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					          AE-1351

         PROPULSION-II


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              UNIT-1


         AIRCRAFT GAS TURBINES




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Gas Turbine
• 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
  power:

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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
  accessories.



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                TYPES OF TURBINES




• AXIAL FLOW TURBINE
• RADIAL FLOW TURBINE




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             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
or wheel.
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        RADIAL FLOW TURBINE




 The radial flow turbine is similar in design and construction to the
centrifugal-flow compressor .
Advantages
-ruggedness and simplicity
-relatively inexpensive and easy to manufacture when compared to
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the axial-flow turbine.
              BLADE TYPES
• IMPULSE TYPE
• REACTION TYPE




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Single-rotor, Single-stage Turbine          Multiple-rotor, Multiple-stage Turbine




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                             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.

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  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
  in hours.
• 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.


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   Difference between impulse and
           reaction turbine




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           UNIT-II
RAMJET PROPULSION




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• 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
  helicopter rotors.
• 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


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         RAMJET ENGINE PARTS




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         RAMJET ENGINE THRUST




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  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




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                                              INLETS
•   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.




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                 INLET EFFICIENCY
•   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.


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                                     SCRAMJET
•   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
    Mach 30+.
•   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


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         Axial turbine blade cooling




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           UNIT-III
FUNDAMENTALS OF ROCKET
 PROPULSION




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          INTERNAL BALLISTICS
•   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.




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         NOZZLES




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         ROCKET NOZZLE
         CLASSIFICATION
• FIXED NOZZLE
• MOVABLE NOZZLE
• SUBMERGED NOZZLE




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           UNIT-IV
CHEMICAL ROCKETS




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         ADVANTAGES OF SOLID
            ROCKET MOTOR
• Advantages/Disadvantages
• 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
  ingredient.

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         SOLID ROCKET MOTOR




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• 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

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•   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
    the thrust:
•   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




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• 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

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                                 Liquid propellant
•   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,
    aka SS-1-c,-d,-e
•   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



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           UNIT-V
ADVANTAGES OF PROPULSION
 TECHNIQUES




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• 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.




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• 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.
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•   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.

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•   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
    thruster.
•
•   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.
                                                   •




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         Nuclear rocket motor




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• 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
  create thrust.
• The nuclear reactor's energy replaces the chemical
  energy of the reactive chemicals in a traditional rocket
  engine.
• 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.

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      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.



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         Solar sail




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• 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.



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             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




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         THANK U




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