AIRCRAFT GAS TURBINES
• In an aircraft gas turbine the output of the
turbine is used to turn the compressor
(which may also have an associated fan or
propeller). The hot air flow leaving the
turbine is than accelerated into the
atmosphere through an exhaust nozzle
(Fig. la) to provide thrust or propulsion
FUNCTION OF TURBINE
A portion of the kinetic energy of the
expanding gases is extracted by the
turbine section, and this energy is
transformed into shaft horsepower which
is used to drive the compressor and
TYPES OF TURBINES
• AXIAL FLOW TURBINE
• RADIAL FLOW TURBINE
AXIAL FLOW TURBINE
It consists of two main elements
A set of stationary vanes followed by a turbine rotor.
Axial-flow turbines may be of the single-rotor or multiple-rotor type.
A stage consists of two main components: a turbine nozzle and a turbine rotor
RADIAL FLOW TURBINE
The radial flow turbine is similar in design and construction to the
centrifugal-flow compressor .
-ruggedness and simplicity
-relatively inexpensive and easy to manufacture when compared to
the axial-flow turbine.
• IMPULSE TYPE
• REACTION TYPE
Single-rotor, Single-stage Turbine Multiple-rotor, Multiple-stage Turbine
Multirotor - Multistage Turbine.
TURBINE BLADE COOLING
• Current turbine inlet temperatures in gas turbine engines are beyond
the melting point of the turbine blade material.
• To prevent the blades from melting, turbine blade cooling methods are
applied to the first turbine stages. Since
• convectively cooled flow fields and temperature fields are coupled and
interact strongly, it is necessary to understand
• the flow physics in order to accurately predict how cooling will
behave. In case of the rotating turbine blades, the effects
• of rotation also influence the flow. Centrifugal forces as well as the
Coriolis force have to be included in the analysis.
• The present research project is an experimental and computational
investigation of the flow through internal turbine
• blade cooling passages. In the first phase, the flow in a straight,
stationary cooling channel is observed. Pressure
• measurements as well as hot-wire and PIV measurements are used to
determine efficacy of different turbulator
• geometries. In the phase 2, the flow in a rotating cooling channel with
an 180° bend will be investigated using PIV.
Advantages of the gas turbine
• It is capable of producing large amounts of useful power for a
relatively small size and weight
• Since motion of all its major components involve pure rotation (i.e.
no reciprocating motion as in a piston engine), its mechanical life is
long and the corresponding maintenance cost is relatively low.
• Although the gas turbine must be started by some external means (a
small external motor or other source, such as another gas turbine), it
can be brought up to full-load (peak output) conditions in minutes as
contrasted to a steam turbine plant whose start up time is measured
• A wide variety of fuels can be utilized. Natural gas is commonly used
in land-based gas turbines while light distillate (kerosene-like) oils
power aircraft gas turbines. Diesel oil or specially treated residual
oils can also be used, as well as combustible gases derived from
blast furnaces, refineries and the gasification of solid fuels such as
coal, wood chips and bagasse.
Difference between impulse and
• A ramjet, sometimes referred to as a stovepipe jet, or an athodyd,
is a form of jet engine that contains no major moving parts. Unlike
most other airbreathing jet engines, ramjets have no rotary
compressor at the inlet, instead, the forward motion of the engine
itself 'rams' the air through the engine. Ramjets therefore require
forward motion through the air to produce thrust.
• Ramjets require considerable forward speed to operate well, and as
a class work most efficiently at speeds around Mach 3, and this type
of jet can operate up to speeds of at least Mach 5.
• Ramjets can be particularly useful in applications requiring a small
and simple engine for high speed use; such as missiles. They have
also been used successfully, though not efficiently, as tip jets on
• Ramjets are frequently confused with pulsejets, which use an
intermittent combustion, but ramjets employ a continuous
combustion process, and are a quite distinct type of jet engine
RAMJET ENGINE PARTS
RAMJET ENGINE THRUST
Axial turbine blade vortex balding
• A turbine blade for a turbine engine having one or more cavities in a
trailing edge of the turbine blade for forming one or more vortices in
inner aspects of the trailing edge.
• In at least one embodiment, the turbine blade may include one or
more elongated cavities in the trailing edge of the blade formed by
one or more ribs placed in a cooling chamber of the turbine blade.
• The elongated cavity in the trailing edge may have one or more
orifices in the rib on the upstream side of the cavity. The orifice may
be positioned relative to a vortex forming surface so that as a gas is
passed through one or more orifices into the elongated cavity, one
or more vortices are formed in the cavity. The gas may be expelled
from the cavity and the blade through one or more orifices in an
inner wall forming the pressure side of the turbine blade
• SUBSONIC INLETS
• For aircraft that cannot go faster than the speed of sound, like large airliners, a simple, straight, short inlet works
quite well. On a typical subsonic inlet, the surface of the inlet from outside to inside is a continuous smooth curve
with some thickness from inside to outside. The most upstream portion of the inlet is called the highlight, or the
inlet lip. A subsonic aircraft has an inlet with a relatively thick lip.
• SUPERSONIC INLETS
• An inlet for a supersonic aircraft, on the other hand, has a relatively sharp lip. The inlet lip is sharpened to
minimize the performance losses from shock waves that occur during supersonic flight. For a supersonic aircraft,
the inlet must slow the flow down to subsonic speeds before the air reaches the compressor. Some supersonic
inlets, like the one at the upper right, use a central cone to shock the flow down to subsonic speeds. Other inlets,
like the one shown at the lower left, use flat hinged plates to generate the compression shocks, with the resulting
inlet geometry having a rectangular cross section. This variable geometry inlet is used on the F-14 and F-15
fighter aircraft. More exotic inlet shapes are used on some aircraft for a variety of reasons. The inlets of the Mach
3+ SR-71 aircraft are specially designed to allow cruising flight at high speed. The inlets of the SR-71 actually
produce thrust during flight.
• HYPERSONIC INLETS
• Inlets for hypersonic aircraft present the ultimate design challenge. For ramjet-powered aircraft, the inlet must
bring the high speed external flow down to subsonic conditions in the burner. High stagnation temperatures are
present in this speed regime and variable geometry may not be an option for the inlet designer because of
possible flow leaks through the hinges. For scramjet-powered aircraft, the heat environment is even worse
because the flight Mach number is higher than that for a ramjet-powered aircraft. Scramjet inlets are highly
integrated with the fuselage of the aircraft. On the X-43A, the inlet includes the entire lower surface of the aircraft
forward of the cowl lip. Thick, hot boundary layers are usually present on the compression surfaces of hypersonic
inlets. The flow exiting a scramjet inlet must remain supersonic.
• An inlet must operate efficiently over the entire flight envelope of the aircraft. At very
low aircraft speeds, or when just sitting on the runway, free stream air is pulled into
the engine by the compressor. In England, inlets are called intakes, which is a more
accurate description of their function at low aircraft speeds. At high speeds, a good
inlet will allow the aircraft to maneuver to high angles of attack and sideslip without
disrupting flow to the compressor. Because the inlet is so important to overall aircraft
operation, it is usually designed and tested by the airframe company, not the engine
manufacturer. But because inlet operation is so important to engine performance, all
engine manufacturers also employ inlet aerodynamicists. The amount of disruption of
the flow is characterized by a numerical inlet distortion index. Different airframes
use different indices, but all of the indices are based on ratios of the local variation of
pressure to the average pressure at the compressor face.
• The ratio of the average total pressure at the compressor face to the free stream total
pressure is called the total pressure recovery. Pressure recovery is another inlet
performance index; the higher the value, the better the inlet. For hypersonic inlets the
value of pressure recovery is very low and nearly constant because of shock losses,
so hypersonic inlets are normally characterized by their kinetic energy efficiency. If
the airflow demanded by the engine is much less than the airflow that can be
captured by the inlet, then the difference in airflow is spilled around the inlet. The
airflow mis-match can produce spillage drag on the aircraft.
• A scramjet (supersonic combustion ramjet) is a variation of a ramjet with the distinction being that some or all of
the combustion process takes place supersonically. At higher speeds, it is necessary to combust supersonically to
maximize the efficiency of the combustion process. Projections for the top speed of a scramjet engine (without
additional oxidiser input) vary between Mach 12 and Mach 24 (orbital velocity). The X-30 research gave Mach 17
due to combustion rate issues. By way of contrast, the fastest conventional air-breathing, manned vehicles, such
as the U.S. Air Force SR-71, achieve approximately Mach 3.4 and rockets from the Apollo Program achieved
• Like a ramjet, a scramjet essentially consists of a constricted tube through which inlet air is compressed by the
high speed of the vehicle, a combustion chamber where fuel is combusted, and a nozzle through which the
exhaust jet leaves at higher speed than the inlet air. Also like a ramjet, there are few or no moving parts. In
particular, there is no high-speed turbine, as in a turbofan or turbojet engine, that is expensive to produce and can
be a major point of failure.
• A scramjet requires supersonic airflow through the engine, thus, similar to a ramjet, scramjets have a minimum
functional speed. This speed is uncertain due to the low number of working scramjets, relative youth of the field,
and the largely classified nature of research using complete scramjet engines. However, it is likely to be at least
Mach 5 for a pure scramjet, with higher Mach numbers (between 7 and 9) more likely. Thus scramjets require
acceleration to hypersonic speed via other means. A hybrid ramjet/scramjet would have a lower minimum
functional Mach number, and some sources indicate the NASA X-43A research vehicle is a hybrid design. Recent
tests of prototypes have used a booster rocket to obtain the necessary velocity. Air breathing engines should have
significantly better specific impulse while within the atmosphere than rocket engines.
• However, scramjets have weight and complexity issues that must be considered. While very short suborbital
scramjet test flights have been successfully performed, perhaps significantly no flown scramjet has ever been
successfully designed to survive a flight test. The viability of scramjet vehicles is hotly contested in aerospace and
space vehicle circles, in part because many of the parameters which would eventually define the efficiency of such
a vehicle remain uncertain. This has led to grandiose claims from both sides, which have been intensified by the
large amount of funding involved in any hypersonic testing. Some notable aerospace gurus such as Henry
Spencer and Jim Oberg have gone so far as calling orbital scramjets 'the hardest way to reach orbit', or even
'scamjets' due to the extreme technical challenges involved. Major, well funded projects, like the X-30 were
cancelled before producing any working hardware
Axial turbine blade cooling
FUNDAMENTALS OF ROCKET
• Internal ballistics is what happens inside a weapon when it is fired. The
firing pin makes a distinct mark on the cartridge. Then explosive pressure
causes the bullet to expand slightly to fill the spiral 'rifling' grooves cut in the
bore. This makes the bullet spin as it passes down the barrel, but it leaves
tell-tale marks on the bullet that are unique to that particular firearm. The
presence of rust or spider silk indicates the gun has not been fired recently.
At close range, particles from a wound may lodge inside the barrel.
• External ballistics is what happens to the bullet and residues outside the
gun, including the direction and velocity of the shot, as well as any deviation
in the trajectory.
• Terminal ballistics looks at the changes in trajectory and speed caused by
ricochet and penetration of objects, as well as the layered deposits on parts
of the bullet accumulated as it contacts these objects. Terminal ballistics
includes examination of the shape of wounds and the extent of tissue
damage. If a bullet cannot be removed for examination, its calibre can be
measured by CT scanning.
• FIXED NOZZLE
• MOVABLE NOZZLE
• SUBMERGED NOZZLE
ADVANTAGES OF SOLID
• Solid fueled rockets are relatively simple rockets. This is
their chief advantage, but it also has its drawbacks.
• Once a solid rocket is ignited it will consume the entirety
of its fuel, without any option for shutoff or thrust
adjustment. The Saturn V moon rocket used nearly 8
million pounds of thrust that would not have been
feasible with the use of solid propellant, requiring a high
specific impulse liquid propellant.
• The danger involved in the premixed fuels of
monopropellant rockets i.e. sometimes nitroglycerin is an
SOLID ROCKET MOTOR
• A simple solid rocket motor consists of a casing, nozzle, grain
(propellant charge), and igniter.
• The grain behaves like a solid mass, burning in a predictable
fashion and producing exhaust gases. The nozzle dimensions are
calculated to maintain a design chamber pressure, while producing
thrust from the exhaust gases.
• Once ignited, a simple solid rocket motor cannot be shut off,
because it contains all the ingredients necessary for combustion
within the chamber that they are burned in. More advanced solid
rocket motors can not only be throttled but can be extinguished and
then re-ignited by controlling the nozzle geometry or through the use
of vent ports. Also, pulsed rocket motors which burn in segments
and which can be ignited upon command are available.
• Modern designs may also include a steerable nozzle for guidance,
avionics, recovery hardware (parachutes), self-destruct
mechanisms, APUs, controllable tactical motors, controllable divert
and attitude control motors and thermal management materials
• Thrust is the force which moves a rocket through the air. Thrust is generated by the rocket engine through the reaction of accelerating a mass of gas. The gas is
accelerated to the the rear and the rocket is accelerated in the opposite direction. To accelerate the gas, we need some kind of propulsion system. We will discuss the
details of various propulsion systems on some other pages. For right now, let us just think of the propulsion system as some machine which accelerates a gas.
• From Newton's second law of motion, we can define a force to be the change in momentum of an object with a change in time. Momentum is the object's mass times the
velocity. When dealing with a gas, the basic thrust equation is given as:
• F = mdot e * Ve - mdot 0 * V0 + (pe - p0) * Ae
• Thrust F is equal to the exit mass flow rate mdot e times the exit velocity Ve minus the free stream mass flow rate mdot 0 times the free stream velocity V0 plus the
pressure difference across the engine pe - p0 times the engine area Ae.
• For liquid or solid rocket engines, the propellants, fuel and oxidizer, are carried on board. There is no free stream air brought into the propulsion system, so the thrust
equation simplifies to:
• F = mdot * Ve + (pe - p0) * Ae
• where we have dropped the exit designation on the mass flow rate.
• Using algebra, let us divide by mdot:
• F / modt = Ve + (pe - p0) * Ae / mdot
• We define a new velocity called the equivalent velocity Veq to be the velocity on the right hand side of the above equation:
• Veq = Ve + (pe - p0) * Ae / mdot
• Then the rocket thrust equation becomes:
• F = mdot * Veq
• The total impulse (I) of a rocket is defined as the average thrust times the total time of firing. On the slide we show the total time as "delta t". (delta is the Greek symbol
that looks like a triangle):
• I = F * delta t
• Since the thrust may change with time, we can also define an integral equation for the total impulse. Using the symbol (Sdt) for the integral, we have:
• I = S F dt
• Substituting the equation for thrust given above:
• I = S (mdot * Veq) dt
• Remember that mdot is the mass flow rate; it is the amount of exhaust mass per time that comes out of the rocket. Assuming the equivalent velocity remains constant
with time, we can integrate the equation to get:
• I = m * Veq
• where m is the total mass of the propellant. We can divide this equation by the weight of the propellants to define the specific impulse. The word "specific" just means
"divided by weight". The specific impulse Isp is given by:
• Isp = Veq / g0
• where g0 is the gravitational acceleration constant (32.2 ft/sec^2 in English units, 9.8 m/sec^2 in metric units). Now, if we substitute for the equivalent velocity in terms of
• Isp = F / (mdot * g0)
• Mathematically, the Isp is a ratio of the thrust produced to the weight flow of the propellants. A quick check of the units for Isp shows that:
• Isp = m/sec / m/sec^2 = sec
• A bipropellant rocket engine is a rocket engine that
uses two propellants (very often liquid propellants) which
are kept separately prior to reacting to form a hot gas to
be used for propulsion.
• In contrast, most solid rockets have single solid
propellant, and hybrid rockets use a solid propellant
lining the combustion chamber that reacts with an
injected fluid. Because liquid bipropellant systems permit
precise mixture control, they are often more efficient than
solid or hybrid rockets, but are normally more complex
and expensive, particularly when turbopumps are used
to pump the propellants into the chamber to save weight
• Thousands of combinations of fuels and oxidizers have been tried over the years. Some of the more common and practical ones are:
• liquid oxygen (LOX, O2) and liquid hydrogen (LH2, H2) - Space Shuttle main engines, Ariane 5 main stage and the Ariane 5 ECA second
stage, the first stage of the Delta IV, the upper stages of the Saturn V, Saturn IB, and Saturn I as well as Centaur rocket stage
• liquid oxygen (LOX) and kerosene or RP-1 - Saturn V, Zenit rocket, R-7 Semyorka family of Soviet boosters which includes Soyuz, Delta,
Saturn I, and Saturn IB first stages, Titan I and Atlas rockets
• liquid oxygen (LOX) and alcohol (ethanol, C2H5OH) - early liquid fueled rockets, like German (World War II) A-4, aka V-2, and Redstone
• liquid oxygen (LOX) and gasoline - Robert Goddard's first liquid-fuel rocket
• T-Stoff (80% hydrogen peroxide, H2O2 as the oxidizer) and C-Stoff (methanol, CH3OH, and hydrazine hydrate, N2H4•n(H2O as the fuel)
- Walter Werke HWK 109-509 engine used on Messerschmitt Me 163B Komet a rocket fighterplane of (World War II)
• nitric acid (HNO3) and kerosene - Soviet Scud-A, aka SS-1
• inhibited red fuming nitric acid (IRFNA, HNO3 + N2O4) and unsymmetric dimethyl hydrazine (UDMH, (CH3)2N2H2) Soviet Scud-B,-C,-D,
• nitric acid 73% with dinitrogen tetroxide 27% (=AK27) and kerosene/gasoline mixture - various Russian (USSR) cold-war ballistic
missiles, Iran: Shahab-5, North Korea: Taepodong-2
• hydrogen peroxide and kerosene - UK (1970s) Black Arrow, USA Development (or study):
• hydrazine (N2H4) and red fuming nitric acid - Nike Ajax Antiaircraft Rocket
• Aerozine 50 and dinitrogen tetroxide - Titans 2–4, Apollo lunar module, Apollo service module, interplanatary probes (Such as Voyager 1
and Voyager 2)
• Unsymmetric dimethylhydrazine (UDMH) and dinitrogen tetroxide - Proton rocket and various Soviet rockets
• monomethylhydrazine (MMH, (CH3)HN2H2) and dinitrogen tetroxide - Space Shuttle Orbital maneuvering system (OMS) engines
• Robert Goddard and his rocket
• One of the most efficient mixtures, oxygen and hydrogen, suffers from the extremely low temperatures required for storing hydrogen and
oxygen as liquids (around 20 K or −253 °C)) and low fuel density (70 kg/m³), necessitating large and heavy tanks. The use of lightweight
foam to insulate the cryogenic tanks caused problems for the Space Shuttle Columbia's STS-107 mission, as a piece broke loose,
damaged its wing and caused it to break up and be destroyed on reentry.
• For storable ICBMs and interplanetary spacecraft, storing cryogenic propellants over extended periods is awkward and expensive.
Because of this, mixtures of hydrazine and its derivatives in combination with nitrogen oxides are generally used for such rockets.
Hydrazine has its own disadvantages, being a very caustic and volatile chemical as well as being toxic. Consequently, hybrid rockets
have recently been the vehicle of choice for low-budget private and academic developments in aerospace technology
ADVANTAGES OF PROPULSION
• Chemical rocket engines, like those on the space shuttle, work by
burning two gases to create heat, which causes the gases to expand
and exit the engine through a nozzle. In so doing they create the
thrust that lifts the shuttle into orbit. Smaller chemical engines are
used to change orbits or to keep satellites in a particular orbit.
• For getting to very distant parts of the solar system chemical
engines have the drawback in that it takes an enormous amount of
fuel to deliver the payload. Consider the Saturn V rocket that put
men on the moon: 5,000,000 pounds of it's total take off weight of
6,000,000 pounds was fuel.
• Electric rocket engines use less fuel than chemical engines and
therefore hold the potential for accomplishing missions that are
impossible for chemical systems. To understand how, we have to
understand a number called specific impulse.
• A resistojet simply uses electricity passing through a
resistive conductor, something like the wires in your
toaster, to heat a gas as it passes over the conductor. As
the conductor heats up the gas is heated, expands, exits
through a nozzle and creates thrust.
diagram of a resistojet
• In real resistojets the conductor is a coiled tube through
which the propellant flows. This is done to get maximum
heat transfer from the conductor to the propellant.
• An arcjet is simply a resistojet where instead of passing the gas through a
heating coil it's passed through an electric arc.
diagram of an arcjet
• Because arcs can achieve temperatures of 15,00 degrees C. this means the
propellant gets heated to much higher temperatures (typically 3,000
degrees C.) than in resistojets and in so doing achieve higher specific
impulses, anywhere from 800 sec for ammonia to 2,000 seconds for
hydrogen. Arcjets tend to be higher power devices, typically 1 to 2 kilowatts,
and used for higher thrust applications, like station keeping of large
satellites. Several are currently in orbit.
• Ion engines:
• Rub a balloon against your hair or shirt and then hold it near your arm, the hairs on your arm will
feel tingly and be attracted to the balloon. Bring the balloon near the carpet and bits of lint will be
pulled to it. What's happening is that electrons have been deposited onto or removed from the
balloon depending on what it was rubbed against, giving it an electrostatic charge, which creates
an electrostatic field. A similar field can be used to produce thrust in a rocket engine called an ion
• ion engine diagram
• As propellant enters the ionization chamber (the small ns on the left), electrons (small -s in the
middle) emitted from the central hot cathode and attracted to the outer anode collide with them
knocking an electron off and causing the atoms of the propellant to become ionized (+s on the
right). This means that they have an electric field around them like the balloon. As these ions drift
between two screens at the right hand side of the ionization chamber, the strong electric field of
the "+" side repels them and the "-" side attracts them, accelerating them to very high velocities.
The ions leave the engine and since the engine pushes on them to accelerate them, they in turn
push back against the engine creating thrust. Ion thrusters typically use Xenon (A very heavy,
inert gas) for propellant, have specific impulses in the 3,000 to 6,000 range and efficiencies up to
60 percent. An average thruster is one to two feet in diameter, produces thrust on the order of
small fractions of a pound and weighs some tens of pounds.
Nuclear rocket motor
• In a nuclear thermal rocket a working fluid, usually
hydrogen, is heated to a high temperature in a nuclear
reactor, and then expands through a rocket nozzle to
• The nuclear reactor's energy replaces the chemical
energy of the reactive chemicals in a traditional rocket
• Due to the higher energy density of the nuclear fuel
compared to chemical ones, about 107 times, the
resulting efficiency of the engine is at least twice as good
as chemical engines even considering the weight of the
reactor, and even higher for advanced designs.
Risk in nuclear rocket motor
• There is an inherent possibility of atmospheric or orbital rocket failure which
could result in a dispersal of radioactive material, and resulting fallout.
Catastrophic failure, meaning the release of radioactive material into the
environment, would be the result of a containment breach. A containment
breach could be the result of an impact with orbital debris, material failure
due to uncontrolled fission, material imperfections or fatigue and human
design flaws. A release of radioactive material while in flight could disperse
radioactive debris over the Earth in a wide and unpredictable area. The
zone of contamination and its concentration would be dependent on
prevailing weather conditions and orbital parameters at the time of re-entry.
However given that oxide reactor elements are designed to withstand high
temperatures (up to 3500 K) and high pressures (up to 200 atm normal
operating pressures) it's highly unlikely a reactor's fuel elements would be
reduced to powder and spread over a wide-area. More likely highly
radioactive fuel elements would be dispersed intact over a much smaller
area, and although individually quite lethal up-close, the overall hazard from
the elements would be confined to near the launch site and would be much
lower than the many open-air nuclear weapons tests of the 1950s.
• The spacecraft arranges a large membrane mirror which
reflects light from the Sun or some other source.
• The radiation pressure on the mirror provides a small
amount of thrust by reflecting photons.
• Tilting the reflective sail at an angle from the Sun
produces thrust at an angle normal to the sail.
• In most designs, steering would be done with auxiliary
vanes, acting as small solar sails to change the attitude
of the large solar sail.
• The vanes would be adjusted by electric motors.
Limitation of solar sail
• Solar sails don't work well, if at all, in low Earth orbit below about
800 km altitude due to erosion or air drag. Above that altitude they
give very small accelerations that take months to build up to useful
speeds. Solar sails have to be physically large, and payload size is
often small. Deploying solar sails is also highly challenging to date.
• Solar sails must face the sun to decelerate. Therefore, on trips away
from the sun, they must arrange to loop behind the outer planet, and
decelerate into the sunlight.
• There is a common misunderstanding that solar sails cannot go
towards their light source. This is false. In particular, sails can go
toward the sun by thrusting against their orbital motion. This reduces
the energy of their orbit, spiraling the sail toward the sun