SYLLABUS
AE6504Propulsion II
UNIT I HYPERSONIC AIRBREATHING PROPULSION 8
Introduction to hypersonic air breathing propulsion, hypersonic vehicles and supersonic
combustion
Hypersonic airbreathing engines are the engine which is operated above Mach 5. 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.
Need for supersonic combustion for hypersonic propulsion
Supersonic combustion is needed in order to produce a hypersonic speed from the
combustion chamber through the nozzle.
Salient features of scramjet engine and its applications for hypersonic vehicles - problems
associated with supersonic combustion
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. This is a
modified Brayton cycle because the final state in the scramjet nozzle is generally not
ambient. Engine specific impulse, or the efficiency of airbreathing ramjet, scramjet and
turbine engines, compared to the rocket is illustrated Figure 1. Specific impulse is the thrust
(Nts) produced per unit mass flow (Kg/s) of propellant utilized, i.e. propellant which is
carried on board.
Application of supersonic & hypersonic vehicles
Military application
Commercial transport
Cruise missiles
Reentry vehicles
engine/airframe integration aspects of hypersonic vehicles
propulsion system and the airframe have to be optimized together. This slide shows the impact of
integrating an airbreathing engine on the lower fuselage of a typical configuration designed for
high speed. The forebody shape is used as a precompression ramp of the engine intake and the
afterbody is used as an additional 2D expansion ramp.
Various types scramjet combustors
Usually made of several cans or combustors. It burns the fuel under difficult conditions as
changing RPM, altitude, air pressure, temperature and various TAS. There are three main types:
can, annular and can-annular. Of all the airflow entering the can only about 20 % is used to burn
the fuel, another 20 % enters the can through holes in the side of the can and the rest is used for
cooling after combustion is completed.
Air temperature from the compressor runs between 200 and 400 °C and goes further up to
600 to 1650 °C in the cans. Flame temperatures are some 2200 °C.
Can
combustors
are used with radial flow compressors as they can be positioned around the compressor
turbine axis and the air flows from the diffuser direct into the can. Annular combustors
are suited for the axial flow compressor. Fuel enters the cans upstream and air is used to
center the flame without it touching the can liner. Secondary air is used to cool the cans.
General
Electric
Aviation
uses
a double
annular
combustor
(DAC)
for reduced CO2 and nitro oxide (NOX) emissions. Now lets wait and see who will build the
first triple annular combustor.
Can-annular
combustors
are a combination of individually placed side by side cans to form a circle inside an annular
chamber. Thus combining the best features of both types in one. Providing an even
temperature distribution along the HP turbine without hot spots should a fuel nozzle become
blocked.
Fuel injection schemes in scramjet combustors - one dimensional models for supersonic
combustion using method of influence coefficients.
Ramjets can be thought of as propulsive devices evolved out of turbojets.
Ramjets operates well only at high speeds, typically between M = 2.0 and 4.0.
The combustion mode being not very different from that of an afterburner, the specific fuel
consumption is comparable to that of afterburner.
The ramjet, unlike turbojet and turbofan does not produce any thrust at zero speed.
A ramjet has no moving parts and achieves compression of intake air by the forward speed of
the air vehicle. Air entering the intake of a supersonic aircraft is slowed by aerodynamic
diffusion created by the inlet and diffuser to velocities comparable to those in a turbojet
augmenter. The expansion of hot gases after fuel injection and combustion accelerates the
exhaust air to a velocity higher than that at the inlet and creates positive push. Ramjets are
mostly contemplated for use in military applications.
UNIT II FUNDAMENTALS OF CHEMICAL ROCKET PROPULSION 9
Operating principle specific impulse of a rocket- internal ballistics performance
considerations of rockets
In a solid propellant rocket, the propellant is entirely stored within the combustion chamber in
the form of one or several shaped blocks which are called grains and are supported by the walls
or by special grids, traps or retainer walls. A typical solid rocket motor is illustrated below.
All the conventional propulsion systems work by causing a change of momentum in a
working fluid in a direction opposite to the intended motion. Rockets fall under the category
of direct acting engines – since the energy liberated by the chemical process is directly used
to obtain thrust. Being non-air breathing devices the basic component of a rocket are
(i) Combustion chamber where exothermic processes produces gases at high temperature and
pressure, and
(ii) Nozzle, which accelerate the fluid to high velocities and discharge them into surrounding
atmosphere thereby deriving the desired force or thrust.
The reaction system does not depend on the surrounding atmosphere. There are no
velocity limitations and altitude ceiling.
(b) Since it has to carry its own oxidizer required for combustion reaction, the specific
propellant consumption is very high. Rockets consume approximately 15kg/kg-hr of
propellant compared to about 1 kg/kg-hr of fuel by turbojet engine.
(c) High pressure operation is possible and hence the ratio of energy liberation per unit
volume (and also unit weight of hardware) is very high.
(d) Main part of the rockets contains no moving element. Hence there is no constraint on
internal aerodynamics and the reliability is high. This also implies quick response times,
which makes them ideal control components.
Internal ballistics Performance consideration:
 Mass ratio
 Specific impulse
 Thrust
 Area ratio
 Burning rate
 Propellant considered
Types of igniters
Pyrotechnic ignitor
Round spiral
Silicon-nitrate ignitor
A pyrotechnic initiator (also initiator or igniter) is a device containing a pyrotechnic
composition used primarily to ignite other, more difficult-to-ignite materials, e.g. thermites, gas
generators, and solid-fuel rockets. The name is often used also for the compositions themselves.
Pyrotechnic initiators are often controlled electrically (called electro-pyrotechnic initiators),
e.g. using a heated bridgewire or a bridge resistor. They are somewhat similar to blasting caps or
other detonators, but they differ in that there is no intention to produce a shock wave. An
example of such pyrotechnic initiator is an electric match.
Preliminary concepts in nozzle-less propulsion
Propulsion by means of uniform cross sectional ducts rather than a varying section (nozzle) is
said to be nozzleless propulsion. In nozzleless propulsion the mass burning rate of the propellant
in the duct is increased and the flow reaches sonic velocity the increased pressure in the duct is
converted in to thrust.
Air augmented rockets
Air-augmented rockets (also known as rocket-ejector, ramrocket, ducted rocket, integral
rocket/ramjets, or ejector ramjets) use the supersonic exhaust of some kind of rocket engine to
further compress air collected by ram effect during flight to use as additional working mass,
leading to greater effective thrust for any given amount of fuel than either the rocket or a ramjet
alone.
It represents a hybrid class of rocket/ramjet engines, similar to a ramjet, but able to give useful
thrust from zero speed, and is also able in some cases to operate outside the atmosphere, with
fuel efficiency not worse than both a comparable ramjet or rocket at every point.
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
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
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
For a combustion process that takes place adiabatically with no shaft work, the temperature of
the products is referred to as the adiabatic flame temperature. This is the maximum temperature
that can be achieved for given reactants. Heat transfer, incomplete combustion, and dissociation
all result in lower temperature. The maximum adiabatic flame temperature for a given fuel and
oxidizer combination occurs with a stoichiometric mixture (correct proportions such that all fuel
and all oxidizer are consumed). The amount of excess air can be tailored as part of the design to
control the adiabatic flame temperature. The considerable distance between present temperatures
in a gas turbine engine and the maximum adiabatic flame temperature at stoichiometric
conditions is shown in Figure 3.24(b), based on a compressor exit temperature of
(922
K).
Figure 15.3: Schematic of adiabatic flame temperature
An initial view of the concept of adiabatic flame temperature is provided by examining two
reacting gases, at a given pressure, and asking what the end temperature is. The process is shown
schematically in Figure 15.3, where temperature is plotted versus the percentage completion of
the reaction. The initial state is and the final state is
, with the final state at a higher
temperature than the initial state. The solid line in the figure shows a representation of the
``actual'' process.
To see how we would arrive at the final completion state the dashed lines break the state of
reaction change into two parts. Process (1) is reaction at constant
and
. To carry out such
a process, we would need to extract heat. Suppose the total amount of heat extracted per unit
mass is
. The relation between the enthalpy changes in Process (1) is
where
is the ``heat of reaction.''
Propellant grain design considerations
A solid fuel rocket can contain several stages of its own composed of several shapes. Each
shape is carefully studied as to its thrust output, from rise to fall. The sequencing of these
shapes are important if the flight path itself contains phases. Each stage, or shape, will be
ignited by the flame from the previous shape. Inter-stage, or inter-shape, points are where
shut-off mechanisms can be installed, but like steerage mechanisms, shut-off devices can be
unneeded
erosive burning in solid propellant rockets
combustion instability
strand burner and T-burner
UNIT IV LIQUID AND HYBRID ROCKET PROPULSION 10
Salient features of liquid propellant rockets selection of liquid propellants
Various feed systems and injectors for liquid propellant rockets –
Pressure feed system
Pump feed system
Gas generator feed system
A liquid-propellant rocket or liquid rocket is a rocket engine that uses liquid propellants.
Liquids are desirable because their reasonably high density allows the volume of the propellant
tanks to be relatively low, and it is possible to use lightweight centrifugal turbopumps to pump
the propellant from the tanks into the combustion chamber, which means that the propellants can
be kept under low pressure. This permits the use of low-mass propellant tanks, resulting in a
high mass ratio for the rocket.
An inert gas stored in a tank at a high pressure is sometimes used instead of pumps in simpler
small engines to force the propellants into the combustion chamber. These engines may have a
lower mass ratio, but are usually more reliable.[1] :186,187 and are therefore used widely in
satellites for orbit maintenance.
Liquid rockets can be monopropellant rockets using a single type of propellant, bipropellant
rockets using two types of propellant, or more exotic tripropellant rocket using three types of
propellant. Some designs are throttleable for variable thrust operation and some may be restarted
after a previous in-space shutdown. Liquid propellants are also used in hybrid rockets, in which a
liquid oxidizer is generally combined with a solid fuel.
Pressure fed engines have simple plumbing and lack complex and often unreliable turbopumps.
A typical startup procedure begins with opening a valve, often a one-shot pyrotechnic device, to
allow the pressurizing gas to flow through check valves into the propellant tanks. Then the
propellant valves in the engine itself are opened. If the fuel and oxidizer are hypergolic, they
burn on contact; non-hypergolic fuels require an igniter. Multiple burns can be conducted by
merely opening and closing the propellant valves as needed. They can be operated electrically, or
by gas pressure controlled by smaller electrically operated valves.
thrust control and cooling in liquid propellant rockets and the associated heat transfer
problems combustion instability in liquid propellant rockets
Thrust vectoring, also thrust vector control or TVC, is the ability of an aircraft, rocket, or
other vehicle to manipulate the direction of the thrust from its engine(s) or motor in order
to control the attitude or angular velocity of the vehicle.
In rocketry and ballistic missiles that fly outside the atmosphere, aerodynamic control
surfaces are ineffective, so thrust vectoring is the primary means of attitude control.
For aircraft, the method was originally envisaged to provide upward vertical thrust as a means to
give aircraft vertical (VTOL) or short (STOL) takeoff and landing ability. Subsequently, it was
realized that using vectored thrust in combat situations enabled aircraft to perform various
maneuvers not available to conventional-engined planes. To perform turns, aircraft that use no
thrust vectoring must rely on aerodynamic control surfaces only, such as ailerons or elevator;
craft with vectoring must still use control surfaces, but to a lesser extent.
In missile literature originating from Russian sources,[1] thrust vectoring is often referred as "gasdynamic steering" or "gas-dynamic control".
peculiar problems associated with operation of cryogenic engines
The thrust comes from the rapid expansion from liquid to gas with the gas emerging from the
motor at very high speed. The energy needed to heat the fuels comes from burning them, once
they are gasses. Cryogenic engines are the highest performing rocket motors. Cryogenic engines
are fundamentally different from electric motors because there isn't anything rotating in them.
They're essentially reaction engines. By 'reaction' I'm referring to Newton's law: "to every action
there is an equal and opposite reaction."
The cryogenic (or rocket) engine throws mass in one direction, and the reaction to this is a thrust
in the opposite direction. Therefore, to get the required mass flow rate, the only option was to
cool the propellants down to cryogenic temperatures (below −183 °C [90 K], −253 °C [20 K]),
converting them to liquid form. Hence, all cryogenic rocket engines are also, by definition,
either liquid-propellant rocket engines or hybrid rocket engines
Introduction to hybrid rocket propulsion-standard and reverse hybrid systemscombustion mechanism in hybrid propellant rockets applications and limitations
UNIT V ADVANCED PROPULSION TECHNIQUES 8
Electric rocket propulsion– types of electric propulsion techniques
Ion propulsion
Ion propulsion is a technology that involves ionizing a gas to propel a craft. Instead of a
spacecraft being propelled with standard chemicals, the gas xenon (which is like neon or helium,
but heavier) is given an electrical charge, or ionized. It is then electrically accelerated to a speed
of about 30 km/second. Plasma is the building block for all types of electric propulsion, where
electric and/or magnetic fields are used to push on the electrically charged ions and electrons to
provide thrust. Ion thrusters are categorized by how they accelerate the ions, using
either electrostatic orelectromagnetic force. Electrostatic ion thrusters use the Coulomb force and
accelerate the ions in the direction of the electric field. Electromagnetic ion thrusters use
the Lorentz force to accelerate the ions. In either case, when an ion passes through an
electrostatic grid engine, the potential difference of the electric field converts to the ion's kinetic
energy.
Nuclear rocket
A nuclear thermal rocket is a proposed spacecraft propulsion technology. In a nuclear thermal
rocket a working fluid, usually liquid hydrogen, is heated to a high temperature in a nuclear
reactor, and then expands through a rocket nozzle to create thrust. In this kind of thermal rocket,
the nuclear reactor's energy replaces the chemical energy of the propellant's reactive chemicals in
a chemical rocket. The thermal heater / inert propellant paradigm as opposed to the reactive
propellants of chemical rockets turns out to produce a superioreffective exhaust velocity, and
therefore a superior propulsive efficiency, with specific impulses on the order of twice that of
chemical engines. The overall gross lift-off mass of a nuclear rocket is about half that of a
chemical rocket, and hence when used as an upper stage it roughly doubles or triples the payload
carried to orbit.
Types:
Solid core
Liquid core
Gas core
Solar sail.
Solar sails will use the sun's power to send us further into deep space. Solar sails (also
called light sails or photon sails) are a form of spacecraft propulsion using the radiation
pressure(also called solar pressure) from stars to push large ultra-thin mirrors to high speeds.
Light sails could also be driven by energy beams to extend their range of operations, which is
strictly beam sailing rather than solar sailing.
Solar sail craft offer the possibility of low-cost operations combined with long operating
lifetimes. Since they have few moving parts and use no propellant, they can potentially be used
numerous times for delivery of payloads.
Solar sails use a phenomenon that has a proven, measured effect on spacecraft. Solar pressure
affects all spacecraft, whether in interplanetary space or in orbit around a planet or small body. A
typical spacecraft going to Mars, for example, will be displaced by thousands of kilometres by
solar pressure, so the effects must be accounted for in trajectory planning, which has been done
since the time of the earliest interplanetary spacecraft of the 1960s. Solar pressure also affects
the attitude of a craft, a factor that must be included in spacecraft design.[1]
The total force exerted on an 800 by 800 meter solar sail, for example, is about 5 Newton’s
(1.1 lbf) at Earth's distance from the Sun, making it a low-thrust propulsion system, similar to
spacecraft propelled by electric engines.