Thermodynamic Analysis and Performance Characteristics

AEROSPACE ENGINEERING school OF MECHANICAL ENGINEERING AND DESIGN THE THERMODYNAMIC ANALYSIS AND PERFORMANCE CHARACTRISTICS OF A TURBO fan JET ENGINE By J. E, Ibok 2011 Supervisor Dr Lionel Ganippa ABSTRACT This civilize focuses on the exploit analytic thinking of a collimate bobbin interracial fertilise win outright kelvin locomotive locomotive. The principal(prenominal) objective was to investigate the effects of victimisation total heat, lamp oil and natural fluid give the axe on the surgery characteristics much(prenominal) as take in s springy, particularised go off expending and propulsive efficacy of the devotee pitchy locomotive.An adjourn re principal of this influence was to move into the concept of exergy and thermoeconomics digest for enoughise bobbin compound feed in turbo fountain railway locomotive and certify the t come aliverical roles that passs the to the broad(prenominal)est degree to the in strength of the railway l ocomotive. A generic postureing was carried out victimisation be adrift Turb 11 parcel to obtain reasonable compendium go outs that were verify with a tangible- cadence JT8D-15A turbo run railway locomotive railway locomotive. The parametric compend was do for everlasting value of tummy f commencement lay out of fire and unbroken turbine uptake temperature for all trey sacks.The prove were rightfully obtained for these digest discip auras and discussed accordingly. Brunel University robotic locomotive locomotive railway locomotiveering Academic Session 2010/2011 marches of Student Johnson Essien Ibok SupervisorDr Lionel Ganippa Title The ope symmetryn Characteristics and Thermodynamics Exergy and Thermoeconomics epitome of a similitude transfer Mixed catamenia Turbolover railway locomotive Operating at 30,000ft at M0 0. victimisation lamp oil, natural Gas and hydrogen open fire. Abstract This figure focuses on the transaction synopsis of a twin spool composite f scummy turbo putting green locomotive locomotive locomotive locomotive engine. A generic accumulation line was carried out employ Gas Turb 11 softwargon program to obtain reasonable abstract results that were corrobo pasture with a real-time JT8D-15A winnow super acid engine engine. The parametric analysis was done with(p) for immutable value of cumulation f moo pass judgment of go off and continual turbine entre temperature for all ternary fuels.The result were rightfully obtained for these analysis human faces and discussed accordingly. Objectives The master(prenominal) perplex of this work is to conduct the parametric make pass manakin of a twin spool commingle f dis modulateed-toned fan-jet engine and investigate the mathematical physical subroutine characteristics of it. An few sepa send aim of this work is to indicate the effects of victimization hydrogen, Kerosene and natural flatulency fuel on the boilersuit effe ctance of the twin spool mixed play fanjet engine engine.Also, the answer of this work is to introduce the use of the act justness of thermodynamics analysis known as exergy and thermoeconomics in analysis the twin spool mixed f hapless turbojet engine engine Background/Applications This work is applicable in so m any shipway when it comes to the general carrying into action optimization and feasibility analysis of a park engine. This work relates to the aerospace and aviation industries since the turbojet engine engine is amongst the vast number of fountain engine employ in propulsion of billetcrafts.thither is increasing compel in the aviation industry to reduce pollution and depletion of zipper resources spot at the equivalent time maintaining reasonable investment damage and high boilers suit consummation. Hence, this research was conducted in hopes of coming up with a current etymon to this problem. Conclusions The main conclusion drawn from the carrying out analysis is that hydrogen fuel produced the highest crusade level and the funkyest peculiar(prenominal) fuel custom surrounded by the triple fuels for a constant push-down list devolve rate of fuel.Kerosene fuel generated squash level house be growthd if it is mixed with a small amount of hydrogen. The expire greens velocity proportionality go alonged constant despite the increasing shunt ratio for all common chord fuels at constant piling liquefy rate of fuel. development the exergetic analysis showed that the blaze put up and the mixer contri howevered the most to the in qualification of the turbofan engine. The amount of exergy transferred into the turbofan engine by hydrogen was depleted in the smallest ratio comp argond to natural botch up and kerosene for constant mass f junior-grade rate of fuel.The thermoeconomics analysis showed that it is favored to use local found cost evaluation to quantity proper(postnominal) thermoeconomics cost of ro botlike press than the globular method since the value was displace. Results The results obtained from the crinkle victimisation Gas Turb 11 produced an error hunt down of 0. 25% 8. 5% when verified with the substantial bear witness entropy of the JT8D-15A turbofan engine. The results obtained for the analysis positd a summons contrive battery-acid at which the parametric analysis was conducted on. The analysis was do in 3 cases as shown clearly in the probe matrix in table 1 at a cast down place. compend Parameters cosmos varied Parameters unbroken eonian Performance Characteristics case 1 * go around ratio * Turbine consumption temperature * HPC constrict balance * LPC contract dimension * strike out bosom ratio * speeding ratio * discharge-Air-ratio * Turbine inlet temperature * Net urge on * particularized deplume out habit * caloric ability * Propulsive efficiency case 2 * rotate proportion * Three diametric fuelsmH2mCH4mC12H23 * Mass come down rate of fuel * HPC haul proportionality * LPC twitch proportion * caramel brown force per social unit bowl dimension instance 3 * Bypass ratio * Three dissimilar fuelsmH2mCH4mC12H23 * Turbine inlet temperature * HPC pull proportionality * LPC Pressure Ratio * buffer Pressure Ratio circuit card 1 The streamlet matrix of the parametric Analysis. The exergy analysis was done for the parametric analysis of case 2 and case 3 where the exergy destruction rates, exergetic efficiency, exergy improvement potential rate and fuel depletion ratio were enumerated. The diffusion of these results throughout apiece element of the turbofan engine was lay outed with draw a blank charts and Grassmann draw. The thermoeconomics analysis was conducted for analysis case 2 apply kerosene fuel.The specialised thermoeconomics cost of confuse was calculated use global and local based cost evaluation methods. ACKNOWLEDGEMENTS First of all, I would manage to thank my parg onnts for their mo straighten outary support and encouragement because without them I would not be here and be able to do this work. I am deeply thankful to my supervisor, Dr Lionel Ganippa for call up in me and giving me the opportunity to work with him in this field of get word. I am in any case thankful to him for giving the necessary guidance and advice and his fervency and innovative ideas inspired me. Finally, I would like to thank Mr Joachim Kurzke for providing me with the necessary softwargon needed for my disquisition. control board of ContentsAcknowledgements i Contents ii List of Notations and Subscripts iv List of Tables vi List of identification numbers vi Chapter 1 knowledgeability1 1. 1. Aims and Objectives2 1. 2. Computational Modeling3 Chapter 2 super acid Engines4 2. 1. Performance characteristics4 2. 1. 1. englut4 2. 1. 2. Thermal faculty5 2. 1. 3. Propulsive efficiency5 2. 1. 4. boilers suit efficiency6 2. 1. 5. Specific dismiss utilisation6 2. 2. give notice and Propellants For outpouring Engines7 Chapter 3 Turbofan Jet Engines 8 3. 1. Introduction 8 3. 2. smorgasbord of Turbofan Engines9 3. 3. Major Components of a Turbofan Engine10 3. 3. 1. diffuse10 3. 3. 2. Fan and Compressor11 3. 3. 3. Combustion bedchamber12 3. 3. 4. Turbine13 3. 3. 5. Exhaust Nozzle14 3. 4.Thermodynamic demonstrate and round of drinks of a equalise bobbin Mixed ladder Turbofan Engine15 Chapter 4 Mathematical and Gas turb 11 Modeling of the turbofan Engine18 4. 1. Station Numbering and Assumptions18 4. 2. Design address turn Simulation of the Turbofan Engine18 4. 3. Off- devise formulaate round Simulation of the Turbofan Engine21 4. 3. 1. faculty/Component Matching 22 4. 3. 2. Off-Design Point Component Modeling22 Chapter 5 Methodology, Results and Discussions26 5. 1. ecumenical Relationship equations of the Major Parameters27 5. 2. Results and Discussions of Parametric roulette wheel Analysis of Case 129 5. 3. Results and Discuss ions of Parametric Cycle Analysis of Case 235 5. 4.Results and Discussions of Parametric Cycle Analysis of Case 343 Chapter 6 Exergy and Thermoeconomics Analysis of the Turbofan Engine49 6. 1. Exergy Analysis49 6. 1. 1. Exergy Analysis Modeling 50 6. 1. 2. Exergy and zip fastener Balance comparisons of the Components58 6. 1. 3. pla unclutterary Relationships in Exergetic Analysis of the Turbofan Engine60 6. 1. 4. Results and Discussions61 6. 1. 5. Grassmann Diagram72 6. 2. Thermoeconomics Analysis74 6. 2. 1. Thermoeconomics Analysis copy74 6. 2. 2. world-wide Based Cost Evaluation76 6. 2. 3. Local Based Cost Evaluation77 6. 2. 4. Results and Discussion of the Thermoeconomics Analysis78 Chapter 7 Conclusions and Future Work80 Reference Appendix A Exergy Analysis Results Appendix B Thermoeconomics Analysis resultsList of Notations and Units ? isentropic efficiency ? derive Pressure ratio m Mass lead tread (kg/s) f enkindle/Air Ratio M Mach Number Pt Total press (kPa) Tt Tota l Temperature (K) NCV Net Calorific Value (MJ/kg) Ht Total Enthalpy (kJ/kg) V hurrying (m/s) ? Bypass Ratio T Static Temperature (K) P Static Pressure (kPa) N Actual spool festinate (RPM) Nc Corrected Spool whet (RPM) mc Corrected Mass Flow roll (kg/s) R universal joint Gas Constant (kJ/kmolK) ?0 Standard Chemical Exergy (kJ/kmol) Ex Exergy regularize (MW) xi Mole Fraction cp Specific Heat at Constant Pressure (kJ/kgK) ? Ratio of Chemical Exergy to NCV ? Exergetic Efficiency give the sack Depletion Ratio W Power estimate of Work done (MW) List of Subscripts LPT Low Pressure Turbine HPT soaring Pressure Turbine CC Combustion Chamber HPC High Pressure Compressor LPC Low Pressure Compressor d Diffuser noz Nozzle mix mixer dest Destruction Rate 0, ambFAR Ambient conditionFuel-Air-Ratio CH Chemical PH Physical KN energizing PN Potential IP Exergy good Potential Rate (MW) CRF Cost recuperation Factor c Specific Thermoeconomic Cost (MJ/kg) STD Standard Temperature and Press ure knocker Turbine Inlet Temperature TSFC Thrust Specific Fuel usage (g/kNs) SFC Specific Fuel Consumption p Propulsive TH ThermalO Overall T Thrust equip Equipment PEC Capital Cost of Equipment List of Tables Table 1 gossip disputations for Design Point Cycle Simulation on Gas Turb 1119 Table 2 Comparison table for the Actual canvass info and Simulated Data victimization gas Turb 1121 Table 3 Comparison Table for Actual Test Data and Simulated Off-Design Point selective information exploitation gas Turb 11. 25 Table 4 equivalence Ratio of the terzetto Fuels Combustion branches 62 Table 5 anticipate Capital costs of Each Component of the Turbofan Engine. 75 Table 6 Flow of Specific Thermoeconomics Cost in all the Components 79 List of dates puzzle out 1 Classification of Turbofan Engine9 mannikin 2 Layout of away Fan Twin Spool Mixed Flow Turbofan16 Figure 3 T-S Diagram for the Forward Fan Twin Spool Mixed Flow Turbofan17 Figure 4 Design Point Cycle Simulation algori thm use Gas Turb 1120 Figure 5 Example of a Compressor Performance Map/Curve24 Figure 6 effects of alter Bypass Ratio at Constant rig of depreciator on Fuel-Air-Ratio30 Figure 7 Effects of Varying Bypass Ratio at Constant determine of depreciator on Exit Velocity Ratio30 Figure 8 Effects of Varying Bypass Ratio at Constant Values of TIT on LPT Exit Pressure Ratio31 Figure 9 Effects of Varying Bypass Ratio at Constant Values of TIT on Net Thrust32 Figure 10 Effects of Varying Bypass Ratio at Constant Values of TIT on Specific Fuel Consumption33 Figure 11 Effects of Varying Bypass Ratio at Constant Values of TIT on Propulsive Efficiency34 Figure 12 Effects of Varying Bypass Ratio t Constant Values of TIT on Thermal Efficiency35 Figure 13 T-S diagram of using Hydrogen Fuel when the go around Ratio is increased36 Figure 14 var. of Fuel-Air-Ratio with Bypass Ratio at Constant Fuel Flow Rate using three contrary Fuels37 Figure 15 pas seul of TIT with Bypass Ratio at Constant F uel Flow Rate using three contrasting Fuels37 Figure 16 Variation of Exit Velocity Ratio with Bypass Ratio at Constant Fuel Flow Rate using three opposite Fuels38 Figure 17 Variation of LPT Exit Pressure Ratio with Bypass Ratio at Constant Fuel Flow Rate using three several(predicate) Fuels39 Figure 18 Variation of Net Thrust with Bypass Ratio at Constant Fuel Flow Rate using three different Fuels40 Figure 19 Variation of Specific Fuel Consumption with Bypass Ratio at Constant Fuel Flow Rate using three different Fuels41 Figure 20 Variation of Thermal Efficiency with Bypass Ratio at Constant Fuel Flow Rate using three different Fuels42 Figure 21 Variation of Propulsive Efficiency with Bypass Ratio at Constant Fuel Flow Rate using three different Fuels43 Figure 22 Variation of Fuel-Air-Ratio with Bypass Ratio at Constant TIT using the three varied Fuels44 Figure 23 Variation of Exit Velocity Ratio with Bypass Ratio at Constant TIT using the three Different Fuels44 Figure 24 Varia tion of LPT Exit Pressure Ratio with Bypass Ratio at Constant TIT using the three Different Fuels45 Figure 25 Variation of Net Thrust with Bypass Ratio at Constant TIT using the three Different Fuels46 Figure 26 Variation of Specific Fuel Consumption with Bypass Ratio at Constant TIT using the three Different Fuels46 Figure 27 Variation of Propulsive Efficiency with Bypass Ratio at Constant TIT using the three Different Fuels47 Figure 28 Variation of Thermal Efficiency with Bypass Ratio at Constant TIT using the three Different Fuels48 Figure 29 Variation of Exergy Destruction Rate Using the three Fuels for Analysis Case 262 Figure 30 Variation of Exergy Destruction Rate Using the three Fuels for Analysis Case 364 Figure 31 Variation of Exergetic Efficiencies Using the three Fuels for Analysis Case 266 Figure 32 Variation of Exergetic Efficiencies Using the three Fuels for Analysis Case 367 Figure 33 Distribution of Exergy Improvement potential Rate Using the three Fuels for Analysi s Case 268 Figure 34 Distribution of Exergy Improvement potential Rate Using the three Fuels for Analysis Case 369 Figure 35 stochastic variable of Fuel Depletion ratio using the Three Fuels for Analysis Case 270 Figure 36 variation of Fuel Depletion ratio using the Three Fuels for Analysis Case 371 Figure 37 Grassmann Diagram for the Exergetic analysis of Case 2 using kerosene Fuel for the Turbofan engine. 72 Chapter 1 Introduction Jet engines are complex thermodynamic dusts that use a series of non-linear equation to define their thermodynamic mental accomplishes and they operate under the principle of Brayton circle.Brayton speech rhythm is a cycle that comprises of the compressor, combustor and turbine operative as a unit. Additionally, the major parameters that dictate the ope logical conditions of the engine at any point during the process are the relative height and Mach number. Mach number is the ratio of the velocity of the outpouring engine to the speed of sound. Basically, the main aspire of this type of thermodynamic system in aerospace industry is to speed a course of circularize and as a result, generate equal shake off needed for shoot. In admission, the rule of jet engines is dependent of what dissolve it forget be employ for in nine to derive its supreme performance.For instance, in military application, jet engines are required to generate upper limit thrust in b identifyline response time which consumes a deal of fuel whereas commercial jet engines are required to less upset generative, less fuel over situationing and at the uniform time score high overall efficiency (El-sayed, 2008). There are certain factors that jet engine manufacturers take into consideration when trope jet engines which are the operate cost, engine noise, environmental emissions, fuel burn and overall efficiency. Accordingly, this has ca utilize a global market competition for engine manufacturers like Rolls Royce, Pratt and Whitney, Gene ral electric automobile and CFM on who so-and-so produce the most efficient jet engines.In fact, Pratt and Whitney Company is working on a adapt turbofan jet engine that they believe evidenceament reduce fuel burn, produce lesser noise and distribute less toxics enchantment General Electric is coming up with simpler ecore jet engines that go away be more fuel efficient than the current jet engines with as much as virtually two fifths of current jet engines (Cassidy, 2008). Taking all that has been express into consideration, it evict easily be take a firm stand that by reducing the fuel consumption of the jet engine, the rack up temperature at the turbine blades guide reduce at that placeby increasing the direct life and overall efficiency of the engine. Also, the total cost of the engine cigarette be down down. Indeed, Dr Pallan cited in (Ward, 2007) supposed that reducing the fuel consumption by as teentsy as 1% is highly longed after by engine manufacturers an d this evoke result in very signifi fuckingt increase in the overall performance.In a general point of view, it can be express that the uttermost point of achievement for jet engine manufacturers would be to design an engine that consumes the minimum amount of work in the compressor unit time generating the maximum amount of work in the turbine unit at minimum fuel supply. The main purpose of this work is to meditate the thermodynamic processes and performance of a jet engine using a modeling withall, exergy and thermoeconomics concept. 1. 1. Aims and Objectives The main objective of this work is to carry out the thermodynamic analysis and show the performance characteristics of a turbofan jet engine. In this work, the vivid history of the thermodynamics processes and cycle of each parting of the turbofan engine starting from the diffusor to the snout will be covered. Also, the first and punt law of thermodynamics with former(a) laws will be applied extensively througho ut this work.However, in the case of performance characteristics of the turbofan engine, a generic simulation will be carried out on a twin spool mixed flow turbofan engine. To relate this work to real life application, a JT8D-15A turbofan engine manufactured by Pratt and Whitney Company will be use as the twin spool mixed flow turbofan for the simulation using the legitimate design data. Indeed, the simulation tool that will be employ is GasTurb 11 which was intentional by Joachim Kurke and for more details on how it works can be found in (Kurke, 2007). This work will use the reference design point of the twin spool mixed flow turbofan at sea level with maximum take-off thrust to obtain the operating(a) point of 30,000ft at M0 0. using the off-design performance simulation which will serve as the operating design point for the analysis in this work since the engine will spend most of its time in the cruise phase between 30000ft to 38000ft. The purpose of carrying this generic s imulation of the turbofan engine is to investigate the effects of varying bypass ratio and turbine inlet temperature (thermic limit parameter) on the performance characteristics of the turbofan engine. In other words, the parametric cycle studies of the turbofan engine. This investigation will be done for three different cases which case 1 will be analyze the effects of varying bypass ratio and turbine inlet temperature on the performance characteristics of the turbofan engine when some(prenominal) of the design choices are kept constant.The second case of report will be the comparison of the performance characteristics of the turbofan engine when three different fuels (kerosene, natural gas and Hydrogen) are used at the same mass flow rate using the same design point in case 1. Finally, the third base case of field of honor will be the comparison of the performance characteristics of the turbofan engine when the three fuels are undergoing the same blaze process that is const ant turbine inlet temperature for the design point in case 1. This aspect of this analysis is very all important(predicate) owing to the growing problem of glasshouse effect and depletion of vim resources. In fact, statistics by the intergovernmental panel shows that aerospace industry is amongst one of the fast growing sources of greenhouse effect and that the emission of carbon dioxide will increase to five round what it is presently which is 3% (Symonds, 2005).Based on this, using utility(a) fuels like hydrogen and natural gas can tend to reduce pollution and consumption of muscle resources risk and this work aims to show how that can be achieved patch the overall efficiency of the engine is still high. Another approach of analysis in this work will be the use of the second law of thermodynamics analysis also known as exergy and thermoeconomics. This aspect of analysis of the turbofan engine will be done for the parametric analysis of case 2 and case 3 in efforts to also c ompare the three fuels that are being considered and show which fuel will cause the turbofan engine atoms to be most inefficient or pass on the most irreversibility.This analysis will be done by calculating the exergy relationships such as exergy transfer rates, exergy destruction rates, exergetic efficiencies, exergy improvement potential rates, and fuel depletion ratios. Furthermore, the exergy analysis will be fight down in a Grassmann diagram for parametric analysis case 2 of paper. However, as for the thermoeconomics analysis of the turbofan engine, entirely parametric analysis case 3 studies will be done for exclusively kerosene fuel and this work will aim to show how to use concept of local and global evaluation of thermoeconomic cost. 1. 2. Computational modelling It will be very expensive and time wasting to design and develop new breedcraft engine whenever an optimization or analysis wants to be done.In fact, Caoa Y, Jin, Meng and Fletcher (2005) utter that new w ays should be developed to reduce spreadcraft engine design, maintenance and manufacturing cost in order to have effective planetary market competition. Surprisingly, computer modelling is one approach of reducing manufacturing cost and time wasting. Computational modelling can simply be delimit as the use of computer codes to replicate a typical system using some of its original data in order to analyse the system at varying conditions. The other side of the medallion shows simulation. There are many types of simulation tools unremarkably used in simulating gas turbines such as Matlab/simulink, Modelica, Gas Turb 11, NPSS and many more. However, the simulation tool that will be adopted for the purpose of this dissertation is Gas Turb 11 designed by Joachim Kurzke.Gas Turb 11 is a language oriented program with a command prompt that calculates the output data without using block diagrams or graphical interface. It is user friendly in a sense that it is indulgent to find the too ls library and to substitute data in for simulation. The Gas Turb 11 is unique(predicate)ally designed for simulation of all kinds of gas turbines starting from cause generators to jet engines. Gas Turb 11 comm further carries out two types of analysis which are the on design cycle point simulation and off-design cycle point simulation. Engine design point cycle simulation makes the larn of comparing gas turbines of different geometry. This cycle design point moldiness be specify before any other simulation can be done.On the other hand, off-design performance cycle point simulation involves the study of the behaviour of a gas turbine with known geometry. This cycle outlines the performance characteristics of each component such as performance maps, Overall efficiency. The type of simulation that will be done in this dissertation will involve the off-design and design point cycle. Chapter 2 Jet Engines 2. 1. Performance Parameter of Jet Engines 2. 2. 1. Thrust Thrust is the wa y of quantifying the ability of a jet engine to effectively utilise the energy added to it in order to propel or push itself forward in the opposite counselling of the exiting jet in the exhaust nose.In other words, it is the reactive force to the force imparted by the exiting jet in the beak in accordance to Isaac normalitys third law of motion. It is the most important parameter that has to be obtained for any jet engine and it depends heavily on the ingested mass of air, exiting velocity and blackmail, the area of the nose, the flight velocity and close conditions. In fact, the numerical expression for thrust which incorporates these factors is shown below as. Thrust=meVe-m0V0+Pe-P0Ae Where, e=the exit conditions at the exhaust nozzle, 0=ambient conditions at the inlet me=m0+mfuel Momentum Thrust=meVe This is the thrust obtained from the reaction of the hot exhaust gases high velocity.Momentum Drag= m0V0 This the friction or drag force caused by the high velocity ingestio n of air mass at the inlet. Pressure Thrust=Pe-P0Ae This force is generated as a result of the higher(prenominal) exit nonoperational pull compared to the ambient wardrobe level which pushes back at the engine. Gross Thrust=meVe+Pe-P0Ae It is the maximum obtainable positive thrust a jet engine can have when the drag forces are ignored. specific Cases of Thrust Take-off Thrust It is the thrust a jet engine can generate with its own reason at static or low power setting which heart the momentum drag component of thrust is ignored and the power of the engine at this point is equivalent to zero.This can be used to explain why the thrust of an engine at take-off condition is usually higher than at cruise condition since in that location is no momentum drag and effects of varying ambient condition. This only applies to turbojet, turbofan, and propjet jet engines but when it comes to ramjet and scramjet, the air flow has to be accelerated by a booster system before it can start pr oducing a positive take-off thrust. Pressure Thrust Component This is the thrust generated as a result of the static racks of the exiting jet and ambient environment. In ideal cases where the nozzle has perfectly expand the jet exit mechanical press to that of the ambient condition, the atmospheric stuff thrust component will disappear which this case is not possible in reality.However, if the nozzle is choked which indicates that the ambient blackjack is lower than the exit pressure of the jet, the pressure thrust component will have a positive effect on the net thrust. Also, if the nozzle tends to over expand the jet because of low energy addition to the jet and the exit pressure is lower than the ambient pressure, the pressure thrust component will have a negative effect on net thrust. 2. 2. 2. Thermal efficiency It is simply the measure at which energy in the engine system is born-again. In other words, it is the measure at which total energy supplied to the engine system as change transfer is converted to kinetic energy.In another way, it can easily be said to be the ratio of the power generated in the engine air flow to the rate at which energy is supplied in the fuel. ?TH=Power Generated in the Engine air flowRate of push Supplied in the Fuel =12? meVe2-12? m0V02mfuel? NCV 2. 2. 3. Propulsive efficiency It is a measure at which kinetic energy possessed by air as it passes through the engine is converted into power of the propulsion of the engine. In mathematical terms, it is simply known as the ratio of thrust power to the power generated in the engine airflow. ?p=Thrust PowerPower Generated in the Engine Airflow = T? V012? meVe2-12? m0V02 2. 2. 4. Overall EfficiencyAs the name overall depicts, it is the resultant efficiency of a jet engine can have which is simply the intersection of the thermal and propulsive efficiencies. In mathematical terms, it is represented as shown below. ?O=? TH?? p =12? meVe2-12? m0V02mfuel? NCV? T? V012? meVe2-12? m0V02 =T? V0mfuel? NCV 2. 2. 5. Specific Fuel Consumption Specific fuel consumption as any other performance characteristics is a ratio and surprisingly it has a major effect on the economics of the aircraft as it is used to determine the aircrafts flight ticket costs. Specific fuel consumption has different expressions depending on what type of jet engine it is. For instance, in ramjet, turbojet and turbofan jet engines, it is the measure of the fuel mass flow rate to the thrust force generated.Also, it is sometimes called the thrust specific fuel consumption (TSFC). TSFC=mfT However, in turbopropeller jet engines, it is the ratio of the fuel mass flow rate to the power generated in the engine shaft by the turbomachinery. It is sometimes referred to as the brake-specific fuel consumption (BSFC). TFSC=mfSP 2. 2. Fuel and Propellants for Jet Engines Fuels can implicitly be be as substances used to add heat energy to a system through combustion or other processes. Fuels are mostly h ydrocarbons like kerosene, diesel, petrol, alcohol, paraffin and butane and can also be in the form of individually free reactive molecular substances like hydrogen or chemical composites like natural gas, coal, wood.The gaseous state substances used as fuels such as hydrogen, and natural gas (94% methane and 6% ethane) are usually made into a cryogenic state as in liquefied at very low temperature because of their low boiling point. It can easily be asserted by anyone that the only purpose that fuels have in jet engines is to add energy but pocket-sized do they know that the purposes grows as the speed of the aircraft increases. For instance, Kerrebrock (2002) stated that supersonic aircrafts which creates very high stagnation temperature that can create destabilization to the airframe structure, engine component and organic substances like lubricants, uses its fuel as a coolant to this parts or components.The energy added by the fuel burned per unit mass of air flow is called t he passion value of the fuel and it is a very life-or-death parameter to be defined before any combustion process analysis is done on a jet engine since it shows how be intimate the combustion process is through efficiency. The heating system value can all be said to be higher or lower depending on if the water product of combustion is a vapour or a liquid. Since the combustion process in jet engine produces vaporised water, the lower heating value of the fuel is used. The most frequently used fuels for jet engines are kerosene jet A1, A2, JP10 and many more but diesel can also be used. The disadvantages of these fuels are their inevitable emission of toxic substances that contribute to greenhouse effect and their risk of depletion.Accordingly, this has been the driving force for the use of alternative fuels such as cryogenic hydrogen and natural gas which is believed will reduce toxic emissions. Besides, hydrogen is a carbon-free energy carrier and possesses almost no risk of toxic emission since most of its combustion product will be water Chiesa and Laozza (2005). Chapter 3 Turbofan Jet Engine 3. 1. Introduction Between 1936 and the next decade when turbofan engines were invented, people showed tiny or no interest in them as they described them to be a complicated version of a turbojet engine. However, in 1956, the benefits of turbofan engines started to be noticed as major companies like Rolls-Royce and General Electric began manufacturing them.Since then, it is been one of the most used jet engine for commercial purposes because of its low fuel consumption and less noise production. In fact, it has been concluded to be the most reliable jet engine ever manufactured El Sayed (2008). The turbofan jet engine gas generator unit comprises of a fan unit, compressor section, combustion chamber and turbine unit. Fundamentally, a turbofan jet engine operates as a result of the compressors pressuring air and supplying it afterwards for progress processing. T he majority of the pressurised air is bypassed around the core of the engine through a duct to be mixed or exhaust whereas the rest of it flows into the main engine core where it combusts with the fuel in the combustion chamber.The hot expanded gas products from the combustion process passes through the turbine thereby rotating the turbine as it leaves the engine. Consequently, the rotating turbine spins the engine spool which in turn rotates the other turbo machinery in the engine. This causes the front fan to pressurise more and more air into the engine for the process to start all over again in continuous state. The turbofan engine is believed to be the perfect combination of the turboprop and turbojet engine and as a result, its advantages are usually compared to that of the turboprop and turbojet. In fact, Kerrebrock (1992) said that turbofan engine provides a better way of improving the propulsive efficiency of a basic turbojet.It is asserted that at low power setting, low t op condition and low speed, the turbofan engine is more fuel efficient and has better performance than a turbojet engine. Unlike turboprop engine where vibration occurs in the propeller blades at relative low velocities, the fan in the turbofan engine can attain high relative velocities of Mach 0. 9 before vibration occurs. Also, since the fan in turbofan engines has many blades, it is more stable than the single propeller so counterbalance if the vibration velocity is reached, the vibration will not destabilize the airflow because the vibrations are almost negligible. Since the flow into the diffuse of the turbofan is usually subsonic, there very slim chances of shock waves being developed at the entrance. 3. 2. Classification of Turbofan EnginesThere are various types of turbofan engine ranging from high and low bypass ratio, afterburning and non-afterburning, mixed and unmixed flow with multi-spool, after fan and geared or ungeared. The classification of the various types of t urbofan engines is shown below in physique 1. Nonetheless, the type of turbofan engine that would be used for the purpose of this dissertation is a forward fan two spool mixed flow turbofan engine. This type of turbofan engine was chosen because it is the compromise of a simple and complex turbofan engine. This is said because it comprises of almost all the classes of a turbofan which are low bypass ratio, forward fan with mixed flow, twin spool with ungeared fan.Moreover, because of the mixed flow introduced, it produces additional thrust in the hot nozzle compared to the high bypass and it can also permit the addition of afterburner which produces a lot of thrust while consuming a lot of fuel which makes it fitting for military application which shows little worry on fuel consumption. In essence, carrying out a study on this type of turbofan engine will be of expectant relevance to the military air force sector especially if new research is discovered. TURBOFAN ENGINES Low Bypa ss Ratio Aft Fan Forward Fan Nonafterburning Afterburning High Bypass Ratio Geared Fan Single Spool Short Duct Ungeared Fan ii Spool Mixed Fan and Core Flow Unmixed Flow Long Duct Three SpoolFigure 1 Classification of Turbofan Jet Engines (El-sayed, 2008) 3. 3. Major Components of Turbofan Engine 3. 4. 1. Diffuser or Inlet Diffuser is the first component that air encounters as it flows into the engine. Basically, the purpose of a distributed is to suck in air smoothly into the engine, reduce the velocity of the air, increase the static pressure of the air and finally, supply the air in a uniform flow to the compressor. Given the fact that overall performance of an engine is highly dependent on the pressure supplied to the burner, it is necessary to design a diffuse that incurs the minimum amount of pressure loss.To demonstrate this, Flack (2005) stated that if the diffuser incurs a large total pressure loss, the total pressure in the burner will be reduced by the compressor tot al pressure ratio time this loss. In other words, a small pressure drop in the diffuser can read into a square drop in the total pressure supplied to the burner. Another point interpreted into consideration when designing a diffuser is the angle because if the angle is too big, there will be tendency of eddy flow generation delinquent to early separation. The major causes of pressure losses in the diffuser are as follows. First, losses due(p) to generation of shock waves outside(a) the diffuser and it majorly occur in supersonic diffusers.Secondly, the loss due to the admonishing or adverse pressure gradient of the diffuser geometry which makes the flow separate a lot earlier and generates eddies. This separation causes a convergent area which makes the velocity not to be reduced by much. Due to the separation, the surround shear deteriorates the static pressure even come on. Further analysis done by El-Sayed (2008), describes ways of accounting for this losses like using Fa nno line flow and combined area and friction. Thermodynamic Process Equation In this analysis, the loss due to heat transfer is negligible so the process can be adiabatic. The initial kinetic energy is used to bone the static pressure p0 to the total pressure ? =pt2pt0 (inlet pressure recovery) efficiency ? d=IdealReal=ht2s-h0ht2-h0 assuming the gas is ideal and the specific heat at constant pressure is constant efficiency ? d=Tt2s-T0Tt2-T0 simplifying the equation inclined that ht0=ht2=ht2s and Tt2=Tt0and pt2s=pt2 TtT0=1+? -12M02 and TtT0=ptp0? -1? pt2p0=1+ ? d? -12M02?? -1 3. 4. 2. Fan And Compressors Compressor is a very crucial component for the operation of an engine in the sense that it prepares the air for the combustion process in the burner. The main purpose of a compressor as the first rotating component is to use its rotating blades to add kinetic energy to the air and later translate it into total pressure increase.There are basically two types of compressors which are the outward-developing and the axial compressor. Firstly, centrifugal compressor as the name implies changes the direction of an axial airflow to a radial outflow of the air. It was the early compressors adapted in jet engines. It comprises of three main parts which are the impellers, the diffusers and the compressor manifold. The purpose of the impeller is to change the direction of the flow from axial to radial and at the same time increases its static pressure. The diffuser slows down the airflow and further increase the static pressure as it is supplied axially by the compressor manifold to the combustion chamber.The centrifugal compressor is profitable because the cost of manufacturing it is low compared to axial compressor and as a result is capable for small engines like turboshafts and turboprops. It is also advantageous because the pressure ratios at single stage are higher than that of the axial compressor. The centrifugal compressor has the tendency of attaining low flow rates and as a result is ideally equal for helicopters and small aircrafts which require low flow rates. On the other hand, the centrifugal compressor cannot attain high pressure ratio and so it is not suitable when high peak efficiency is required. It incurs a lot of losses due to the change in direction. Secondly, an axial compressor is the most reliable type of compressor and is usually applied when higher pressure ratios of up to 401 are required.An axial compressor does not change the axial flow direction of the air but increases the total pressure. Indeed, an axial compressor comprises of three major components which are the rotor coil with blades, stator can and the inlet guide vane. A stage is a combination of a stator and a rotor. The assembly of the full rotor blade and stator can form the number of stages in a compressor and the greater the number of stages, the higher the total pressure ratio. In this arrangement, the air flows into the inlet guide vane and then into the rotor and stator assembly where compression starts. Also, the length of the rotor and stator reduces along the whole unit which signifies a reduction in volume which induces the increase in pressure.A fan or low pressure compressor is a type of axial compressor but the only differences are that the blades are longer, the total pressure ratio is lower than the typical compressor and the number of stages is usually 1 or 2. The main purpose of creating a fan is to compress more air and to create a bypass air which can be used to generate addition thrust or used for mixing process. Fan Equation Process Given that, isentropic efficiency ? fan= Ideal CycleActual cycle=ht3s-ht2ht3-ht2 Since the specific heat is constant, the equation deduces to ? fan=Tt3s-Tt2Tt3-Tt2 Simplifying the equation whenpt3s=pt3, Tt3sTt2=pt3pt2? -1? , ? fan=pt3pt2 and ? fan=Tt3Tt2 ? fan=? fan? -1? -1? fan-1 Bypass Ratio=msma where ms is the bypass flow rate and ma is the engine core flow rate.For the high pressure compressor, the equations remain the same as that of the fan except the changes in put up numeration and the bypass ratio. 3. 4. 3. Combustion Chamber/ Burner The combustion chamber as the Brayton cycle implies is the only source of heat energy addition to the system. Accordingly, the combustion chamber causes very significant increase in the temperature of the air which results in the air gaining enormous internal energy. This energy gained is extracted to be used to power the turbine while the rest is used to create highly accelerated gases from the nozzle. There are three types of combustor namely the can combustor, the annular combustor and the cannular combustor.The main considerations when designing a combustion chamber is to ensure that the combustion process is land up with no fuel waste, the combustor should have long life materials because any trial can lead to engine explosion. The other consideration is that the air mustiness be heated enough above the igni tion fuel temperature in order to ensure stoichiometric combustion. Equations of the Combustion Chamber In the real process of the combustor, total and static pressure drops and the temperature also drop. The major causes of pressure losses are the high level of irreversibility or non-isentropic process and glutinous effects in the burner. The burner pressure ratio ? =pt5pt4Burner temperature ratio ? b=Tt5Tt4 Since no work is done only heat transfer, the efficiency of the burner is analysed using the heating value NCV of the fuel used. Thus, efficiency ? b=heat addedHeating value of fuel=ma+mfht5-maht4NCVmf Given that f=mfma, ? b= 1+fht5- ht4NCVf Equivalence Ratio of combustion It is the ratio of the actual fuel to air ratio of the combustion process to the stoichiometric fuel to air ratio. This ratio produces a manner of classifying the combustion process to show whether it is a lean, rich or stoichiometric combustion. The mathematical expression for this is as shown below ? =Act ual FARStiochiometric FAR <1 Lean combustion process ?=1 Stiochiometric combustion process ?>1 Rich combustion process 3. 4. 4. Turbine Turbine can simply be said to be the antonym of a compressor. In response, a turbine extracts molecular kinetic energy from the air and uses it to drive the turbo machineries which results in the pressure and temperature of the air to drop. If lawfulness be told, Flack (2005) asserted that the turbine uses 70% to 80% of the total energy gained by the air in the combustion chamber to drive the turbo machineries while the remaining 20% to 30% is used to generate thrust in the nozzle.Since the geometry of a turbine have favourable pressure gradient unlike the compressor which is adverse, the efficiency of the turbine is usually very high. Since the turbine is the opposite of the compressor, it has exactly the same configuration of rotor and stator but the volume increase across it which induces the pressure drop. nonpareil major problem faced whe n design a turbine is the deterioration of the blades due to high inlet temperature from the combustion chamber. Based on this, (Song et al. 2002) demonstrated that General Electric uses about 16. 8% of the compressor air to cool the turbine blades of GE 7f engine. Turbine Equation Analysis Given that, Turbine efficiency ? T=ActualIdeal=ht6-ht5ht6s-ht5 T=Tt6-Tt5Tt6s-Tt5 Simplifying the equation given that pt6s=pt6 Tt6sTt5=pt6pt5? -1? ?T=pt6pt5 ? T=Tt6Tt5 ?T=? T-1? T? -1? -1 3. 4. 5. Exhaust Nozzle The nozzle is the final component of the jet engine that the air passes through. The main purposes of the nozzle is to add extra acceleration to the high velocity exiting air, reduces its total pressure to that of ambient condition and finally generate sufficient thrust. There are two conditions that occur in the exit of the nozzle depending on the ambient pressure. The first condition is termed under-expansion which occurs when the ambient pressure is less than the exit pressure of the ga ses.The result of this is that the exit velocity will be lower than it normally is and this makes the momentum component of thrust to be lower than ideal. On the other hand, it will create a positive thrust component for the pressure terms. The second case termed as overexpansion which occurs when the ambient pressure is greater than the exit pressure of the gases. Consequentially, the opposite of what happens in the under-expansion condition occurs where the pressure term is lower and the momentum is higher. Nozzle efficiency ? n=ActualIdeal=ht8-h9ht8-h9s=Tt8-T9Tt8-T9s for constant specific heat Using the steady state energy equation and balancing it out, U9=2ht8-h9 . When specific heat is constant U9=2cpTt8-T9 p9pt8=T9sTt8? -1? T9Tt8=11+? -12M92 p9pt8=11+? -12M92-1+ ? n ? n 3. 4.Thermodynamic Process and Cycle of Twin Spool Mixed Flow Turbofan Engine Before any explanation is done from Figure 2, the blue arrows represent the entrance air into the diffuser and the red represent th e air flow into the core of the engine while the black arrow represent the bypass air flow through the fan. Finally, the brown arrow represents the air flow after the bypass air and the core air flow have mixed. Based on the arrangement of the turbofan engine in manikin 2, it can be seen that air at ambient condition is sucked into the diffuser where the air velocity is reduced and some of its kinetic energy is used to increase the static pressure to the total pressure. The air exiting the diffuser enters the fan or low pressure compressor where it is compressed. Indeed, the molecules of the air gains kinetic and internal energy by colliding rapidly with one another and as a result increase the enthalpy and static pressure.Also, in the fan, some of the compressed air is bypassed through a duct to be used for the mixing process later while the rest of the air enters into the high pressure compressor of the engine core. In the high pressure compressor, the air is further compressed w here the enthalpy and pressure increases as it is released into the combustion chamber. Also, in the high pressure compressor, some of the air mass flow rate is bled out to be used to cool the turbine blades and for air conditioning in the aircraft. In the combustion chamber, the incoming fuel reacts with the air in an oxidation process at constant pressure where the by-product gases gain molecular kinetic energy thereby increasing the enthalpy.This high temperature gases operates into the high pressure turbine where it is expanded and the gases lose some of their kinetic molecular energy as it enthalpy and static pressure reduces. In other words, it can be said that the molecular kinetic energy of the gases is being converted to mechanical work which is used to power the high pressure spool. Consequently, the gases enters into the low pressure turbine where it is further expanded to a lower pressure and enthalpy as their molecular kinetic energy is converted to mechanical work to power the low pressure spool. These gases escaping from the low pressure turbine enters the mixing zone or mixer after it has disconnected most of its total enthalpy and mixes with the bypassed coolness air from the duct to further reduce its enthalpy as that of the cold air increases.In other words, the cold air absorbs some of the heat energy from the hot gases until they both attain equilibrium enthalpy. The mixture of the cold air and hot gases both escape at the same equilibrium enthalpy and pressure through the nozzle where their velocity is increased and the pressure is reduced considerably to that of the ambient condition. Furthermore, the exhausted high velocity gases is used to produced thrust for propulsion according to Newtons third law of motion (In every action, there is equal and opposite reaction). 2 4. 5 6 4 13 0 HPC DIFFUSER FAN/LPC HPT LPT NOZZLE COMBUSTION CHAMBER 2. 5 3 5 8 16 BYPASS DUCT HP Spool LP Spool MIXING ZONEFigure 2 Layout of a Forward Fan Twin Spool Mixed Flow Turbofan Engine P0 P3 P4. 5 P5 P8 P6 P2. 5 P2 P13 P4 ENTROPY (S)(kJ/kg) TEMPERATURE (K) Figure 3 T-S Diagrams for the Forward Fan Twin Spool Mixed Flow Turbofan Engine Chapter 4 Mathematical and Gas Turb 11 Modelling of the Engine 4. 1. Station Numbering and Assumptions Station amount is a very crucial step that has to be taken when analysis of any thermodynamic system involving many processes is to be done. Moreover, send out numbering contributes immensely to showing how the properties of one process relate to another and how the interaction between these processes derives the functional relationship of the thermodynamic system.Returning to the work in hand, the station numbering system that has been adopted for this work on a JT8D-15A turbofan engine is in accordance with the Aerospace Recommended Practice (ARP) and it is shown in figure 2. Assumptions The spare-time activity assumption were made based on Mattingly (2002) and Kurzke (2007) in order to perform th e modelling as listed below * The air flow through the engine is assumed to be steady and one dimensional * The fan and the low pressure Compressor are driven by the low pressure turbine * The overall engine is assumed to have no bleeds in mass flow or power off-take in turbine. * The nozzle of the engine is choked which means the exit pressure will be greater than the ambient pressure. The air is assumed to act as a half ideal gas where the specific heat and ratio is dependent on temperature only. * The areas of each station of the engine is assumed to be constant 4. 2. Design Point Cycle Analysis of the Turbofan Engine The off-design or performance cycle analysis cannot be done without the design point cycle being defined. The design point cycle in this analysis is obtained using exactly the same data used in the actual screen analysis for a JT8D-15A turbofan engine operating at sea level with maximum take-off thrust as shown in (JT8D Typical Temperature and Pressure) and (ICAO). Some of the input parameters such as the isentropic efficiencies and pressure ratios from the actual test data had to be calculated.Since not all the input parameters were given from the actual test data, some of the parameters like inlet corrected mass flow rate, diffuser pressure ratio and efficiency mechanical spool efficiency had to be guessed in order to complete the analysis and the data are represented below in Table 1. With all the stimulant Parameter being stipulate as shown in table 1, the design point cycle simulation of the JT8D-15A turbofan Engine using the Gas Turb 11 parcel can then be performed. All the steps taken to model the mixed flow turbofan engine on Gas Turb 11 is clearly represented in the algorithm shown in figure 3 below. COMPONENT INPUT PARAMETER DIFFUSER Pressure Ratio (? d) 1 Inlet Corrected Mass Flow Rate (mc2) 138. 618 kg/s FAN Pressure Ratio (? fan) 2. 054 Isentropic Efficiency (? fan) 0. 78 Bypass Ratio (? ) 1. 08 Low Pressure Compressor (LPC ) Pressure Ratio (? LPC) 4. 7 Isentropic Efficiency (? LPC) 0. 88 Nominal Low Pressure Shaft stimulate (NLP) 8160RPM High Pressure Compressor (HPC) Pressure Ratio (? HPC) 3. 77 Isentropic Efficiency (? HPC) 0. 864 Nominal Low Pressure Shaft Speed (NHP) 11420RPM Combustion Chamber (cc) Pressure Ratio (? CC) 0. 934 Isentropic Efficiency (? CC) 0. 99 Burner Exit Temperature (TIT) 1277. 15K High Pressure Turbine (HPT) Isentropic Efficiency (? HPT) 0. 9 HP Spool Mechanical efficiency (? m) 1 Low Pressure Turbine (LPT) Isentropic Efficiency (? LPT) 0. 91 LP Spool Mechanical efficiency (? m) 1 Table 1 Input Parameters for the Design Point Cycle Simulation STARTSpecify all the input data gotten from the actual test data as shown in Table 1 Run the Gasturb 11 software and select mixed flow turbofan from the drag down Tab list. Set the scope to More, set the Calculation Mode as Design and click Run subscribe to the Units to either Imperial or SI and Select the type of fuel from to d rop down list to Kerosene, Natural Gas or Hydrogen Estimate the inlet Corrected mc2 Mass Flow rate to the FAN/LPC Choose Single Cycle for Select a Task pickax and click Run Check if the Thrust, SFC, ? HPT, ? LPT and EPR are within (0-10) % of the actual test Experiment END YES NO Figure 4 Design Point Cycle Simulation Algorithm Using Gas Turb 11 validation of the Design Point simulation ResultsSince not all the input parameters were specified in the actual test data and some of them had to be guessed, it is without any doubt that errors are bound to generate in the simulation results using the Gas Turb 11 software. In order to ensure that the errors accumulated in the simulation were within range, the major output parameters obtained such as net thrust, fuel flow rate, Engine exit pressure ratio, etc were compared to the actual test data as shown in Table 2 and the error range was calculated to be between 0. 25% to 8. 5% which is within an unimpeachable range. PARAMETERS ACTUAL T EST DATA simulated DATA USING GASTURB 11 Net Thrust 69307. 74 69320 Engine Exit Pressure Ratio P8P0 2. 09 2. 167Burner Fuel Flow 1. 100843 1. 09781 HPT pressure Ratio (? HPT) 0. 415 0. 449 LPT Pressure Ratio (? LPT) 0. 3294 0. 3514 HPT temperature Ratio (? HPT) 0. 8097 0. 8435 LPT temperature Ratio (? LPT) 0. 7718 0. 793 Table 2 Comparison Table for the Actual Test Data and Simulated Data Using GasTurb 11 4. 3. Off-Design Point Cycle Simulation of the Turbofan Engine The off-design or performance cycle simulation takes into account the concept of module matching of each component through performance maps. This cycle analysis enables the determination of different operating point of the engine at a given design point of the engine.Considering the work in hand, the design point have been defined and verified for the JT8D-15A turbofan engine operating at sea level with maximum take-off thrust which means that different operating points of the engine can be defined with the concept of off-design module matching of the engine. Indeed, the off-design operating point that was considered for the parametric analysis in this work was 30,000ft at M0 0. 8 for the turbofan engine. The off-design modelling of the JT8D-15A engine for the operating point of 30,000ft at M0 0. 8 based on the reference design point defined earlier is clearly demonstrated as follows. The off-design performance cycle simulation may contain some errors because of the component performance maps that were used for the simulation. 4. 3. 1. Module/Component Matching This process only applies to the off-design performance cycle point of the engine.It can simply be defined as the act of synchronising each component of a jet engine to coexist as a unit in order to derive the overall performance characteristics of the jet engine. Component matching involves the process closely studying the ramifications of the actual jet engine overall performance behaviour on the components major characteristics such as pressure ratio, temperature ratio, efficiency and spool speed. This process introduces the concept of empirically determined component performance maps that establishes the relationship between the thermodynamic properties and the geometry of the jet engine itself. 4. 3. 2. Off-Design Component Modelling Diffuser The diffuser was assumed to be adiabatic and the pressure ratio ? d=1 The Isentropic Efficiency was assumed to be 1 For Sea Level,Pamb=101325pa , Tamb=288. 15K For 30,000ft and M0 0. 8, Tamb=288. 15-0. 0065? 9144 =288. 15-59. 436 =228. 71K Pamb=101325? Tamb288. 155. 2561 =30. 09kpa Tt1=228. 71? 1+? -12M02 =228. 71? 1+1. 4-12? 0. 82 =258K pt1p0=1+ ? d? -12M02?? -1 pt1=30. 09? 1+ 1? 1. 4-120. 821. 41. 4-1 pt1=45. 8674kPa pt1=pt2 Tt1=Tt2 Fan and Low Pressure Compressor The inlet corrected mass flow rate is estimated as 138. 618kg/s , As for the off design simulation using the component performance maps for the altitude of 30000ft and Mach no. 0. 8, the actual spool speeds and inlet mass flow rate are calculated based on the estimated inlet corrected mass flow rate as shown below.Low and High pressure spool mechanical efficiency is assumed to be=1 HP spool Speed=11420RPM, LP spool Speed=8160RPM m2=Pt2PSTD? mc2Tt2TSTD =45. 878101. 325? 138. 618258288. 15 Actual Mass flow rate m2=66. 3323kg/s N=Tt2TSTD? NcLP=228. 71288. 15? 8160=7722 RPM The calculated actual mass flow rate and spool speed were used to evaluation the isentropic efficiency and the pressure ratio of the LPC for that operating condition from the compressor performance map. Figure 5 Example of a Compressor Performance Map/Curve The diagram above in figure 4 depicts a typical compressor performance map that was used for the off-design point analysis in this work.It can be seen that the x-axis represents the inlet corrected mass flow rate mc2 into the compressor, the y-axis represents the compressor pressure, the red signifier lines represents the isentropic efficiencies and the black curved lin es represent the relative corrected spool speed. To add to that, the red dash line that ends the speed lines and efficiency lines represent the surge margin which is also known as the stall line that must be avoided since the flow will become unstable in that region. In this work, the inlet corrected mass flow rate and spool speed were calculated which were interpolated on the performance map to obtain the pressure ratio and the isentropic efficiency.For instance, the yellow dot on the map represents a design point traced for a given pressure ratio, High Pressure Compressor The inlet corrected mass flow rate into the HPC mc2. 5=mc21+? mc2. 5=138. 6182. 08=66. 64kgs m2. 5=Pt2. 5PSTD? mc2. 5Tt2. 5TSTD N=Tt2. 5TSTD? NcHP The same equation used for the LPC is used to calculate the actual mass flow rate and spool speed which is used to evaluate the isentropic efficiency and pressure ratio when it is operating at an altitude of 30000ft at M0 =0. 8. Verification of the off-design modelling for 30000ft at Mo 0. 8 In order to verify the simulation result gotten for the operational design point of 30000ft at M0 0. , the actual test data results gotten from Mattingly, Heiser and Pratt (2002) for the same operating condition was compared. Due to the difficulties in obtaining a lot of output parameters for this operating point, the result will be verified with only the net thrust generated and the specific fuel consumption. Indeed, the error accumulated was 1. 71% for the net thrust and 0. 83% for the specific fuel consumption. PARAMETERS ACTUAL TEST DATA SIMULATED DATA USING GASTURB 11 Net Thrust (lb) 4920 4836 Specific Fuel Consumption(lb/lbh) 0. 779 0. 7855 Table 3 Comparison Table for the Actual Test Data and Simulated Off-design Data Using GasTurb 11 Chapter 5Methodology, Results and Discussions Given that the design point of the JT8D-15A turbofan engine at sea level has been obtained and verified with the actual test data, the operating point of 30000ft at M0 0. 8 wa s simulated and obtained which now served as the design point for the analysis in this work. Moreover, the procedure taken to define this design point of 30000ft at M0 0. 8 of the JT8D-15A turbofan engine has been clearly stated earlier which gives the permission to conduct the parametric cycle study of the turbofan engine. The parametric cycle studies were done for three different cases for the operational design point of 30000ft at M0 0. of the JT8D-15A turbofan engine as explained as follows. 1. The first parametric analysis case 1 aim to create an understanding of the effects of varying major design parameters on the performance parameters of the turbofan engine when some of the design choices are kept constant. In other words, the bypass ratio and thermal limit parameter (turbine inlet temperature) were varied when the design choices such as the compressor pressure ratio, fan pressure ratio and isentropic efficiencies were kept constant in order to investigate their effects on the performance parameters such as the net thrust, specific fuel consumption, propulsive efficiency, thermal efficiency, and fuel-air-ratio.Much interest is shown nowadays in using alternative fuels like hydrogen and Natural gas in efforts to reduce the cancer known as pollution and the risk of depletion of energy resources. Based on this, conducting a research that focuses of comparing different fuels consumption rate, their risk of pollution and their contribution to the performance of the engine will be really valuable. Based on this, a parametric analysis had to be done on the JT8D-15A turbofan engine using three different fuels which are the design point fuel kerosene, hydrogen and natural gas. Since the original design point of the JT8D-15A turbofan was obtained using kerosene fuel, the design points of using hydrogen and natural gas was obtained using the same design choices as that of kerosene.Now that the design points of the JT8D-15A turbofan engine had been defined when u sing the three different fuels, it had given a go ahead to perform whatever parametric cycle studies of the turbofan engine using the three fuels. In order to compare the performance characteristics of the turbofan engine when it is using the three different fuels, different approaches had to be devised to compare them effectively on a rational basis which defines the last two parametric analysis cases as follows. 2. The second case of parametric analysis was that the fuel flow rate would be kept constant for the three fuels that would be used as the bypass ratio is varied with design choices remaining the same. 3.The third case of study was to make the energy supply into the combustion chamber of the turbofan engine the sa