The Aircraft Engine Design Project Fundamentals Of Engine .

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gGE AviationGE Aircraft EnginesThe Aircraft Engine Design ProjectFundamentals of Engine CyclesSpring 20091Ken GouldPhil Weed

gGE Aviation Technical HistoryGE Aircraft EnginesU.S. jet engineU.S. turboprop engineV i bl statorVariablet t engineiMach 2 fighter engineMach 3 bomber engineHigh bypass engineI-A - First U.S. jet engine(Developed in Lynn, MA, 1941)GE90 on testVariable cycle turbofan engineUnducted fan engine30:1 pressure ratio engineDemonstration of 100k engine thrustCertified double annular combustor engineFirst U.S. turboprop powered aircraft, Dec. 19452TLI KFOP

gFlowdown of RequirementsGE Aircraft EnginesThe Customer: Overall system requirementsMTOW, Range, Cost per seat per mileThe Airframer: Sub-system requirementsTechnical: Wing (lift/drag),Engines(Thrust/SFC)Program: CostC and ScheduleSFAA/JAASafety/reliabilityEngines Systems: Module requirementsTechnical: Pressure ratio, efficiency,ffweight, lifefProgram: NRE, part cost, schedule, validation plan3Design &ValidationNoise/emissionsQualified Product

gGE Aircraft EnginesThe Aircraft Engine Design ProjectFundamentals of Engine CyclesCombustorHPTCompressorExhaustairflowInlet4T b j t EngineTurbojetE i

gTurbojet StationsGE Aircraft EnginesEngine Modules and ComponentsCompressorHPTCombustorInletExitHP Spool20345Turbojetj EnginegCross-SectionMulti-stage compressor modulepowered by a single stage turbine59

gIdeal Brayton Cycle: T-S RepresentationGE Aircraft EnginesHP Turbine Inlet4ExpansionTurbine Exit Pressure5T5CombustorInletΔ pressure available forexpansion across Exhaust NozzleAmbient Pressure3CompressionLines of Constant Pressure2Compressor InletS6P Δ21 h0WNote: 1) Flight Mach 02) Pt2 Pamb3) P power4) W mass flow rate5) h0 total enthalpy

gReal Brayton Cycle: T-S RepresentationGE Aircraft EnginesHP Turbine Inlet4ExpansionTurbine Exit Pressure5Δ pressure available forexpansion across Exhaust NozzleTCombustorInletAmbient Pressure3Impact of Real Efficiencies:Decreased Thrust @ if T4 is maintainedCompressionLines of Constant Pressure2OrIncrease Temp (fuel flow) to maintainthrust!Compressor InletS7P Δ21 h0W

gJet Engine Cycle AnalysisGE Aircraft EnginesEngine Inlet Flow capacity (flow function relationship)Startingg with the conservation of mass and substitutingg the total tostatic relations for Pressure and Temperature, can derive:W Density * Area* VelocityW*(sqrt(Tt)) M *sqrt(gc*γ/R)Pt* Ae[1 ((γ-1)/2)*M2] (γ 1)/[2*(γ-1)]where M is Mach numberTt is total temperature (deg R)Pt is total pressure (psia)W is airflowf(lbm/sec)(/)Ae is effective area (in2)gc is gravitational constant 32.17 lbm ft/(sec2 lbf)γ is ratio of specific heatsR is gas constant(ft-lbf)/(lbm-deg R)8TurbojetCompressorExitInlet0Combustor HPTHP Spool23459

gJet Engine Cycle AnalysisGE Aircraft EnginesCompressor From adiabatic efficiency relationshipηcompressor Ideal Work/ Actual Work Cp*(Texit’ – Tinlet)Cp*(Texit – Tinlet) (Pexit/Pinlet)(γ-1)/γ - 1Texit/Tinlet - 1wherePexit is compressor exit total pressure (psia)Pinlet is compressor exit total pressure (psia)Tinlet is compressor inlet total temperature (deg R)Texit is compressor exit total temperature (deg R)Texit’ is ideal compressor exit temperature (deg R)TurbojetCompressorExitInlet90Combustor HPTHP Spool23459

gJet Engine Cycle AnalysisGE Aircraft EnginesCombustor From Energy balance/ Combustor efficiency relationship:ηcombustor Actual Enthalpy Rise/ Ideal Enthalpy Rise (WF W)*CpcombustorTexit – W*CpcombustorTinletWF * FHVwhere W is airflow (lbm/sec)WF is fuel flow (lbm/sec)FHV is fuel heating value (BTU/lbm)Tinlet is combustor inlet total temperature (deg R)Texit is combustor exit total temperature (deg R)Cp is combustor specific heatBTU/(lbm-degBTU/(lbmdeg R)Can express WF/W asfuel to air ratio (FAR)0CompressorCombustor HPTExitInlet10TurbojetHP Spool23459

gJet Engine Cycle AnalysisGE Aircraft EnginesTurbine From efficiency relationshipηturbine Actual Work/Ideal Work Cp*(TinletCp (Tinlet – Texit)Cp*(Tinlet – Texit’) 1 - (Texit/Tinlet)( 1)/1 - (Pexit/Pinlet)(P it/Pi l t)(γ-1)/γ Work Balance: From conservation of energyTurbine Work Compressor Work Losses(W WF)* CpC turb* (Ti(Tinletl t - Texit) T it) turb W * CpC compressor* (T(Texitit - Tinlet) Ti l t) compTurbojetwherePexit is turbine exit total ppressure (p(psia))Pinlet is turbine exit total pressure (psia)Tinlet is inlet total temperature (deg R)Texit is exit total temperature (deg R)Texit’ is ideal exit total temperature (deg R)Cp is specific heat for turbine or compressorBTU/(lbm-deg R)11CompressorExitInlet0Combustor HPTHP Spool23459

gJet Engine Cycle AnalysisGE Aircraft EnginesNozzle Isentropicp relationship,p can determine exhaust ppropertiespTt/Ts (Pt/Ps)(γ-1)/γ 1 ((γ -1)/2) * M2 From Mach number relationship can determine exhaust velocityv M*awhere a, speed of sound sqrt(γ*gc*R*Ts)whereTt is total temperature (deg R)Pt is total pressure (psia)Ps is static pressure (psia)Ts is static temp (deg R)gc is gravitational constant 32.17 lbm ft/(sec2 lbf)γ is ratio of specific heatsR is gas constant (ft-lbf)/(lbm-deg R)v is flow velocityy ((ft/sec))a is speed of sound (ft/sec)M is Mach number12TurbojetCompressorExitInlet0Combustor HPTHP Spool23459

gJet Engine Cycle AnalysisGE Aircraft EnginesE i PerformanceEngineP f Thrust relationship: from conservation of momentumFnet W9 V9/ gc - W0 V0/ gc (Ps9-Ps0) A9If flight Mach number is 0, v0 0pto ambient,, PS9 Ps0 andand if nozzle expandsFnet W9 V9/ gcwhere gc is gravitational constant Specific Fuel Consumption (SFC)TurbojetSFC Wf/ Fnet (lbm/hr/ lbf)(l(lowerSFC isi better)b tt )CompressorExitInlet130Combustor HPTHP Spool23459

gModern Afterburning Turbofan EngineSingle-stage HPT moduleGE Aircraft EnginesA/B w// VVariablei bl ExhaustE ht NozzleNl3-stageg fan moduleSingle Stage LPT moduleAnnular Combustormulti-stage compressor module14Terms:bladevanestagePLArotating airfoilstatic airfoilrotor/stator pairpilot’s throttleTypical Operating Parameters:OPR25 125:1BPR0.34ITT2520oFAirflow142 lbm/secThrust Class16K-22K lbf

gThermodynamic Station RepresentationWf ABWf ABWf comb7GE Aircraft Engines8954.5FN432.52NozzleExpansionA/B Temp RiseLP TurbineexpansionW2Fan Pr(P25/P2)15HPC Pr(P25/P2)Overall Pressure Ratio (P3/P2)CombTempRiseHP TurbineexpansionA8 (nozzlearea)

gGE Aircraft EnginesFanCompressorBypass FlowInletAir FlowHPT LPTCombustorAfterburnerExitHP SpoolLP SpoolAugmented Turbofan Engine Cross-Section16General Electric Aircraft Engines

gGE Aircraft EnginesDesign ConsiderationsConsiderations- Process Centering and VariationOff-TargetVariationXXXX XX XXX XX uceSpreadSix SigmagMethodologygy AppliesppStatistical Analysesyto CenterProcesses and Minimize Variation17General Electric Aircraft Engines

General Electric Aircraft EnginesgGE Aircraft EnginesProbabilistic Designg TechniquesqAccount for Process VariationForecast: Margin-: Average Off Target2,000 TrialsDForecast: Margin: High VariationMFrequencyO ChartT0 Outliers.0542,000 Trials108LSL.041041TDMFrequency ChartHV49 .25.0000000.000000-2.000.002.004.00Certainty is 92.50% from 0.00 to InfinityD6.0033.750-1.00M1.003.005.00Certainty is 95.05% from 0.00 to InfinityD7.00MForecast: Margin:On Target-Low Variation2,000 TrialsCenterProcessO TL V 2.000.002.00D18104TLSL.0391 Outlier4.00LSL LS6.00MUnderstanding and Accounting for Process Variation AssuresCompliance with Design LimitsL

The Aircraft Engine Design Project Fundamentals of Engine Cycles Ken Gould Spring 2009 Phil Weed 1. g GE Aviation Technical History GE Aircraft Engines U.S. jet engine U.S. turboprop engine Vibl tt iVariable stator engine Mach 2 fighter engine Mach 3 bomber engine High bypass engine

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