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AE2255 propulsion

AE2255 propulsion

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    propulsion 1 propulsion 1 Presentation Transcript

    • RAJLAKSHMI ENGINEERING COLLEGE DEPARTMENT OF AERONAUTICAL ENGINEERING AE-2255 PROPULSION-I Prepared by MR.SURESH CHANDRA KHANDAI
    • UNIT-1 FUNDAMENTALS OF GAS TURBINE ENGINES
    • INTRODUCTION • Comprehend the thermodynamic processes occurring in a gas turbine. • Comprehend the basic components of gas turbine engines and their basic operation. • Comprehend the support systems associated with gas turbine engines.
    • ADVANTAGES OF GTE’s – – – – – – – – Weight reduction of 70% Simplicity Reduced manning requirements Quicker response time Faster Acceleration/deceleration Modular replacement Less vibrations More economical
    • DISADVANTAGES OF GTE’s • • • • • Many parts under high stress High pitched noise Needs large quantities of air Large quantities of hot exhaust (target) Cannot be repaired in place
    • TYPES OF JET ENGINE • • • • • • • TURBOJET ENGINE TURBOSHAFT ENGINE TURBOPROP ENGINE TURBOFAN ENGINE RAMJET ENGINE SCRAMJET ENGINE PULSEJET ENGINE
    • SCRAMJET ENGINE
    • BRAYTON CYCLE Unlike diesels, operate on STEADY-FLOW cycle Open cycle, unheated engine 1-2: Compression 2-3: Combustion 3-4: Expansion through Turbine and Exhaust Nozzle (4-1: Atmospheric Pressure)
    • BASIC COMPONENTS
    • NUMBERING OF TURBINE ENGINES intake compressor burner turbine afterburner (AB) nozzle (n) (diffuser) (e.g., turbofan)
    • COMPRESSOR • Supplies high pressure air for combustion process • Compressor types – Radial/centrifugal flow compressor – Axial flow compressor
    • COMPRESSOR • Radial/centrifugal flow – Adv: simple design, good for low compression ratios (5:1) – Disadv: Difficult to stage, less efficient • Axial flow – Good for high compression ratios (20:1) – Most commonly used
    • THE THRUST EQUATION
    • FACTORS AFFECTING THRUST • • • • • • PRESSURE TEMPERATURE DENSITY HUMIDITY ALTITUDE FORWARD VELOCITY
    • METHODS OF THRUST AUGMENTATION • • • AFTER BURNING INJECTION OF WATER & ALCOHOL MIXTURE BLEED BURN CYCLE
    • UNIT-II SUBSONIC & SUPERSONIC INLETS FOR JET ENGINES
    • INTRODUCTION Inlets are very important to the overall jet engine performance & will greatly influence jet engine thrust output. The faster the airplane goes the more critical the inlet duct design becomes. Engine thrust will be high only if the inlet duct supplies the engine with the required airflow at the highest possible pressure.
    • The nacelle/duct must allow the engine to operate with minimum stall/surge tendencies & permit wide variation in angle of attack & yaw of the aircraft. For subsonic aircraft, the nacelle shouldn’t produce strong shock waves or flow separations & should be of minimum weight for both subsonic & supersonic designs. For certain military applications, the radar cross sectional control or radar reflectance is a crucial design requirements.
    • Inlet ducts add to parasite drag skin friction+ viscous drag) & interference drag. It must operate from static ground run up to high aircraft Mach number with high duct efficiency at all altitude, attitudes & flight speeds. It should be as straight & smooth as possible & designed such a way that Boundary layer to be minimum. It should deliver pressure distribution evenly to the compressor.
    • Spring loaded , blow-in or such-in-doors are sometimes placed around the side of the inlet to provide enough air to the engine at high engine rpm & low aircraft speed. It must be shaped such a way that ram velocity is slowly & smoothly decreases while the ram pressure is slowly & smoothly increases.
    • DUCT EFFICIENCY The duct pressure efficiency ratio is defined as the ability of the duct to convert the kinetic or dynamic pressure energy at the inlet of the duct to the static pressure energy at the inlet of the compressor without a loss in total pressure . It is in order of 98% if there is less friction loss.
    • RAM RECOVERY POINT The ram recovery point is that aircraft speed at which the ram pressure rise is equal to the friction pressure losses or that aircraft speed at which the compressor inlet total pressure is equal to the outside ambient air pressure. A good subsonic duct has 257.4 km/h.
    • SINGLE ENTRANCE DUCT
    • SUBSONIC DUCTS
    • VARIABLE GEOMETRY DUCT FOR SUPERSONIC A/C
    • NORMAL SHOCK RELATION
    • OBLIQUE SHOCK RELATIONS
    • BOUNDARY LAYER
    • UNIT-III COMBUSTION CHAMBERS
    • COMBUSTION Combustion in normal, open cycle, gas turbine is a continuous process in which fuel is burned in the air supplied by the compressor, an electric spark is required only for initiating the combustion process and thereafter the flame must be selfsustaining.
    • CHAMBER GEOMETRY Over the years combustion chamber geometry has evolved considerably with respect to the objective of: 1.Improving flame stability both at sea altitudes. level and 2.Reducing chamber size (while still burning all the fuel and maintaining reasonable pressure drop) 3.Reducing emissions of the oxides of nitrogen, carbon monoxide and unburned hydrocarbons. 4. Increasing chamber life. 5. Controlling the temperature distribution at inlet to the turbine.
    • TYPES OF COMBUSTION CHAMBER • CAN TYPE • CAN-ANNULAR TYPE • ANNULAR TYPE
    • CAN TYPE • The earliest aircraft engines made use of Can (or tubular) combustor in which the air leaving the compressor splits into a number of separate streams, each supplying a separate chamber. • These chambers are spread around the shaft connecting the compressor & turbine, each chamber having its own fuel jet fed from a common supply line. • This arrangement was well suited to engines with centrifugal compressors, where the flow was divided into separate streams in the diffuser. Example Rolls-Royce Dart
    • ADVANTAGES • Development could be carried out on a single can using only a fraction of the overall airflow and fuel flow. • Ease of control of the fuel air ratio and simplicity. • Low cost of replacement of a damaged liner.
    • DISADVANTAGES • Relatively large size & weight of the chamber components. • Relatively large pressure drop & need to provide numerous igniters. • It is not easy to provide nearly simultaneous ignition of all chambers.
    • CONCLUSION • The disadvantages have out weighted the advantages & tubular chambers are now seldom used for large gas turbines for aircraft applications.
    • CAN-ANNULAR
    • ADVANTAGES • Ease ignition. • Minimum total cross-sectional area. • Minimum pressure drop. • Minimum length & weight.
    • DISADVANTAGES • Difficulties in development to obtain circumferentially uniform fuel-air ratio & outlet temperature. • A failure of the liner in one spot means replacement of relatively expensive component. • There tend to be a heavy buckling load, due to thermal expansion on the outer surface of the chamber liner. • Requires large airflow rates during testing.
    • ANNULAR
    • ADVANTAGES • Compact dimensions, in which maximum use is made of the space available within the specified diameter , this should reduce the pressure loss & results in an engine of minimum diameter.
    • DISADVANTAGES • It is more difficult to obtain an even fuel/air distribution & an even outlet temperature distribution. • Structurally weaker & it is difficult to avoid buckling of the hot flame tube walls. • For testing full air mass flow is required.
    • CONCLUSION • Annular combustors are universally used in modern aircraft engines. Some of the examples are, The Olympus 593. PT-6, PW 530, V2500
    • FACTOS AFFECTING COMBUSTOR DESIGN 1. The temperature of the gas after combustion must be comparatively low to suit the highly stressed turbine materials. Development of improved materials & method of blade cooling, enabled permissible combustor outlet temperatures to raise from about 1100K to as much as 1850K for aircraft applications.
    • 2. At the end of the combustion space the temperature distribution must be of known form if the turbine blades are not suffer from local overheating. In practice, the temperature can increase with radius over the turbine annulus, because of the strong influence of temperature on allowable stress and the decrease of blade centrifugal stress from toot to tip. 3.Combustion must be maintained in a stream of air moving with a high velocity in the region of 30-60m/s, and stable operation is required over a wide range of air/fuel ratio from full load to idling conditions.
    • 4. The formation of carbon deposits must be avoided, particularly the hard brittle variety. Small particles carried into the turbine in the high velocity gas stream can erode the blades and block cooling air passages. Aerodynamically excited vibration the combustion chamber might cause sizeable pieces of carbon to break free resulting in even worse damage to the turbine. 5. In aircraft gas turbines, combustion must also be stable over a wide range of chamber pressure because of the substantial change in this parameter with altitude &forward speed. It also capable of relighting at high altitude in the event of an engine flame out.
    • 6.Avoidance of smoke in the exhaust is of major importance of all types of gas turbine. Smoke trails in flight were a problem for military aircraft, permitting them to be seen from a great distance. 7. Although gas turbine combustion systems operate at extremely high efficiencies, they produce pollutants such as oxides of nitrogen, carbon monoxides and unburned hydrocarbons and these must be controlled to very low levels.
    • ZONES OF THE COMBUSTOR
    • • The function of recirculation zone is to evaporate, partly burn, and prepare the fuel for rapid combustion within the remainder of the burning zone. • The function of the burning zone to burn all the fuels. • The function of dilution zone is solely to mix the hot gas with the dilution air.
    • PERFORMANCE PARAMETERS The main factors of importance in assessing combustion chamber performance are •Pressure loss •Combustion efficiency •Outlet temperature distribution •Stability limits •Combustion intensity
    • Overall stagnation Pressure loss can be regarded as the sum of the fundamental loss( a small component which is a function of T02/T01) & the frictional loss. Pressure loss factor, PLF =
    • FLAME STABILIZATION 1.With the help of swirl vanes surrounding the fuel nozzle, strong vortex flow occurs in the combustion air in the combustion region. 2.A low pressure region is created at the combustor axis, which causes recirculation of the flame toward the fuel nozzle. 3.At the same time, radial holes around the liner supply air to the center of the vortex, making the flame grow to some extent. 4.Jet angles and penetration from the holes are such that jet impingement along the combustor axis results in upstream flow. 5.The upstream flow forms a torroidal recirculation zone,
    • Flame stabilization created by impinging jets and general airflow pattern.
    • POLLUTION CONTROL 1.Use of rich primary zone in which little NO formed, followed by rapid dilution in the secondary zone. 2.Use of a very lean primary zone to minimize peak flame temperature by dilution. 3.Use of water or stream admitted with the fuel for cooling the small zone downstream from the fuel nozzle. 4.Use of inert exhaust gas recirculated into the reaction zone. 5.Catalytic exhaust cleanup.
    • BASIC REQUIREMENTS OF FUELS 1.Heating value 2.Cleanliness 3.Corrosivity 4.Deposition and fouling tendencies 5.Availability
    • SUPERSONIC COMBUSTION The losses associated with subsonic ramjet combustion can be substantial. If ramjets are applied to hypersonic flight, additional problems arise because of extremely high temperature at the entrance of the combustion chamber. This not only makes vehicle cooling very difficult, but it leads to severe combustion loss due to dissociation. For hypersonic flight (M>8) the temperature of the air in the chamber is quite dependent on the pressure. The higher the pressure, the less dissociation and the higher the temperature of the mixture. The temperature of the combustion products is likewise pressure dependent.
    • For combustion pressure of 10atm and a flight Mach number of 10, there is no temperature rise due to combustion. All of the combustion energy is absorbed by dissociation. At sufficiently high Mach number, the temperature of the combustion products can be lower than that of the incoming air. Consideration with the speed with which the fuel and air can be converted into dissociation products may show that there is sufficient residence time in the combustion chamber to approach equilibrium composition.
    • During the subsequent expansion in the nozzle, it is quite possible that the expansion will be too rapid for the composition to readjust, after each step of temperature and pressure reduction, to a new equilibrium composition. If the expansion is extremely rapid, the mixture may be effectively “frozen” with the initial(high temperature) composition. This would mean that little of the combustion energy of the fuel would be available for acceleration of the combustion product to provide thrust. The chemical kinetics of the recombination processes in the nozzle have a strong effect on the thrust and propulsive efficiency of the engine.
    • COMBUSTOR MATERIALS 1.For resistance against fatigue, Nimonic 75 has been used with Nimonic 80 and Nimonic 90. 2.Nimonic 75 is an 80-20 nickel-chromium alloy stiffened with a small amount of titanium carbide. 3.Nimonic 75 has excellent oxidation and corrosion resistance at elevated temperatures, a reasonable creep strength, and good fatigue resistance. It is easy to press, draw and mold. 4.HA-188, a Cr, Ni-based alloy, has been employed in the latter section of some combustion liners for improved creep rupture strength. 5.In addition to the base material changes, many of today’s combustors also have Thermal Barrie Coatings(TBCs), which have an insulation layer of the total thickness used is 0.4-0.6mm and based on ZrO2-Y2O3 and can reduce metal temperature by 50-150C
    • 2 MARK QUSTIONS 1. What is need for supersonic combustion? 2. Define equivalence ratio and stoichiometric fuel air ratio. 3. Define efficiency of the combustor. 4. What is the purpose of primary air in combustion chamber? 5. What is the purpose of secondary air in combustion chamber? 6. What is the purpose of dilution air in combustion chamber? 7. Define combustion intensity? 8. State the advantages and disadvantages of annular
    • LONG TYPE QUESTIONS 1.What are the important factors affecting combustor design? 2.Write down the methods of flame stabilization and explain with sketch. 3.What are the three types of combustion chamber? Compare its advantages and disadvantages. 4.Name the material used for combustion chamber and discuss the special qualities of the material used for combustion chamber? 5. With the aid of a simplified picture explain the operation of a flame holder. 6.With a neat sketch explain the working of a combustion chamber. 7. Consider n-decane fuel, balance the chemical equation for the stoichiometric combustion of this fuel in air and find the stoichiometric fuel-to-air ratio.
    • UNIT-IV NOZZLES
    • INTRODUCTION • The primary objective of a nozzle is to expand the exhaust stream to atmospheric pressure, and form it into a high speed jet to propel the vehicle. For air breathing engines, if the fully expanded jet has a higher speed than the aircraft's airspeed, then there is a net rearward momentum gain to the air and there will be a forward thrust on the airframe.
    • • Many military combat engines incorporate an afterburner (or reheat) in the engine exhaust system. When the system is lit, the nozzle throat area must be increased, to accommodate the extra exhaust volume flow, so that the turbo machinery is unaware that the afterburner is lit. A variable throat area is achieved by moving a series of overlapping petals, which approximate the circular nozzle crosssection.
    • • At high nozzle pressure ratios, the exit pressure is often above ambient and much of the expansion will take place downstream of a convergent nozzle, which is inefficient. Consequently, some jet engines (notably rockets) incorporate a convergentdivergent nozzle, to allow most of the expansion to take place against the inside of a nozzle to maximise thrust. However, unlike the fixed con-di nozzle used on a conventional rocket mo , when such a device is used on a turbojet engine it has to be a complex variable geometry device, to cope with the wide variation in nozzle pressure ratio encountered in flight and engine throttling. This further increases the weight and cost of such an installation.
    • • The simpler of the two is the ejector nozzle, which creates an effective nozzle through a secondary airflow and spring-loaded petals. At subsonic speeds, the airflow constricts the exhaust to a convergent shape. As the aircraft speeds up, the two nozzles dilate, which allows the exhaust to form a convergent-divergent shape, speeding the exhaust gasses past Mach 1. More complex engines can actually use a tertiary airflow to reduce exit area at very low speeds. Advantages of the ejector nozzle are relative simplicity and reliability. Disadvantages are average performance (compared to the other nozzle type) and relatively high drag due to the secondary airflow. Notable aircraft to have utilized this type of nozzle include the SR-71, Concorde, F-111, and Saab Viggen
    • NOZZLE
    • 1-D ANALYSIS OF GAS
    • MASS FLOW RELATION
    • UNIT-V COMPRESSORS
    • • Compressor – Draws in air & compresses it • Combustion Chamber – Fuel pumped in and ignited to burn with compressed air • Turbine – Hot gases converted to work – Can drive compressor & external load
    • COMPRESSOR • Controlling Load on Compressor – To ensure maximum efficiency and allow for flexibility, compressor can be split into HP & LP sections – Vane control: inlet vanes/nozzle angles can be varied to control air flow • Compressor Stall – Interruption of air flow due to turbulence
    • USE OF COMPRESSED AIR • Primary Air (30%) – Passes directly to combustor for combustion process • Secondary Air (65%) – Passes through holes in perforated inner shell & mixes with combustion gases • Film Cooling Air (5%) – Insulates/cools turbine blades
    • VELOCITY TRIANGLE
    • THANK U