant grain design considerations – erosive burning in solid
propellant rockets – combustion instability – strand burner and T-burner – applications and advantages of
solid propellant rockets.
UNIT IVLIQUID AND HYBRID ROCKET PROPULSION 10
Salient features of liquid propellant rockets – selection of liquid propellants – various feed systems and
injectors for liquid propellant rockets -thrust control and cooling in liquid propellant rockets and the
associated heat transfer problems – combustion instability in liquid propellant rockets – peculiar problems
associated with operation of cryogenic engines - Introduction to hybrid rocket propulsion – standard and
reverse hybrid systems- combustion mechanism in hybrid propellant rockets – applications and
limitations

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UNIT V ADVANCED PROPULSIONTECHNIQUES 8
Electric rocket propulsion– types of electric propulsion techniques - Ion propulsion – Nuclear rocket –
comparison of performance of these propulsion systems with chemical rocket propulsion systems – future
applications of electric propulsion systems - Solar sail.
TOTAL: 45 PERIODS
OUTCOMES
 Understanding various propulsion systems
 Knowledge in rocket propulsion systems
 Knowing the applications and principles of liquid and solid-liquid propulsion systems
 Application of nuclear propulsion in rocketry.
TEXT BOOKS:
1. Sutton, G.P., “Rocket Propulsion Elements”, John Wiley & Sons Inc., New York, 5th
Edition,
1993.
2. Mathur, M.L., and Sharma, R.P., “Gas Turbine, Jet and Rocket Propulsion”, Standard Publishers
and Distributors, Delhi, 1988.
REFERENCES:
1. James Award,"Aerospace Propulsion System"
2. Hieter and Pratt,"Hypersonic Air Breathing Propulsion"

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UNIT I
HYPERSONIC AIRBREATHING PROPULSION
Introduction to hypersonic air breathing propulsion- hypersonic vehicles and supersonic
combustion need for supersonic combustion for hypersonic propulsion – salient features of
scramjet engine and its applications for hypersonic vehicles – problems associated with
supersonic combustion – engine/airframe integration aspects of hypersonic vehicles – various
types scramjet combustors – fuel injection schemes in scramjet combustors – one dimensional
models for supersonic combustion using method of influence coefficients.
1.1 INTRODUCTION TO HYPERSONIC AIR BREATHING PROPULSION
DEFINING HYPERSONIC FLOW
A precise definition of hypersonic flow conditions is elusive because the term merely
connotes very high Mach numbers and is inherently qualitative. For propulsion engineers,
defining and locating a useful boundary is a relatively straightforward matter.
In subsonic aerodynamics most of the total temperature is invested in static enthalpy, and
changes of Mach number are largely due to changes of velocity. In transonic aerodynamics there
are substantial fractions of both static enthalpy and kinetic energy, and the Mach number
changes rapidly because both the speed of sound and the velocity are changing. In the hypersonic
region the bulk of the total temperature is invested in kinetic energy, and the Mach number
changes largely because the static temperature and speed of sound are changing.
This leads to one internal flow or propulsion-based definition of the hypersonic regime,
namely that this is where most of the total temperature exists as kinetic energy, and changes in
Mach number have little effect on the kinetic energy or velocity of the flow.
INTRODUCTION TO SCRAMJET
At higher temperature around 2500K, the walls of ramjet will tend to fail structurally.
Thus, like turbojets, conventional ramjets are also limited by material problem at high flight
Mach numbers. Moreover, if the temperature of the air entering the combustor is too high, when
the fuel is injected, it will be decomposed by the higher temperatures rather than be burned; that

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is, the fuel will absorb rather than release energy, and the engine will become a drag machine
rather than a thrust producing device.
The velocity of the incoming air, as seen from the frame of reference of the vehicle or
engine, also represents relative kinetic energy. When the air flow is decelerated by the scramjet,
the relative velocity and kinetic energy both decrease and observation of energy requires that any
missing kinetic energy will reappear as internal energy, with the result that the pressure,
temperature, and density of the flow entering the burner are considerably higher than in the free
stream. When the flight Mach number exceeds about 6, this effect becomes so pronounced that it
is no longer advantages to decelerate the flow to subsonic speeds. Depending upon flight
conditions and details of the diffuser operation, the adverse consequences can include pressures
too high for practical burner structural design, excessive performance losses due to normal shock
wave system, excessive wall heat transfer rates, and combustion conditions that large fraction of
the available chemical energy to dissociation.
A logical way to solve this problem is to only partially compress and decelerate the
oncoming flow, avoiding in particular the normal shock wave system, with the result that the
flow entering the burner is supersonic, the resulting engine is known as a supersonic combustion
ramjet or scramjet.
Even though the diffuser is responsible for some of the desired compression and deceleration
match of it is invariably accomplished by oblique shock waves emanating from the vehicle fore-
body located upstream of the engine. This allows the engine to take advantage of the inevitable
compression of the free stream by the vehicle and reduces the burdens upon the diffuser.
Moreover, since, the diffuser exit flow is supersonic the geometry is entirely convergent. Fuel is

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injected in to the supersonic flow just downstream of the diffuser, and the emphasis is upon
achieving rapid and through mixing (especially when all the entering oxygen is to be consumed)
because the time available for the combustion process is short. The heat loads are highest in the
burner primarily because of the combination of the high gas density due to compression, rather
than to the ongoing combustion. The exhaust nozzle need only be divergent because the
accelerating glow is supersonic throughout, and some of the acceleration can take place outside
the confining duct by using the after body of the vehicle as a free expansion surface.
1.2 HYPERSONIC VEHICLES AND SUPERSONIC COMBUSTION NEED FOR
SUPERSONIC COMBUSTION FOR HYPERSONICPROPULSION
Clearly for hypersonic flight at very high Mach numbers, something else must be done.
This problem has led to the concept of a “supersonic combustion ramjet” and the scramjet. Here,
the flow entering the diffuser is at high Mach number, say Ml=M =6. However, the diffuser
decelerates the air flow only enough to obtain a reasonable pressure ratio P2/P1; the flow is still
supersonic upon entering the combustor. Fuel is added to the supersonic stream, where
supersonic combustion takes place. In this way, the flow field throughout the scramjet is
completely supersonic; in turn, the static temperature remains low, and the material and
decomposition problems associated with the conventional ramjet are circumvented. Therefore
the power plant for a hypersonic transport in the future will most likely be a scramjet.
For the hypersonic regime, it has been proposed that the initial diffusion should be
limited and that combustion should occur at supersonic velocities. If then the heat is added when
M>1, the process occurs along the lower branch of a Rayleigh line and may again take place
until the checking point is reached. The Rayleigh lines are different ones, the specific mass flow,
G=m/A = v being greater for this case although the heat added is less and the efficiency is
lower. It is possible that shock is an adiabatic process and at constant area, all states must be on
the initial Rayleigh line, and constant line, and the total amount of heat added is not changed.
The balance between a conventional ramjet (CRJ) and the supersonic combustion ramjet (SCR
or Scramjet) lies in the efficiencies of the processes (e.g. diffusion) and the initial temperature
level (ie., Mach number). Also these high temperatures, the degree of frozen or equilibrium flow
can have a considerable effect and this is dependent on the fuel. It should be noted that subsonic

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combustion necessitates a decrease of static pressure (for Acceleration of the flow) and that
supersonic combustion implies an increase of static pressure (decelerating of the flow) but both
are accompanied by a loss of stagnation pressure. As in indication of performance, Fig 2.10
shows a comparison of a CRJ and a SCRJ at a particular condition.
The successful burning at high velocity is very difficult and the situation is an order of
magnitude higher or more severe. The temperature of course is very much higher, which
promotes very fast combustion. Pressures too are high, again helpful. A fuel like hydrogen is
almost necessary and so is a flow situation which permits some stability. It would seem that a
standing shock wave somewhere is required, produced by wedges on the duct walls or along its
axis or at the discharge of a premixed fuel.
The advantage of the SCRJ, in addition to the lower initial temperature and absence of the
dissociation sink, is in having lower pressures and no subsonic diffusion, hence a saving of
weight in length and strength.
1.3 SALIENT FEATURES OF SCRAMJET ENGINE AND ITS APPLICATIONS FOR
HYPERSONIC VEHICLES
The scramjet uses a slightly modified Brayton Cycle to produce power, similar to that used
for both the classical ramjet and turbine engines. Air is compressed; fuel injected, mixed and
burned to increase the air – or more accurately, the combustion products - temperature and
pressure; then these combustion products are expanded. For the turbojet engine, air is
mechanically compressed by work extracted from the combustor exhaust using a turbine. In
principle, the ramjet and scramjet works the same. The forward motion of the vehicle compresses
the air. Fuel is then injected into the compressed air and burned. Finally, the high-pressure
combustion products expand through the nozzle and over the vehicle after body, elevating the
surface pressure and effectively pushing the vehicle.
Thrust is the result of increased kinetic energy between the initial and final states of the
working fluid, or the summation of forces on the engine and vehicle surfaces.

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PARTS OF SCRAMJET ENGINE
The different parts of a scramjet engine: air inlet, isolator, combustor and nozzle. With the
actual technology, as it is mentioned in Chapter 3, the scramjet engine must be integrated with
the fuselage of the aircraft, specially the air inlet and the nozzle. Part of the forebody aircraft
fuselage makes the function of air inlet compressing the freestream air, and similarly, the aft
body acts as a nozzle expanding the gases from the combustion.
AIR INLET:
It can be considered as a diffuser in which takes place the compression of the freestream air
gathered. This compression is achieved by successive shock waves. The performance of the air
inlet compression can be separated into two key parameters: capability, or how much
compression is performed, and efficiency, or what level of flow losses does the inlet generate
during the compression process. A common parameter used to quantify the efficiency of the
forebody/inlet compression is the kinetic energy efficiency. The definition is the ratio of the
kinetic energy of the compressed flow would achieve if it were expanded isentropically to
freestream pressure, relative to the kinetic energy of the freestream.

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Hypersonic inlets used in scramjets fall into three-different categories, based on the type of
compression that is utilized. These three types are: external compression, mixed compression and
internal compression. A schematic of these types is shown in figure below.
In the external compression all the compression is performed by flow turning in one direction
by shock waves that are external to the engine. These inlet configurations have large cowl drag,
as the flow entering the combustor is at a large angle relative to the freestream flow; however,
external compression inlets are self-starting and spill flow when operated below the design Mach
number (this is a desirable feature for inlets that must operate over a large Mach number range).
In a mixed compression inlet the compression is performed by shocks both external and
internal to the engine, and the angle of the external cowl relative to the freestream can be made
very small to minimize external drag. These inlets are typically longer than external compression
configurations, but also spill flow when operated below the design Mach number. Depending on
the amount of internal compression, however, mixed compression inlets may need variable
geometry in order to start.
In internal compression inlet the compression is performed by shock waves that are internal
to the engine. This type of inlet can be shorter than a mixed compression inlet, but it does not
allow easy integration with the vehicle. It maintains full capture at Mach numbers lower than the
design point, but its most significant limitation is that extensive variable geometry is always
required for it to start.

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ISOLATOR:
At flight speeds below Mach 8, combustion in a scramjet engine can generate a large local
pressure rise and separation of the boundary layer on the surfaces of the combustion duct. This
separation, which can feed upstream of fuel injection, acts to further diffuse the core flow in the
duct, and will affect the operation of the inlet, possibly causing an unstart of the engine. The
method use to alleviate this problem is the installation of a short duct between the inlet and the
combustor known as an isolator. In some engines (those which operate in the lower hypersonic
regime between Mach 4 and 8) the combination of the diffusion in the isolator and heat release in
the combustion decelerate the core flow to subsonic conditions, in what is called dual-mode
combustion. At speeds above Mach 8 the increased kinetic energy of the airflow through the
engine means that the combustion generated pressure rise is not strong enough to cause boundary
layer separation. Flow remains attached and supersonic throughout, and this is termed pure
scramjet. In this case an isolator is not necessary.
The structure of the supersonic flow in confined ducts under the influence of a strong adverse
pressure gradient is of interest in the design of scramjet isolators. As shown in figure below, a
pressure gradient is imposed on the incoming supersonic flow, and with the presence of a
boundary layer, a series of crossing oblique shocks are generated.
This phenomenon, known as pseudo shock or shock-train, is characterized by a region of
separated flow next to the wall, together with a supersonic core that experiences a pressure
gradient due to the area restriction of the separation and forms the series of oblique shocks
mentioned before. Finally, the flow reattaches at some point and mixes out to conditions that

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match the imposed back-pressure. Being able to predict the length scale of this flow structure is
the key component of isolator design for dual-mode scramjets.
COMBUSTOR :
The combustor chamber is a duct where the combustion between freestream air and fuel takes
place. This combustion is supersonic, so there are some aspects that require more attention on the
contrary of the conventional combustion. At very high velocities, a properly fuel injection and
mixing could be a problem, as well as holding the flame. Some techniques used today for fuel
injection in scramjet engines are: wall, ramp, strut, pylon and pulsed injectors. And for keeping
the combustion, there is a technique quite used called cavity flame holders.
Another significant aspect to take into account is the dissociation. At the entrance of the
combustor the flow static temperature and pressure are very high, and with the heat release due
to chemical reactions, the temperature and pressure could reach extremely high values which
involve dissociation of combustion products.
Due to the heat addition, the velocity or Mach number decreases while the static temperature
and pressure increases. The total temperature is raised and the total pressure is reduced. The total
pressure loss is proportional to the square of Mach number; hence, it is better to have a small
combustor inlet Mach number, on the contrary for the dissociation phenomenon.
The fuel used in scramjet engines is hydrogen or hydrocarbons. Hydrogen is most used
because it has more advantages in front of hydrocarbons. The reason for using liquid hydrogen
for scramjet fuel rests with its high specific impulse and its potential for cooling parts of the
vehicle. The heat value (which represents the amount of energy released when a fuel is
combusted) for hydrogen is two and a half times that of hydrocarbons. Another advantage over
hydrocarbons is that hydrogen is a clean fuel as it doesn‟t produce any harmful pollutants like
carbon monoxide (CO) or carbon dioxide (CO2) during the combustion process.
Although it may appear that hydrogen is the ideal fuel for scramjet propulsion it does present
some drawbacks. Liquid hydrogen is not a dense fuel, having a density of only 0.09 kg/m3. For
example, JP-8 on the other hand has a density of 800 kg/m3 in similar conditions, very much

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higher. Having a low density does save weight; however, a large volume is needed in order to
store enough chemical energy for practical use.
NOZZLE:
The nozzle is a divergent duct that accelerates the supersonic flow and at the same time
expands it reducing its static temperature and pressure. The expansion process converts the
potential energy of the combusting flow to kinetic energy and then it results in thrust.
The weight of a fully-expanded nozzle would be prohibitive at most hypersonic flight
conditions; hence under-expansion losses are usually traded against vehicle structural weight.
Dissociation losses result from chemical freezing in the rapid expansion process in the nozzle,
essentially locking up energy that cannot be converted to thrust. Flow angularity losses are
product of varying flow conditions in the nozzle, and viscous losses are associated with friction
on the nozzle surfaces.
The choice of combustor inlet Mach number is a key aspect for the performance of the
scramjet and it is related to the nozzle expansion. If the static temperature at the combustor
entrance is too high, dissociation will be present and then chemical energy is not available as
thermal energy for conversion to kinetic energy in the nozzle.
APPLICATIONS FOR SCRAMJET ENGINES
There is a range of possible applications for scramjet engines, including missile propulsion,
hypersonic cruiser propulsion, and part of a staged space access propulsion system Figure below
displays the approximate performance range in terms of engine specific impulse and Mach
number for various types of propulsion systems . It can be seen that at Mach numbers higher than
approximately 6-7, the only available propulsion systems are rockets and scramjets. Compared to
rockets, scramjets have much higher specific impulse levels; therefore, it is clear why it is
advantageous to develop the scramjet, if for this reason only. Contrary to rockets, scramjets do
not require that an oxidizer be carried on board the aircraft as it is an air breathing engine,
collecting oxygen from the atmosphere. This decreases the required weight of the overall
propulsion system and fuel, resulting in a higher allowable payload weight or increased range.

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There are other reasons that the development of the scramjet is advantageous as well. Air
breathing engines produce higher engine efficiency, have a longer powered range, possess the
ability for thrust modulation to ensure efficient operation, have higher maneuverability, and are
completely reusable .
The general consensus is that hydrogen fuel should be used for air breathing flight faster than
Mach 8-10, due to its “higher cooling capacity” and its faster reactions. Though hydrogen can
perform at higher speeds above the hydrocarbon upper limit, with current capabilities the
hydrogen fueled scramjet will only offer acceptable performance to about Mach 15.
There are many advantages in applying the scramjet as the propulsion system for the second
stage of a two-stage-to-orbit (TSTO), hydrocarbon-fueled aerospace plane. It would provide for a
small TSTO vehicle as well as a small single-stage-to-orbit (SSTO) vehicle or military
hypersonic cruiser that uses a hydrocarbon-fueled scramjet.
The rationale for hypersonic missile capability lies in the fact that a Mach 6-8 missile
increases the possible range within a given flight time, or similarly, decreases the flight time
required for a given range.
The goal of scramjet development is to give hypersonic vehicles a more efficient alternative to
rockets. The vehicle that could most quickly benefit from current scramjet research is the cruise
missile; however, a hypersonic cruiser aircraft is an alternative to traditional turbojet

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transportation for civilian or military application. Scramjets could also be used in conjunction
with rockets for space launchers, thereby requiring less on-board oxidizer for transport to space.
1.4 PROBLEMS ASSOCIATED WITH SUPERSONIC COMBUSTION
At high flight speeds, the residence time for atmospheric air ingested into a scramjet inlet and
exiting from the engine nozzle is on the order of a millisecond. Therefore, fuel injected into the
air must efficiently mix within tens of microseconds and react to release its energy in the
combustor. The overall combustion process should be mixing controlled to provide a stable
operating environment; in reality, however, combustion in the upstream portion of the
combustor, particularly at higher Mach numbers, is kinetically controlled where ignition delay
times are on the same order as the fluid scale. Both mixing and combustion time scales must be
considered in a detailed study of mixing and reaction in a scramjet to understand the flow
processes and to ultimately achieve a successful design.
Airframe structural and heat transfer limitations constrain flight Mach numbers to specific
altitudes and corresponding freestream conditions. Cycle efficiency considerations, together with
temperature limitations imposed by materials and combustion product gas dissociation, dictate
the combustion system entry Mach number and thermodynamic state.
The maximum combustion temperature occurs when hydrocarbon fuel molecules are mixed
with just enough air so that all of the hydrogen atoms form water vapor H20, and all of the
carbon atoms form carbon dioxide C02.
Gas-phase chemical reactions occur by the exchange of atoms between molecules as a result
of molecular collisions. Consequently, fuel and air must be mixed to near-stoichiometric
proportions at the molecular level before combustion can take place.
Compressible shear/mixing layers and jets provide good model problems for studying the
physical processes occurring in high-speed mixing and reacting flow in a scramjet. Mixing layers
are characterized by large-scale eddies that form due to the high shear that is present between the
fuel and air streams. These eddies entrain fuel and air into the mixing region.

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The shock and expansion wave structure in and about the mixing layer can interact with the
turbulence flow field to affect mixing layer growth. Shock and expansion waves interacting with
the layer result from the engine internal structure. Experiments have shown that the shocks that
would result from wall and strut compressions appear to enhance the growth of the two-
dimensional eddy structure (rollers) of a mixing layer.
1.5 ENGINE/AIRFRAME INTEGRATION ASPECTS OF HYPERSONICVEHICLES
High speed flight requires proper aerodynamic integration of the ramjet or scramjet with the
remainder of the vehicle. It is found, for example, that making hypersonic engines axisymmetric
and attaching them to the vehicle by means of pylons or struts can produce enough external drag
on the pylon and the cowl to virtually cancel the internal thrust, and creates internal passages so
narrow that the flow is dominated by wall effects and difficult to manage. Furthermore, this
configuration cannot easily capitalize on the vehicle surfaces for compression and expansion.
The recent engineering design work on aerospace planes has shown that success is also
dependent upon careful integration of engine and airframe structures, materials, cooling,
controls, and subsystems. As the vehicles fly to the very high altitudes necessary to avoid the
excessive pressures that would result from hypersonic speeds, they also fly where the air is very
rare, the density being only one hundredth or one-thousandth or less of the sea level value. Air
breathing engines require airflow in order to generate the thrust that lifts and accelerates the
vehicle, and they cannot be allowed to suffocate.
One way to capture the required airflow is to use the entire fore body underneath the vehicle
as a compression surface, shaping it carefully for high efficiency, and recognizing that, since the
air has no warning of the presence of the vehicle until it encounters the first oblique shock wave,
a stream tube of air enormously larger than the physical opening of the engine inlet can be
directed into the engine, as shown in figure below. Once inside the engine, the air that has been
compressed and burned would then require a nozzle exit area even larger than the original free
stream capture area in order to make good on the available thrust, but a conventional nozzle of
this geometry would be too cumbersome to carry along. Instead, the entire after body underneath
the vehicle is used as a free expansion surface, and the entire underside of the vehicle must be
carefully designed to accommodate engine performance under all flight conditions.

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HYPERSONIC AIRBREATHING PROPULSION CHALLENGES
Operating efficiently and reliably over an extraordinarily large range of flight conditions,
including Mach numbers from 0 to 25 (orbital speed) and altitudes from sea level to the top of
the atmosphere.
It is the fact that practical aerospace plane configurations allow only one "hole" for the engine,
and that ramjets and scramjets produce no thrust while standing still or when the atmosphere is
too thin, and further complicated because the airframe and engine are totally integrated, not only
with regard to generating net thrust, but also inevitably from the standpoint of the ordinary
"aerodynamic" effects such as lift and drag forces and stability and control moments.
The current challenges in the development of the scramjet engine can be gathered in three
main areas: air inlet, combustion, and structures and materials.
Air inlet :
The overall performance of a scramjet is largely dictated by the aerodynamic
performance, geometric size, and weight of the hypersonic inlets. Commonly, hypersonic inlets
have a wide Mach number range, but the shock-on-lip condition can be met only at the design
Mach number, since shock angles vary with the upstream Mach numbers. Thus, at Mach
numbers higher than the design one, the ramp shocks move inside the inlet and evolve into a
strong incident shock, causing strong slip layers, remarkable total pressure loss, boundary-layer
separation, and possible engine unstart. At Mach numbers lower than the design one, the ramp
shocks move away from the cowl lip, causing loss of the precompressed airflow and the so-
called spillage drag. To avoid these performance penalties at offdesign conditions, the control of

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the ramp shock system is needed. Hence, variable geometric approaches for ramp shock control
are widely considered and studied.
Combustion :
Supersonic combustion is very difficult to maintain and continues to be a formidable task. the
largest problem associated with combustion is the mixing between freestream air and fuel. If fuel
cannot be properly injected and mixed into the air stream it will not ignite, regardless of pressure,
temperature or equivalence ratio. Due to compressibility effects, fuel injection presents
challenging obstacles. The air stream is at such a high pressure and velocity, that fuel injected
into the stream has a tendency to be pushed against the wall and rendered ineffective. In addition
to the problem of mixing, ignition and flame holding at these high velocities is extremely
difficult. Another challenge to increase the performance is the need of a variable geometry
combustion chamber. A fixed geometry combustor associated to a variable capture area air inlet
does not benefit from the enhanced efficiency of the air inlet. A fully variable geometry – air
inlet + combustion chamber – can increase the performance.
Structures and materials :
A rocket that passes nearly vertically through the atmosphere on its way to orbit, a scramjet
would take a more leveled trajectory. Because of the thrust-to-weight ratio of a scramjet being
low compared to modern rockets, the scramjet needs more time to accelerate. Such a depressed
trajectory implies that the vehicle stays a long time in the atmosphere at hypersonic speeds,
causing atmospheric friction to become a problem. This is not only for space launch applications
but also in missile or commercial transport applications. Heat addition produced by the
combustion at these high velocities and temperatures is another significant factor to take into
account. Therefore, the materials chosen for the structure must have good properties and be
adequate in front of these phenomena. Furthermore, cooling of the engine‟s structure by fuel or
radiation is essential.
Summarize the challenges of SCRAMJET development:
 Accomplishing stable, efficient mixing and combustion in a supersonic flow within a
burner of reasonable size.
 Providing the structural integrity necessary for a reusable system despite the extremely
hostile environmental conditions.

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 Developing the analytical tools that enable confident control over the engine design and
reliable prediction of the actual behavior.
 Proving that the aerospace plane and engine are ready for routine operations by means of
analysis, ground testing, and flight testing of experimental vehicles.
1.6. VARIOUS TYPES OF SUPERSONIC COMBUSTORS
The use of supersonic combustion can result in higher levels of performance at increased
flight Mach number than can be achieved with subsonic combustion.
DESIGN CONSIDERATIONS OF SUPERSONIC COMBUSTOR:
Supersonic combustor considerations essentially dictate many of the inlet requirements,
and the resulting combustor exit flow chemistry can also limit key exhaust nozzle options.
A. Inlet Flow:
For a subsonic combustion ramjet the captured supersonic airflow is first decelerated in a
supersonic diffuser to a lower supersonic Mach number, then brought to subsonic conditions
through a normal shock system, and finally diffused further to a lower subsonic Mach number.
This action is typically accomplished by a supersonic diffuser employing a number of oblique
shocks and a throat section followed by a subsonic diffuser. Obviously for a scramjet inlet, only
a supersonic diffuser is required. Consequently, the question of how much supersonic diffusion
or inlet compression should be provided at the entrance of the supersonic combustor becomes an
important issue. The answer involves integration of the design of the inlet, combustor, and
nozzle such that the required engine performance is achieved at acceptable weight, operational
reliability, and cost.

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B. COMBUSTOR FLOW:
Heat release rate of hydrogen over a wide range of pressure and temperature, often
further accelerated by the presence of local shock waves from nonaxial fuel injection into the
airstream, the actual combustion pressure rise can be expected to be rather abrupt with an
initially high adverse pressure gradient .This sharp pressure gradient coupled with the
impracticality of scramjet inlet Boundary Layer removal for such high-temperature flow raises
the concern of Boundary Layer separation.
The assumption of 100% mixing implies that provided adequate combustor length ER = 1
mixing of the fuel and air is achieved homogeneously in the combustor. The assumption of
equilibrium chemistry implies that for steady flow the specie concentrations are not (or no
longer) a function of time (forward and backward reactions occur at the same rate). Because the
individual chemical reactions require a finite time for completion as a function of pressure and
temperature, the preceding calculations also implied that the combustor was long enough for the
finite rate chemical reactions to reach equilibrium. Consequently, if the combustor length is not
long enough, then the net heat release for stoichiometric combustion of hydrogen is reduced.
TYPES OF COMBUSTORS:
1. STEP COMBUSTOR:
One design approach to the Boundary Layer separation problem is the use of a step
combustor as schematically illustrated in figure below. The resulting A^/A2 > 1 provides a basis
for accommodating the combustor pressure rise. Fuel is typically injected in close proximity to
the base of the step, which can also provide small subsonic recirculation regions for ignition and
flame holding. Boundary-layer separation occurs downstream of the station 2 interface but is
only a concern in that the resulting fuel injection and combustion-induced step pressure must not
subject the station 2 Boundary Layer to such a high back pressure that significant upstream
separation results.

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The step combustor can be a design solution when Boundary Layer separation is likely to
occur in a constant-area combustor. The step area increase provides a mechanism for
accommodating wall forces resulting from fuel injection and combustion pressures Pstep>P2
without separating the inlet Boundary Layer.
2. ISOLATOR COMBUSTORS:
Another design solution when Boundary Layer separation is likely to occur in a constant
area combustor is the use of an inlet isolator in conjunction with an increased area combustor.
The increased area combustor could be a step combustor, a divergent-area wall combustor, or a
combination. The isolator section is intended to provide the length required for precombustion
increase of the inlet pressure to the combustor pressure level without unstarting the inlet.
Typically, such backpressuring of the viscous inlet flow is accomplished by a shock train (close-
coupled series of shock-wave Boundary Layer interactions). The isolator is designed to contain
this shock train, thus preventing it from extending further upstream into the inlet compression
system. The higher the combustor back pressure, the greater the throttling required, and
consequently the smaller is the flow area A2A required to satisfy the conservation equations at
the isolator exit.

21
1.7 FUEL INJECTION SCHEMES IN SCRAMJET COMBUSTORS
For given inlet conditions the net heat release achieved in a scramjet combustor is driven
by the efficiency and effectiveness of the fuel injection. Efficiency is reflected in the degree of
fuel/air mixing achieved; effectiveness is associated with minimization of the combustor exit
stream thrust losses incurred in the mixing process and the extent of the additional wall cooling
or thermal protection risk associated with the fuel injector.
A. WALL JETS:
Wall jets, although minimizing intrusion of the combustor flow path, result in a relatively
complex flow structure in the immediate vicinity of the jets as characterized shown below in
figure for normal injection through a round hole on a flat plate. The resulting regions of high-
pressure gradient on the wall near the jet are also a source of increased local heat flux.
When in-line jets are spaced too far apart, the benefit of increased momentum flux per unit
frontal area is likely to diminish, approaching single hole performance in the limit. On the other
hand, too close an axial spacing could also be a problem in terms of constraining fore and aft

22
expansion (high-pressure jet proximity) at the expense of the frontal area increasing lateral
expansion.
B. IN-STREAM INJECTORS:
Fuel jets can be vertically spaced as required to achieve desired local F/A distribution across
the entire combustor height, matching even non uniform flow air profiles. For fuel injected from
the sidewalk or out of the top of in-stream injectors, much of the previous discussion of wall jets
applies. For fuel injected axially downstream from the trailing edge, full axial fuel momentum is
also achieved but at what can be a significant mixing rate penalty. At high combustor velocities
such axial injection can lead to increased distance (combustor length) to achieve adequate
mixing. The major issues are thermal protection and structural integrity.
C. HYPERMIXERS:
In hypermixer type axial downstream fuel injection permits maximum momentum recovery
from the fuel (the importance of which increases with flight Mach number). Vortex shedding
from the corners, step-type recirculation behind the aft surfaces, and impingement of the
reflected ramp shock just aft of the injectors are relied upon to enhance the otherwise relatively
slow mixing of the axially injected hydrogen.

23
The fuel injector ramp and reflected shocks are a source of non combustion entropy increase and
unlike well-designed inlet compression ramps they do not provide permanent increased inlet
contraction for the losses incurred. Intentional creation of vortices for mixing enhancement also
introduces additional axial momentum losses. the challenge of the hypermixer concept is to
achieve adequate mixing of near axially injected fuel in reasonable combustor lengths without
increasing drag.
D. MIXING:
Scramjet combustors primarily dependent on free shear-layer mixing (such as unenhanced
simple axial fuel injection) can be expected to require an unacceptably long length to achieve
acceptable mixing. Fuel injector designs that introduce curved shock waves and local Boundary
Layer separation increase the rotationality of the flow and the losses induced provide near-field
mixing over that of free shear flow. Combustion in the near and far field, along with induced
vortices, also increases mixing relative to free shear flow.
1.8 ONE DIMENSIONAL MODELS FOR SUPERSONIC COMBUSTION USING
METHOD OF INFLUENCE COEFFICIENTS.
The sole function of the isolator is to prevent inlet unstart, by providing sufficient additional
adiabatic compression above its entry pressure p2 to match or support whatever back pressure P3
.There are two very different physical mechanisms which cause a back pressure p3 > p2 to arise at
burner entry: in scramjet mode, thermal blockage unrelieved by area expansion which causes
unwanted flow separation, and in ramjet mode, the required thermal choking which insures that
the flow into the burner will be subsonic.
For the case of frictionless flow without mass addition, but with change in both cross-sectional
area A and total temperature Tt due to heat "addition," the governing ordinary differential
equation (ODE) for axial variation of Mach number is given by
When M > 1, it can be seen from inspection of Eq. (6-90) that occluding the flow, either by
decreasing A or by increasing Tt, causes M to decrease in the axial direction. Conversely,
relieving the flow, either by increasing A or decreasing Tt, will cause M to increase. However, in

24
the dual-mode combustion system, A(x) will never decrease with x, as there is no physical throat,
and Tt(x) will never decrease.
A(x) and Tt(x) are chosen as independent variables, and their coefficients are the influence
coefficients of the respective independent variables on the single dependent variable M. for
supersonic combustion (scramjet mode), both mixing and chemical heat release rates are greatest
at onset, and relax asymptotically toward their respective fully-mixed and chemical equilibrium
(or kinetically frozen) zero-rate values with infinite convective time or distance.
The axial total temperature gradient dTt/dx is greatest shortly after ignition, usually near
burner entry, and decreases to its least value at burner exit. For a wide variety of scramjet mixers,
and to represent both mixing transition delay and induction or ignition delay, Tt(x) can be
usefully represented in non dimensional form by a rational function given by
Tb is the overall total temperature rise ratio in the burner.
Ө is an empirical constant of order 1 to 10 which depends on the mode of fuel injection
and fuel-air mixing.
To determine a unique solution ordinary differential equation (ODE), it is necessary to specify
the burner entry Mach number M3, as well as the forcing functions A(x) and Tt(x).
The following additional equations are useful for plotting the burner process path on H-K
coordinates:
Most analyses of supersonic combustion have assumed constant-A combustion, as the
constant-A geometry is both easy to fabricate and to analyze. an adverse pressure gradient is
created whenever heat is added to a frictionless, supersonic flow in a constant-A burner, and if
the pressure rise is too great, it will cause separation of the boundary layer. It is important to note
that cause-and-effect relationships are somewhat different in the combustion system.
In the case of a frictionless constant-Area (A) burner, the causal chain is: (1) the pressure rises
as a result of heat addition to the supersonic flow, and (2) if the pressure rise exceeds some
threshold value, the boundary layer separates, so that (3) the oncoming supersonic flow is turned

25
into itself by the effective area blockage of the separated flow near the wall, and is compressed
into a confined core flow through an oblique shock train, until the confined core flow pressure
matches the pressure in the region of separated flow in the burner.

26
UNIT II
FUNDAMENTALS OF CHEMICAL ROCKET PROPULSION
Operating principle – specific impulse of a rocket – internal ballistics – performance
considerations of rockets – types of igniters- preliminary concepts in nozzle-less propulsion – air
augmented rockets – pulse rocket motors – static testing of rockets & instrumentation –safety
considerations
2.1 OPERATING PRINCIPLE:
A rocket propulsion device, though working on the same principle as that of a jet engine i e.
of obtaining a propulsive force as a reaction to the acceleration of a mass of fluid, characterizes
itself in that it carries, with it, its own supply of oxidant. this enables the rocket to operate in
higher altitudes and even at vacuum.
A rocket motor is a device for converting the thermo chemical energy of one or more
propellants into exhaust jet kinetic energy: the term “propellants” is applied to any material.
Solid or liquid, consumed in the rocket motor. It is generally assumed, and there is some
evidence to support the assumption, that under the pressures and temperatures occurring in a
rocket motor combustion chamber and nozzle the chemical reactions take place under conditions
which approach chemical equilibrium.
The chemicals utilized for producing the high-pressure, high – temperature gases in a
rocket motor are either solids or liquids, but for certain special applications the reaction of gases,
like gaseous hydrogen with gaseous oxygen, must be useful. Regardless of the type of propeller
(or propellant), the main objective is to produce the largest possible jet velocity. Since the
pressure of the gases enthalpy the exhaust nozzle ranges from 250 to 3000psi and the back
pressure is 14.7psi or less, the exhaust nozzle always operates under supercritical conditions.
Hence if the isentropic flow is assumed the gas speed in the throat of the nozzle will be the
critical speed of sound a*, which is dependant of the basic pressure. Furthermore, to obtain the
largest possible values of jet speed, the exhaust gases are always the converging-diverging the
exhaust gases are always ejected with supersonic speed.

27
Thrust :
Thrust is useful to begin by examining the performance of a rocket under static tests. Consider
the thrust of a stationary rocket indicated schematically in figure below for simplicity, the flow
may be assumed one dimensional, with a steady exit velocity Ve and propellant flow rate &
consider a stationary control surface “S” which interests the jet perpendicularly through the exit
plane of the nozzle. Positive thrust “T” acts in the opposite direction to Ve. The reaction to the
thrust is shown in fig as it acts on the control volume. If the expelled fluid can be considered a
continuum, it is necessary to consider the pressures just inside the exit plane of the nozzle, pe ,the
cross-sectional area of the jet Ae is the exit area of the nozzle.
The momentum equation for any such control volume M is,
∑ ∫ ∫ (2.1)
Where
Fx = Force component in the x direction
= density of fluid
Ux = velocity component of the fluid in x direction
v = Volume
m = mass flow rate of flow (positive for out flow)

28
In words the resultant force acting on the control volume is equal to the time rate of
increase of linear momentum with in the control volume flux the nut efflux of linear momentum
from the control volume.
Where, the subscripts CV and CS denote the control volume and control surface
respectively.
Since Ux is zero within the propellant tank or tanks, and the flow is steady within the thrust
chamber, the time derivative term is zero. Also the momentum – flux can be written as
∑ ∫ (2.2)
Considering the pressure on the control surface to be uniformly Pa. expect the plane of the jet,.
The force summation may be written
∑ - (2.3)
Thrust is actually a result of pressure or stress distribution over interior and exterior surface, as
shown typically for a chemical rocket in figure.
Combining equation 1,2 and 3, we obtain,
mVe = - (2.4)
If the pressure in the exhaust plane is the same as the ambient the pressure, the thrust is given by
= mVe
The condition pa =pe is called correct or optimum expansion because it corresponds to
maximum thrust for given chamber conditions. Conveniently, one can define an effective
exhaust velocity, Vj , such that
T = m Vj (2.5)
Where

29
Effective exhaust velocity, Vj= Ve + (Pe-Pa) (2.6)
2.2 SPECIFIC IMPULSE OF A ROCKET:
The impulse per unit mass of propellant will be shown to be an important performance
variable. If the effective exhaust velocity Vj is constant equation shows that the total impulse I
imparted to the vehicle during acceleration is
I= ∫ (2.7)
Where, m = the total mass of expelled propellant. This impulse per unit mass of propellant is
therefore,
= = (2.8)
He term specific impulse, Isp is usually defined by,
Isp = = (2.9)
Where, g is the acceleration due to gravity at the earth‟s surface. The presence of g in the
definition is arbitrary, but it does have the advantage that n all common systems of units the
specific impulse is expressed in seconds.
Analysis of an Ideal Rocket: Chemical rockets, whether powered by liquid or solid propellants,
consist in varying complexity of propellant supply and feed system, a combustion chamber, and
exhaust nozzle. To simplify our analysis of the thrust chamber, we assume the following.
1. The working substances (or chemical reaction products) is homogenous.
2. All the species of the working fluid is gaseous. Any condensed phases (liquid or solid)
add a negligible amount to the total mass.
3. The working substance obeys the perfect gas law.
4. These are no heat transfer across the rocket walls: therefore, the flow is adiabatic.
5. These are no appreciable friction and all boundary layer effects are neglected.
6. These are no shock waves or discontinuities in the nozzle flow.

30
7. The propellant flow is steady and constant. The expansion of the working fluid is uniform
and study, without vibration. Transient effects (i.e. start up and shut down) are of very
short duration and may be neglected.
8. All exhaust gases leaving the rocket have an axially directed velocity.
9. The gas velocity, pressure, temperature, and density are all uniform across any section
normal to the nozzle axis.
10. Chemical equilibrium is established with in the rocket chamber and the gas composition
does not change in the nozzle (frozen flow)
11. Stored propellants are at room temperature. Cryogenic propellants are at their boiling
points.
Assuming adiabatic nozzle expansion, the energy equation requires constant stagnation enthalpy
in the nozzle.
h02 = h0e
Assuming the expansion to be Isentropic,
(2.10)
Putting Cp in terms of universal gas constant R0 , molecular weight M, and ratio of specific heats
, we get
(2.11)
If Q is the heat supplied in the form of chemical energy per u nit mass of propellant, we get

31
(2.12)
From Eqn 2.11 we can write
(2.13)
If At is the area of the throat of the nozzle and Pt the pressure at throat it can be proved that the
mass flow rate is given by
(2.14)
The thrust produced is given by
(2.15)
putting eqn 2.13 & 2.14 in 2.15
IDEAL THRUST COEFFICIENT:
where pc = combustion chamber pressure,
At = nozzle throat area
t
T
F
A
F
C
c
P


32
Depends primarily on (pc/pa) so a good indicator of nozzle performance – dominated by
pressure ratio.
CHARACTERISTIC VELOCITY (C*):
• Calculated from standard test data.
• It is independent of nozzle performance and is therefore used as a measure of
combustion efficiency – dominated by T
c
(combustion chamber temperature).
2.3 INTERNAL BALLISTICS:
The parameters that govern the burning state and mass discharge rate of motors are called
internal ballistic properties; they include r, k, k, p, and the influences caused by pressures,
propellants ingredients, gas velocity, or acceleration.
K = A6 /A1
p- temperature sensitivity of the burning rate
k– temperature sensitivity of pressure
r - burning rate.
2.4 PERFORMANCE CONSIDERATIONSOF ROCKETS:
The analysis of performance is usually divided into two somewhat separate sets of
calculations:
The combustion process is the first part. It usually occurs in the combustion chamber at
essentially constant chamber pressure (isobaric) and the resulting gases follow Dalton's law. The
chemical reactions or the combustions occur very rapidly. The chamber volume is assumed to be
large enough and the residence time in the chamber long enough for attaining chemical
equilibrium in the chamber.
The nozzle gas expansion process constitutes the second set of calculations. The fully reacted,
equilibrated gas combustion products enter the nozzle and undergo an adiabatic expansion in the
ejects
t
*
m
A
C

c
P


33
nozzle. The entropy remains constant during a reversible (isentropic) nozzle expansion, but in
real nozzle flows it increases slightly.
The thrust, depends upon the pressure in the combustion chamber, the properties of the
propellant and the geometrical shape of the rocket. To obtain high thrust the molecular weight of
the propellants must be as low as possible.
Another factor which limits the thrust obtainable is the maximum allowable temperature as
well as the maximum temperature which can be produced by chemical reactions. At very high
temperatures, dissociation does not allow the whole of the heat energy to be converted into the
kinetic energy and the maximum obtainable temperatures are limited.
The effect of the characteristics of the propellant and the area-ratio of the nozzle and its shape
also play an important part in dictating the performance of the chemical rocket as it affects the
velocities obtainable as well as the drag on the rocket.
The optimum expansion condition is that when the static pressure at the exit of the rocket
nozzle is the same as the ambient pressure. The length of the diverging nozzle passage is
increased over that corresponding to the optimum expansion ratio (Pa =pe) ; this will result in
further expansion in the nozzle and pressure at exit of the nozzle will be less than ambient
pressure pa.
On the other hand, any length of the nozzle which is less than that corresponding to Pe=Pa
will result in lesser expansion and hence in reduced exhaust velocity and thrust will again be
reduced. So the condition Pe=pa gives the best expansion condition for a rocket nozzle with a
given threat diameter.

34
2.5 TYPES OF IGNITERS:
PYROTECHNIC IGNITERS
In industrial practice, pyrotechnic igniters are defined as igniters (other than pyrogen-type
igniters as defined further on) using solid explosives or energetic propellant-like chemical
formulations (usually small pellets of propellant which give a large burning surface and a short
burning time) as the heat-producing material.
This definition fits a wide variety of designs, known as bag and carbon igniters, powder can,
plastic case, pellet basket, perforated tube, combustible case, jellyroll, string, or sheet igniters.
The common pellet-basket design in figure below is typical of the pyrotechnic igniters. Ignition
of the main charge, in this case pellets consisting of 24% boron-71% potassium perchlorate-5%
binder, is accomplished by stages; first, on receipt of an electrical signal the initiator releases the
energy of a small amount of sensitive powdered pyrotechnic housed within the initiator,
commonly called the squib or the primer charge; next, the booster charge is ignited by heat
released from the squib; and finally, the main ignition charge propellants are ignited.

35
PYROGEN IGNITERS:
A pyrogen igniter is basically a small rocket motor that is used to ignite a larger rocket
motor. The pyrogen is not designed to produce thrust. All use one or more nozzle orifices, both
sonic and supersonic types, and most use conventional rocket motor grain formulations and
design technology. Heat transfer from the pyrogen to the motor grain is largely convective, with
the hot gases contacting the grain surface as contrasted to a highly radiative energy emitted by
pyrotechnic igniters. For pyrogen igniters the initiator and the booster charge are very similar to
the designs used in pyrotechnic igniters. Reaction products from the main charge impinge on the
surface of the rocket motor grain, producing motor ignition. Common practice on the very large
motors is to mount externally, with the pyrogen igniter pointing its jet up through the large motor
nozzle.
SAFETY DEVICES IN IGNITERS
Two approaches are commonly used to safeguard against motor misfires, or inadvertent
motor ignition; one is the use of the classical safe and arm device and the second is the design of
safeguards into the initiator. Energy for unintentional ignition--usually a disaster when it
happens--can be (1) static electricity, (2) induced current from electromagnetic radiation, such as
radar, (3) induced electrical currents from ground test equipment, communication apparatus, or
nearby electrical circuits in the flight vehicle, and (4) heat, vibration, or shock from handling and
operations. Functionally, the safe and arm device serves as an electrical switch to keep the igniter
circuit grounded when not operating; in some designs it also mechanically misaligns or blocks
the ignition train of events so that unwanted ignition is precluded even though the initiator fires.

36
When transposed into the arm position, the ignition flame can be reliably propagated to the
igniter's booster and main charges.
Electric initiators in motor igniters are also called squibs, glow plugs, primers, and sometimes
headers; they always constitute the initial element in the ignition train and, if properly designed,
can be a safeguard against unintended ignition of the motor.
Both (a) and (b) structurally form a part of the rocket motor case and generically are headers.
In the integral diaphragm type (a) the initial ignition energy is passed in the form of a shock
wave through the diaphragm activating the acceptor charge, with the diaphragm remaining
integral. This same principle is also used to transmit a shock wave through a metal case wall or a
metal insert in a filament-wound case; the case would not need to be penetrated and sealed. The
header type (b) resembles a simple glow plug with two high-resistance bridgewires buried in the
initiator charge. The exploding bridgewire design (c) employs a small bridgewire (0.02 to 0.10
mm) of low-resistance material, usually platinum or gold, that is exploded by application of a
high voltage discharge.

37
2.6 PRELIMINARY CONCEPTS IN NOZZLE-LESS PROPULSION:
The nozzleless solid rocket motor utilizes a solid-fuel grain geometrically contoured to
function as a rocket engine nozzle during operation.
The nozzleless solid rocket motor is in reality an advanced mass fraction system. By avoiding
the inherent weight penalty of a nozzle, the nozzleless motor potentially offers improved payload
or terminal velocity capability. The nozzleless motor could also be simple, low cost and reliable.
The attractiveness of nozzleless solid rocket motors is due not only to cost savings and
simplification but their increased potential for flight missions. Using the terminal velocity or V
capability of a rocket as a rmeasure of performance, nozzleless rockets may be compared to
conventional rockets. The ideal V equation is:
By simply removing the nozzle from the conventional motor and firing it, about 70 percent of
its standard total impulse is achieved. The most important reason is the propellant burn rate.
Conventional motors are generally designed for operation In the pressure regime of 1000 to 2000
psi. Burn rates and nozzle sizes are matched appropriately. Port areas are usually much larger
than the throat area in the motors to avoid erosive burning. When the nozzle is removed, they

38
port becomes the throat, thus reducing the chamber pressure and propellant burn rate. The lower
pressures result in lower total impulse. Decreasing the burn rate of the propellant without
changing other motor parameters can thus have a significant effect upon the delivered total
impulse. A motor designed for operation without a nozzle would use the volume originally taken
up by the nozzle for additional propellant; would have a higher web fraction; would use a
propellant having more desirable propellant properties (such as a low pressure exponent); and
would probably have a lighter case. All these changes would greatly enhance the comparison
between nozzleless and conventional motors in terms of performance.
Elimination of the nozzle from a conventional rocket motor will result in a 30 percent
decrease in total impulse. By placing additional propellant in the volume originally occupied by
the nozzle, utilizing improved grain design, and selecting proper propellant3, the decrease in
total impulse can be reduced to 10 percent or less. Preliminary studies indicate that nozzleless
motors would cost 20 to 30 percent less than the conventional motors, and the use of nozzleless
rocket motors appears to be attractive for small, unguided, air-launched solid rockets.
SHORTCOMINGS:
Considerable development is necessary to operate at low pressure and to avoid the instability
problem. It is assumed that proper propellant tailoring and grain design could accomplish stable
operation. In addition, the nozzleless motor designs thus far envisioned cannot accommodate the
traditional fins necessary for flight stability.
2.7 AIR AUGMENTED ROCKETS:
A ducted rocket, sometimes called as an air-augmented rocket, combines the principles of
rocket and ramjet engines, it give higher performance (specific impulse) than a chemical rocket
engine, while operating within the earth‟s atmosphere. Usually the term air-augmented rocket
denotes mixing of air with the rocket exhaust (full-rich for after burning) in proportions that
enabled the propulsion device to retain the characteristics of the rocket engine, for example, high
static thrust and high thrust to weight ratio. In contrast, the ducted rocket often is like a ramjet in
that it must be boosted to operating speed and uses the rocket components more as a fuel-rich
generator (liquid, solid, or hybrid), igniter, and air ejector pump).

39
The principles of the rocket and ramjet can be combined so that the two propulsion systems
operate in sequence and in tandem and yet utilize a common combustion chamber and volume.
The low-volume configuration, known as an integral rocket-ramjet, can be attractive in air
launched missiles using ramjet propulsion. The transition from the rocket to the ramjet requires
enlarging the exhaust nozzle throat (usually by ejecting rocket nozzle parts), opening the ramjet
air inlet-combustion chamber interface, and following these two events with the normal ramjet
starting sequence.
A solid fuel ramjet uses a grain of solid fuel that gasifies or ablates and reacts with air. Good
combustion efficiencies have been achieved with a patented boron-containing solid fuel
fabricated into a grain similar to a solid propellant and burning in a manner similar to a hybrid
rocket propulsion system.
2.8 PULSE ROCKET MOTORS:
A pulsed rocket motor is typically defined as a multiple Pulse (physics) solid-fuel rocket motor.
This design overcomes the limitation of solid propellant motors that they cannot be easily shut
down and reignited. The pulse rocket motor allows the motor to be burned in segments (or
pulses) that burn until completion of that segment. The next segment (or pulse) can be ignited on
command by either an onboard algorithm or in pre-planned phase. All of the segments are
contained in a single rocket motor case as opposed to staged rocket motors. The pulsed rocket
motor is made by pouring each segment of propellant separately. Between each segment is

40
a barrier that prevents the other segments from burning until ignited. At ignition of a second
pulse the burning of the propellant generally destroys the barrier.
The benefit of the pulse rocket motor is that by the command ignition of the subsequent pulses,
near optimal energy management of the propellant burn can be accomplished. Each pulse can
have different thrust level, burn time, and achieved specific impulse depending on the type of
propellant used, its burn rate, its grain design, and the current nozzle throat diameter.
2.9 STATIC TESTING OF ROCKETS & INSTRUMENTATION:
Static rocket system tests with complete propulsion system on test stand is carried out in two
methods:
(a) partial or simulated rocket operation (for proper function, calibration, ignition, operation--
often without establishing full thrust or operating for the full duration)
(b) complete propulsion system tests (under rated conditions, off-design conditions, with
intentional variations in environment or calibration).
For a reusable or restartable rocket propulsion system this can include many starts, long-duration
endurance tests, and post operational inspections and reconditioning.
These tests can be performed on at least three basic types of programs:
1. Research on and development or improvement of a new (or modified) rocket engine or motor
or their propellants or components.
2. Evaluation of the suitability of a new (or modified) rocket engine or motor for a specified
application or for flight readiness.
3. Production and quality assurance of a rocket propulsion system.
For chemical rocket propulsion systems, each test facility usually has the following major
systems or components:
1. A test cell or test bay where the article to be tested is mounted, usually in a special test fixture.
If the test is hazardous, the test facility must have provisions to protect operating personnel and
to limit damage in case of an accident.
2. An instrumentation system with associated computers for sensing, maintaining, measuring,
analyzing, correcting, and recording various physical and chemical parameters. It usually
includes calibration systems and timers to accurately synchronize the measurements.
3. A control system for starting, stopping, and changing the operating conditions.

41
4. Systems for handling heavy or awkward assemblies, supplying liquid propellant, and
providing maintenance, security, and safety.
5. For highly toxic propellants and toxic plume gases it has been required to capture the
hazardous gas or vapor (firing inside a closed duct system), remove almost all of the hazardous
ingredients (e.g., by wet scrubbing and/or chemical treatment), allow the release of the nontoxic
portion of the cleaned gases, and safely dispose of any toxic solid or liquid residues from the
chemical treatment. With an exhaust gas containing fluorine, for example, the removal of much
of this toxic gas can be achieved by scrubbing it with water that contains dissolved calcium; it
will then form calcium fluoride, which can be precipitated and removed.
6. In some tests specialized test equipment and unique facilities are needed to conduct static
testing under different environmental conditions or under simulated emergency conditions. For
example, high and low ambient temperature tests of large motors may require a temperature-
controlled enclosure around the motor; a rugged explosion-resistant facility is needed for bullet
impact tests of propellant-loaded missile systems and also for cook-off tests, where gasoline or
rocket fuel is burned with air below a stored missile. Similarly, special equipment is needed for
vibration testing, measuring thrust vector forces and moments in three dimensions, or
determining total impulse for very short pulse durations at low thrust.
INSTRUMENTATION AND DATA MANAGEMENT:
Some of the physical quantities measured in rocket testing are as follows:
1. Forces (thrust, thrust vector control side forces, short thrust pulses).
2. Flows (hot and cold gases, liquid fuel, liquid oxidizer, leakage).
3. Pressures (chamber, propellant, pump, tank, etc.).
4. Temperatures (chamber walls, propellant, structure, nozzle).
5. Timing and command sequencing of valves, switches, igniters, etc.
6. Stresses, strains, and vibrations (combustion chamber, structures, propellant lines,
accelerations of vibrating parts)
7. Time sequence of events (ignition, attainments of full pressure).
8. Movement and position of parts (valve stems, gimbal position, deflection of parts under load
or heat),
voltages, frequencies, and currents in electrical or control subsystems.

42
9.Visual observations (flame configuration, test article failures, explosions) using high-speed
cameras or video cameras.
10.Special quantities such as turbopump shaft speed, liquid levels in propellant tanks, burning
rates, flame luminosity, or exhaust gas composition.
MEASUREMENT SYSTEM TERMINOLOGY:
Range refers to the region extending from the minimum to the maximum rated value over which
the measurement system will give a true and linear response.
Errors in measurements are usually of two types: (1) human errors of improperly reading the
instrument, chart, or record and of improperly interpreting or correcting these data, and (2)
instrument or system errors, which usually fall into four classifications: static errors, dynamic
response errors, drift errors, and hysteresis errors.
Static errors are usually fixed errors due to fabrication and installation variations; these errors
can usually be detected by careful calibration, and an appropriate correction can then be applied
to the reading.
Drift error is the change in output over a period of time, usually caused by random wander and
environmental conditions.
Dynamic response errors occur when the measuring system fails to register the true value of the
measured quantity while this quantity is changing, particularly when it is changing rapidly.
A maximum frequency response refers to the maximum frequency (usually in cycles per
second) at which the instrument system will measure true values. The natural frequency of the
measuring system is usually above the limiting response frequency.
Linearity of the instrument refers to the ratio of the input (usually pressure, temperature, force,
etc.) to the output (usually voltage, output display change, etc.) over the range of the instrument.
Resolution refers to the minimum change in the measured quantity that can be detected with a
given instrument. Dead zone or hysteresis errors are often caused by energy absorption within
the instrument system or play in the instrument mechanism; in part, they limit the resolution of
the instrument.
Sensitivity refers to the change in response or reading caused by special influences. For example,
the temperature sensitivity and the acceleration sensitivity refer to the change in measured value
caused by temperature and acceleration.

43
2.10 SAFETY CONSIDERATIONS:
Most rocket propulsion testing is now accomplished in sophisticated facilities under closely
controlled conditions. Modern rocket test facilities are frequently located several miles from the
nearest community to prevent or minimize effects of excessive noise, vibrations, explosions, and
toxic exhaust clouds.
Prior to performing any test, it is common practice to train the test crew and go through repeated
dry runs, to familiarize each person with his or her responsibilities and procedures, including the
emergency procedures.
Typical personnel and plant security or safety provisions in a modern test facility include the
following:
1. Concrete-walled blockhouse or control stations for the protection of personnel and instruments
remote from the actual rocket propulsion location.
2. Remote control, indication, and recording of all hazardous operations and measurements;
isolation of propellants from the instrumentation and control room.

44
3. Automatic or manual water deluge and fire-extinguishing systems.
4. Closed circuit television systems for remotely viewing the test.
5. Warning signals (siren, bells, horns, lights, speakers) to notify personnel to clear the test area
prior to a test, and an all-clear signal when the conditions are no longer hazardous.
6. Quantity and distance restrictions on liquid propellant tankage and solid propellant storage to
minimize damage in the event of explosions; separation of liquid fuels and oxidizers.
7. Barricades around hazardous test articles to reduce shrapnel damage in the event of a blast.
8. Explosion-proof electrical systems, spark-proof shoes, and non spark hand tools to prevent
ignition of flammable materials.
9. For certain propellants also safety clothing (see Fig. 20-4), including propellant- and fire-
resistant suits, face masks and shields, gloves, special shoes, and hard hats.
10. Rigid enforcement of rules governing area access, smoking and safety inspections.
11. Limitations on the number of personnel that may be in a hazardous area at any time.
MONITORING AND CONTROL OF TOXIC MATERIALS:
Open-air testing of chemical rockets frequently requires measurement and control of exhaust
cloud concentrations and gas movement in the surrounding areas for safeguarding personnel,
animals, and plants. A toxic cloud of gas and particles can result from the exhaust gas of normal
rocket operation, vapors or reaction gases from unintentional propellant spills, and gases from
fires, explosions, or from the intentional destruction of vehicles in flight or rockets on the launch
stand. Environmental regulations usually limit the maximum local concentration or the total
quantity of toxic gas or particulates released to the atmosphere. One method of control is for tests
with discharges of moderately toxic gases or products to be postponed until favorable weather
conditions are present.

45
UNIT III
SOLID ROCKET PROPULSION
Salient features of solid propellant rockets – selection criteria of solid propellants- estimation of
solid propellant adiabatic flame temperature - propellant grain design considerations – erosive
burning in solid propellant rockets – combustion instability – strand burner and T-burner –
applications and advantages of solid propellant rockets.
3.1 SALIENT FEATURES OF SOLID PROPELLANT ROCKETS:
The principal components and features of relatively single solid propellant rocket motors.
The grain is the solid body of the hardened propellant and typically accounts for 82% to 94% of
the total motor mass. The igniter (electrically activated) provides the energy to start the
combustor. This grain configuration has a central cylinder cavity with light tapered slots,
forming an 8-pointed star. Many grains have slots, grooves, holes, or other geometric features
and they alter the initial burning surface, which determines the initial mass flow and the initial
thrust. The hot reaction gases flow along the perforation or port cavity towards the nozzle. The
inner surface of the case (really a pressure vessel), which are exposed directly to hot gas have a
thermal protection or insulation layer to keep the case from becoming too hot, in which case it
could no longer carry its pressure and other loads. The case is either made of metal(such as steel,
aluminum or titanium) or a composite fiber-reinforced plastic material.
The nozzle accelerates the hot gas; it is made of high temperature materials (usually graphite
and/or an ablative material to absorb the heat) to withstand the high temperature. The majority of

46
all solid rockets have a simple fixed nozzle. as shown here, but some nozzle have provision to
rotate it slightly so as to control the direction of the thrust to allow vehicle steering.
3.2 SELECTION CRITERIA OF SOLID PROPELLANTS:
The propellant selection is critical to rocket motor design. The desirable propellant
characteristics are listed below. The requirement for any particular motor will influence the
priorities of these characteristics:
1. High performance or high specific impulse; really this means high gas temperature and/or
low molecular mass.
2. Predictable, reproducible, and initially adjustable burning rate to fit the need of the grain
design and the thrust-time requirement.
3. For minimum variation in thrust or chamber pressure, the pressure or burning rate exponent
and the temperature coefficient should be small.
4. Adequate physical properties (including bond strength) over the intended operating
temperature range.
5. High density (allows a small-volume motor).
6. Predictable, reproducible ignition qualities (such as reasonable ignition overpressure)
7. Good aging characteristics and long life. Aging and life predictions depend on the
propellant‟s chemical and physical properties, the cumulative damage criteria with load
cycling and thermal cycling and actual tests on propellant samples and test date from failed
motors.
8. Low absorption of moisture, which often causes chemical deterioration.
9. Simple, reproducible, safe, low-cost, controllable, and low-hazard manufacturing.
10. Guaranteed availability of all raw materials and purchased components over the reduction
and operating life of the propellant, and good control over undesirable impurities.
11. Low technical risk, such as a favorable history of prior applications.
12. Relative insensitivity to certain energy stimuli
13. Non-toxic exhausts gases.
14. Not prone to combustion instability.

47
3.3 ESTIMATION OF SOLID PROPELLANT ADIABATIC FLAME TEMPERATURE:
All of these theoretical analyses are only approximations of what really happens in rocket
combustion and nozzle flow, and they all require some simplifying assumptions. The analysis is
usually divided into two somewhat separate sets of calculations:
The combustion process is the first part. It usually occurs in the combustion chamber at
essentially constant chamber pressure (isobaric) and the resulting gases follow Dalton's law. The
chemical reactions or the combustions occur very rapidly. The chamber volume is assumed to be
large enough and the residence time in the chamber long enough for attaining chemical
equilibrium in the chamber.
The nozzle gas expansion process constitutes the second set of calculations. The fully reacted,
equilibrated gas combustion products enter the nozzle and undergo an adiabatic expansion in the
nozzle. The entropy remains constant during a reversible (isentropic) nozzle expansion, but in
real nozzle flows it increases slightly.
The principal chemical reactions occur inside the combustion chamber of a liquid propellant
rocket engine or inside the grain cavity of a solid propellant rocket motor, usually within a short
distance from the burning surface.
Rocket propulsion systems usually do not operate with the proportion of their oxidizer and fuel
in the stoichiometric mixture ratio. Instead, they usually operate fuel-rich because this allows
lightweight molecules such as hydrogen to remain unreacted ; this reduces the average molecular
mass of the reaction products, which in turn increases the specific impulse.
Dalton's law applies to the gas resulting from the combustion. It states that a mixture of gases at
equilibrium exerts a pressure that is the sum of the partial pressures of the individual gases, all at
the same temperature. The subscripts a, b, c, etc. refer to individual gas constituents.
The perfect gas equation p V = RT applies very closely to high temperature gases. Here V is the
specific volume or the volume per unit mass of gas mixture, and the gas constant R for the
mixture is obtained by dividing the universal gas constant R' (8314.3 J/kg-mol-K) by the average
molecular mass (often erroneously called the molecular weight) of the gas mixture.

48
Using Dalton's law, the above equation can be written as,
The effective average molecular mass / of a gas mixture is given by
In most rocket propulsion the heat of reaction is determined for a constant-pressure combustion
process. In general the heat of reaction can be determined from sums of the heats of formation of
the products and the reactants, namely
Here nj is the molar fraction of each particular species j.
The free energy is a function of temperature and pressure. It is another property of a material,
just like enthalpy or density; only two such independent parameters are required to characterize a
gas condition. The free energy may be thought of as the tendency or driving force for a chemical
material to enter into a chemical (or physical) change.
The chemical reaction occurs instantaneously but isothermally at the reference temperature,
and the resulting energy release then heats the gases from this reference temperature to the final
combustion temperature. The heat of reaction is
Here h is the increase in enthalpy for each species multiplied by its molar fraction, and Cp is
the molar specific heat at constant pressure.
Once the gases reach the nozzle, they experience an adiabatic, reversible expansion process
which is accompanied by a drop in temperature and pressure and a conversion of thermal energy
into kinetic energy.

49
3.4 PROPELLANT GRAIN DESIGN CONSIDERATIONS:
The grain is the solid body of the hardened propellant and typically accounts for 82 to 94% of
the total motor mass. Design and stresses of grains are described later in this chapter. Propellants
are described in the next chapter. The igniter (electrically activated) provides the energy to start
the combustion. The grain starts to burn on its exposed inner surfaces. This grain configuration
has a central cylindrical cavity with eight tapered slots, forming an 8-pointed star. Many grains
have slots, grooves, holes, or other geometric features and they alter the initial burning surface,
which determines the initial mass flow and the initial thrust. The hot reaction gases flow along
the perforation or port cavity toward the nozzle.
The inner surfaces of the case (really a pressure vessel), which are exposed directly to hot gas,
have a thermal protection or insulation layer to keep the case from becoming too hot, in which
case it could no longer carry its pressure and other loads. The case is either made of metal (such
as steel, aluminum or titanium) or a composite fiber-reinforced plastic material.
The nozzle accelerates the hot gas; it is made of high temperature materials (usually a graphite
and/or an ablative material to absorb the heat) to withstand the high temperatures and the
erosion. The majority of all solid rockets have a simple fixed nozzle, as shown here, but some
nozzles have provision to rotate it slightly so as to control the direction of the thrust to allow
vehicle steering.
PROPELLANT BURNING RATE:
The burning surface of a propellant grain recedes in a direction essentially perpendicular
to the surface. The rate of regression usually exposed in un/sec, mm/sec, or in/sec, is the burning
rate.
Success in rocket motor design and development depends significantly on knowledge of
burning rate behavior of the selected propellant under all motor operating conditions and design
limit conditions. Burning rate is a function of the propellant composition, for composite
propellant, it can be increased by changing the propellant characteristics.
1. Add a burning rate catalyst, often called burning rate modifier (0.1 to 3.0% of
propellant) or increase the percentage of existing catalyst
2. Decrease the oxidizer percentage.
3. Increase oxidizer percentage.

50
4. Increase the heat of combustion of the binder and (or the plasticizer).
5. Imbed wires or metal staples in the propellant.
A side from the propellant formulation and propellant manufacturing process, burning rate in a
full scale motor can be increased by the following
i. Combustion chamber pressure.
ii. Initial temperature of the solid propellant prior to start.
iii. Combustion gas temperature.
iv. Velocity of the gas flow parallel to the burning surface.
v. Motor motion (acceleration and spin-induced grain stress)
Burning rate date are usually obtained in three ways namely, from testing by:
i. Standard strand burner, often called craw ford burners
ii. Small scale ballistic evaluation motors.
iii. Full scale motors with good instrumentation.
The burning rate of propellant in the motor is a function of many parameters, and at any instant
governs to the mass flow rate m of hot gas generated and flowing form the motor (stable
combustion)
m=Ab X r X pb
Here, Abis the burning of the propellant grain; r is the burning rate, and pb is the solid propellant
density prior to motor start. The total mass “m” of the effective propellant burned can be
determined by integration the equation.
m=ʃ m dt = pb Ar X r dt
Where Ab and r vary with the time and pressure.
BURNING RATE RELATION WITH PRESSURE:
Unless otherwise stated, burning rate for most propellant is expressed for 70*F or 294K
propellant (prior to ignition) burning at a reference chamber pressure of 1000psi or 6.894Mpa.
With many propellants it is possible to approximate the burning rate as a function of chamber
pressure, at least over a limited range of chamber pressures. For most production type
propellants, the empirical equation for burning rate is
r=a Po
n

51
Where r - the burn rate is usually in centimeters per second and the chamber pressure „a‟ is an
empirical constant influenced by ambient grain temperature. This equation applies to all the
commonly used double – base, composite, or composite double-base propellants. Also „a‟ is
known as the temperature co-efficient and it is not dimensionless. The burning rate exponent „n‟
sometimes called the combustion index, is independent of the in initial grain temperature and
describes the influence of chamber pressure on the burning rate. The change in ambient
temperature does not change the chemical energy released in the combustion; it merely changes
the rate of reaction at which energy is released.
For a particular propellant and for wide temperature and pressure limits, the burning rate can
vary by factor of 3 or 4. For all propellants, they range form about 0.05to 75mm/sec; The high
values are difficult to achieve, even with considerable burning rate catalyst additives, embedded
metal wires, or high pressures (above 14Mpa or 2000Mpa)
The burning rate very sensitive to the exponent n for stable operation, n values greater then O
and less than, I, High values of n give a rapid change of burning rate with pressure. This implies
that even a small change in chamber pressure produces substantial changes in the amount of hot
gas produced. Most production propellants have a pressure exponent n ranging between 0.2 and
0.6. In practice, as „n‟ approaches 1, burning rate and chamber pressure become very sensitive to
one another and disastrous rise in chamber pressure can occur in a few milliseconds. When the
„n‟ value is low and comes closed to zero, burning can become unstable and may even
extinguish itself. Some propellants display a negative „n‟ which is very important for „restorable‟
motors or gas generators. A propellant having a wide pressure range. Plate an propellants are
those that exhibit a nearly constant burning rate over a limited pressure range.
BURNING RATE RELATION WITH TEMPERATURE:
Temperature affects chemical reaction rates and the initial ambient temperature of a propellant
grains prior to combustion influences burning rate. The motor performance characteristics must
stay within specified acceptable limits. For air launched missile motors, the extremes are usually
219K and 344K. Motors using typical composite propellants experience a 20 to 35% variation in
chamber pressure and a 20 to 30% variation in operating time over such a range of propellant
temperatures. In large rocket motors, an uneven melting of the grain can cause a sufficiently
large difference in burning rate so that a slight thrust misalignment can be produced.

52
The sensitivity of burning rate to propellant temperature can be expressed in the form of
temperature co-efficient,
[ ] [ ]
[ ] [ ]
Where is known as temperature sensitivity of burning rate, expressed as percent change of
burning rate per degree change in propellant temperature at a particular value of chamber
pressure, and as the temperature sensitivity of pressure expressed as percent change of
chamber pressure per degree change in propellant temperature at a particular value of k. Hence k
is the geometric function, namely the ratio of the burning surface, Ab to nozzle throat area Ak

53
CLASSIFICATION OF SOLID PROPELLANT ROCKET MOTOR:
Processed modern propellants can be classified in several ways, as described below. This
classification is not rigorous or complete. Sometimes the same propellant will fit into two or
more of the classification.
Propellants are often tailored to and classified by specific applications such as space launch
booster propellants or tactical missile propellants; each has somewhat specific chemical
ingredients, different burning rates, different physical properties, and different performance.
Table shows four kinds of rocket motor applications (each has somewhat different propellants)
and several gas generator applications. Propellants for rocket motors have hot (over 2400k) gases
and are used to produce thrust, but gas generator propellants have lower-temperature combustion
gases (800 to 1200k) and they are used to produce power not thrust. Historically, the early rocket
motor propellants used to be grouped into two classes: double-base (DB*) propellants were used
as the first production propellants, and then the development of polymers as binders made the
composite propellants feasible.
Double-base (DB) propellants for a homogeneous propellant grain, usually a nitrocellulose
(NC*), a solid ingredient which absorbs liquid nitroglycerine (NG) plus minor percentages of
additives. Both the major ingredients are explosive and function as a combined fuel and oxidizer.
Both extruded double-base (EDB) and cast double-case (CDB) propellant have found extensive
applications, mostly in small tactical missiles of older design. By adding crystalline nitramines
(HMX or RDX)* the performance and density can be improved; this is sometimes called cast-
modified double-base propellant. A further improvement is to add an elastomeric binder (rubber-
like, such as cross linked poly-butadiene) which improves the physical properties and allows
more nitramine and thus improves the performance slightly. The resulting propellant is called
elastomeric-modified double-base (EMCDB). These four classes of double base have nearly
smokeless exhausts. Adding some solid ammonium per chlorate (AP) and aluminum (A1)
increase the density and the specific impulse slightly, but the exhaust gas is smoky. The
propellant is called composite-modified double-base propellant or CMDB.
Composite propellants form a heterogeneous propellant grain with the oxidizer crystals and
powered fuel (usually aluminum) held together in a matrix of synthetic rubber (or plastic) binder,
such as poly butadiene (HTPB)*. Composite propellants are cast from a mix of solid (AP
crystals, Al powder)* and liquid (HTPB, PPG)* ingredients. The propellant is hardened by cross

54
linking or curing the liquid binder polymer with a small amount of curing agent, and curing it in
an oven, where it becomes hard and solid. In the past three decades the composite propellants
have been the most commonly used class. They have further subdivided below.
1. Conventional composite propellants usually contain between 60 and 72% ammonium per
chlorate (AP) as crystalline oxidizer, up to 22% aluminum powder (AI) as a metal fuel, and 8 to
16% of elastomeric binder (organic polymer) including its plasticizer.
2. Modified composite propellant where an energetic nitramine (HMX or RDX) is added for
obtaining a little more performance and also a somewhat higher density.
3. Modified composite propellant where an energetic plasticizer such as nitroglycerine (used in
double-base propellant) is added to give a little more performance. Sometimes HMX is also
added.
4. A high energy composite solid propellant (with some aluminum), where the organic
elastomeric binder and plasticizer are largely replaced by energetic materials (such as certain
explosives) and where some of the AP is replaced by HMX. Some of these are called elastomer-
modifier cast double-base propellants (EMCDB). Most are experimental propellants. He
theoretical specific impulse can be between 270 and 275 sec at standard conditions.
5. A lower energy composite propellant, where ammonium nitrate (AN) is the crystalline
oxidizer (no AP). It is used for gas generator propellant. If a large amount of HMX is added, it
can become a minimum smoke propellant with fair performance.
Propellants can be classified by the density of the smoke in the exhaust plume as smoky
reduced smoke, or minimum smoke (essentially smoke-less). Aluminium powder, a desirable
fuel ingredient, is oxidized to aluminium oxide, which forms visible small solid smoke particles
in the exhaust gas. Most composite propellant are smoky. By reducing the aluminum content in-
composite propellant, the amount of smoke is also reduced. Carbon (Soot) partials and metal
oxide, such as zirconium oxide or iron oxide, can also be visible if in high enough concentration.
The safety rating for detonation can distinguish propellants as a potentially detonable
material (class 1.1) or as a non-detonable materials (class 1.3). Examples of class 1.1 propellant
are a double-base propellants and composite propellants containing a significant proportion of
solid explosive (e.g., HMX or RDX), together with certain other ingredients.
Propellants can be classified by some of the principal manufacturing processes that are
used. Cast propellant is made by mechanical mixing of solid and liquid ingredients, followed by

55
casting and curing; it is the most common process for composite propellants. Curing of many
cast propellants is by chemical reaction between binder and curing agent at elevated temperature
(45 to 150 C) or hardened by a non chemical process such crystallization. Propellant can also be
made by a salvation process (dissolving a plasticizer in a solid palletized matrix, whose volume
is expanded ). Extruded propellant is made by mechanical mixing (rolling into sheets) followed
by extrusion (pushing through a die at high pressure). Salvation and extrusion process apply
primarily to double-base propellants.
Propellants have also been classified by their principal ingredient. Such as the principal
oxidizer (ammonium per chlorate propellants, ammonium nitrate propellants. or azide-type
propellants) or their principal binder or fuel ingredient and also by propellants with toxic and
nontoxic exhaust gases.
PROPELLANT GRAIN AND GRAIN CONFIGURATION:
The grain is the shaped mass of processed solid propellant inside the rocket motor. The
propellant material and geometrical configuration of the grain determine the motor performance
characteristics. The propellant grain is a cast, molded, or extruded body and its appearance and
feel is similar to that of hard rubber or plastic. Once ignited, it will burn on all its exposed
surfaces to form hot gases that are then exhausted through a nozzle. A few rocket motors have
more than one grain inside a single case or chamber and very few grains have segments made of
different propellant composition
Cartridge-loaded or freestanding grains are manufactured separately from the case (by
extrusion or by casting into a cylindrical mold or cartridge) and then loaded into or assembled
into the case. In case-bonded grains the case is used as a mold and the propellant is cast directly
into the case and is bonded to the case or case insulation.

56
Free-standing grains can more easily be replaced if the propellant grain has aged excessively.
Cartridge-loaded grains are used in some small tactical missiles and a few medium-sized motors.
They often have a lower cost and are easier to inspect. The case-bonded grains give a somewhat
better performance, a little less inert mass (no holding device, support pads, and less insulation),
a better volumetric loading fraction, are more highly stressed, and often somewhat more difficult
and expensive to manufacture. Today almost all larger motors and many tactical missile motors
use case bonding.
Definitions and terminology important to grains include:
Configuration: The shape or geometry of the initial burning surfaces of a grain as it is
intended to operate in a motor.
Cylindrical Grain: A grain in which the internal cross section is constant along the axis
regardless of perforation shape.
Neutral Burning:Motor burn time during which thrust, pressure, and burning surface area
remain approximately constant, typically within about =15%. Many grains are neutral burning.
Perforation: The central cavity port or flow passage of a propellant grain; its cross section
may be a cylinder, a star shape, etc.
Progressive Burning: Burn time during which thrust, pressure, and burning surface area
increases
Regressive Burning: Burn time during which thrust, pressure, and burning surface area
decreases
Silver: Unburned propellant remaining (or lost – that is, expelled through the nozzle) at the
time of web burnout.

PROPULSION II

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