GE Aviation g GE Aircraft Engines The Aircraft Engine Design Project Fundamentals of Engine Cycles Ken Gould Spring 2009 Phil Weed 1 GE Aviation Technical History g GE Aircraft Engines U.S. jet engine U.S. turboprop engine ViblttVariable stator eng ine Mach 2 fighter engine Mach 3 bomber engine I-A - First U.S. jet engine High bypass engine GE90 on test (Developed in Lynn, MA, 1941) Variable cycle turbofan engine Unducted fan engine 30:1 pressure ratio engine Demonstration of 100k+ engine thrust Certified double annular combustor engine First U.S. turboprop powered aircraft, Dec. 1945 T L I K O P 2 F g Flowdown of Requirements GE Aircraft Engines The Customer: Overall system requirements MTOW, Range, Cost per seat per mile The Airframer: Sub-system requirements Technical: Wing (lift/drag),Engines(Thrust/SFC) Program: CSCost and Schedule FAA/JAA Safety/reliability Engines Systems: Module requirements Noise/emissions Technical: Pressure ratio, efficiency, weight, li fe Program: NRE, part cost, schedule, validation plan Qualified Product Design & 3 Validation g GE Aircraft Engines The Aircraft Engine Design Project Fundamentals of Engine Cycles Combustor HPT Compressor Exhaust airflow Inlet TbjtEiTurbojet Engine 4 g Turbojet Stations GE Aircraft Engines Engine Modules and Components Compressor HPT Combustor Inlet Exit HP Spool 0 2 3 4 5 9 Turbojjget Engine Cross-Section Multi-stage compressor module powered by a single stage turbine 5 g Ideal Brayton Cycle: T-S Representation GE Aircraft Engines HP Turbine Inlet 4 Expansion Turbine Exit Pressure 5 5 Δ pressure available for expansion across Exhaust Nozzle T Combustor Ambient Pressure P 2 Inlet = Δ1h0 3 W Compression Note: 1) Flight Mach = 0 2) P =P= P Lines of Constant Pressure t2 amb 3) P = power 4) W = mass flow rate 2 Compressor Inlet 5) h0 = total enthalpy S 6 g Real Brayton Cycle: T-S Representation GE Aircraft Engines HP Turbine Inlet 4 Expansion Turbine Exit Pressure 5 Δ pressure available for expansion across Exhaust Nozzle Combustor T Ambient Pressure P 2 Inlet = Δ h 3 W 1 0 Impact of Real Efficiencies: Compression Decreased Thrust @ if T4 is maintained Lines of Constant Pressure Or 2 Compressor Inlet Increase Temp (fuel flow) to maintain thrust! S 7 g Jet Engine Cycle Analysis GE Aircraft Engines Engine Inlet • Flow capacity (flow function relationship) Startinggg with the conservation of mass and substituting the total to static relations for Pressure and Temperature, can derive: W= Density * Area* Velocity W*(sqrt(Tt)) = M *sqrt(gc*γ/R) Pt* Ae [1+ ((γ-1)/2)*M2] (γ+1)/[2*(γ-1)] where M is Mach number Tt is total temperature (deg R) Turbojet Pt is total pressure (psia) W is airf(/)flow (lbm/sec) Compressor Combustor HPT Ae is effective area (in2) gc is gravitational constant =32.17 lbm ft/(sec2 lbf) Inlet Exit γ is ratio of specific heats HP Spool R is gas constant (ft-lbf)/(lbm-deg R) 8 0 2 3 4 5 9 g Jet Engine Cycle Analysis GE Aircraft Engines Compressor • From adiabatic efficiency relationship ηcompressor = Ideal Work/ Actual Work = Cp*(Texit’ – Tinlet) Cp*(Texit – Tinlet) = (Pexit/Pinlet)(γ-1)/γ - 1 Texit/Tinlet - 1 where Pexit 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) Turbojet Compressor Combustor HPT Inlet Exit HP Spool 9 0 2 3 4 5 9 g Jet Engine Cycle Analysis GE Aircraft Engines Combustor •From Energy balance/ Combustor efficiency relationship: ηcombustor = Actual Enthalpy Rise/ Ideal Enthalpy Rise = (WF + W)*CpcombustorTexit – W*Cp combustorTinlet WF * FHV where 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) Turbojet Cp is combustor specific heat BTU/(lbm-deg R) Combustor Can express WF/W as Compressor HPT fuel to air ratio (FAR) Inlet Exit HP Spool 10 0 2 3 4 5 9 g Jet Engine Cycle Analysis GE Aircraft Engines Turbine • From efficiency relationship ηturbine = Actual Work/Ideal Work = Cp*(Tinlet – Texit) Cp*(Tinlet – Texit’) = 1 - (Texit/Tinlet) 1 - (Pex it/Pin le t)(γ-1)/γ • Work Balance: From conservation of energy Turbine Work = Compressor Work + Losses (W+ WF)* Cpturb*(Tilt* (Tinlet - Tit)|Texit)|turb = W*CW * Cpcompressor*(T* (Tex it - Tilt)|Tinlet)|comp where Turbojet Pexit is turbine exit total pp(p)ressure (psia) Pinlet is turbine exit total pressure (psia) Compressor Combustor HPT Tinlet is inlet total temperature (deg R) Texit is exit total temperature (deg R) exit’ Inlet Exit T is ideal exit total temperature (deg R) HP Spool Cp is specific heat for turbine or compressor BTU/(lbm-deg R) 11 0 2 3 4 5 9 g Jet Engine Cycle Analysis GE Aircraft Engines Nozzle • Isentropppic relationship, can determine exhaust ppproperties Tt/Ts= (Pt/Ps)(γ-1)/γ = 1 + ((γ -1)/2) * M2 • From Mach number relationship can determine exhaust velocity v= M*a where a, speed of sound= sqrt(γ*gc*R*Ts) where Tt is total temperature (deg R) Pt is total pressure (psia) Ps is static pressure (psia) Turbojet Ts is static temp (deg R) gc is gravitational constant Compressor Combustor HPT =32.17 lbm ft/(sec2 lbf) γ is ratio of specific heats R is gas constant (ft-lbf)/(lbm-deg R) Inlet Exit v is flow velocityy( (ft/sec) HP Spool a is speed of sound (ft/sec) M is Mach number 12 0 2 3 4 5 9 g Jet Engine Cycle Analysis GE Aircraft Engines EiEngine PfPerformance • Thrust relationship: from conservation of momentum Fnet = W9 V9/ gc -W0 V0/ gc + (Ps9-Ps0) A9 If flight Mach number is 0, v0 = 0 and if nozzle expp,ands to ambient, PS9=Ps0 and Fnet = W9 V9/ gc where gc is gravitational constant • Specific Fuel Consumption (SFC) Turbojet SFC = Wf/ Fnet (lbm/hr/ lbf) (lower SFC i s be tter) Compressor Combustor HPT Inlet Exit HP Spool 13 0 2 3 4 5 9 g Modern Afterburning Turbofan Engine GE Aircraft Engines A/B w / Var ia ble E xh aust N ozzl e Single-stage HPT module 3-stage fan module Single Stage LPT module Annular Combustor multi-stage compressor module Typical Operating Parameters: Terms: OPR 25: 1 blade rotating airfoil BPR 0.34 vane static airfoil ITT 2520oF stage rotor/stator pair 14 Airflow 142 lbm/sec PLA pilot’s throttle Thrust Class 16K-22K lbf g Thermodynamic Station Representation GE Aircraft Engines Wf_ AB 7 8 9 5 Wf_comb 4.5 FN 4 3 2.5 2 Nozzle Expansion A/B Temp Rise LP Turbine W2 expansion A8 (nozzle HP Turbine Comb expansion area) Temp HPC Pr Rise Fan Pr (P25/P2) (P25/P2) 15 Overall Pressure Ratio (P3/P2) g GE Aircraft Engines Fan Compressor Bypass Flow HPT LPT Inlet Combustor Afterburner Air Flow Exit HP Spool LP Spool Augmented Turbofan Engine Cross-Section 16 General Electric Aircraft Engines g GE Aircraft Engines Design Considerations - Process Centering and Variation Off-Target Variation XXX X X X XX X X XXXX X X X X X X X X On-Target X XXXX X X XXXX Center XXX Reduce Process Spread Six Sigggyppyma Methodology Applies Statistical Analyses to Center Processes and Minimize Variation 17 General Electric Aircraft Engines General Electric Aircraft Engines g GE Aircraft Engines Probabilistic Desiggqn Techniques Account for Process Variation Forecast: Margin-: Average Off Target Forecast: Margin: High Variation D M H V 2,000 Trials D MFrequency O Chart T 0 Outliers 2,000 Trials Frequency Chart 49 Outliers .054 108 .023 45 LSL T LSL T .041 81 .017 33.75 .027 54 .011 22.5 .014 27 .006 11.25 .000 0 .000 0 -2.00 0.00 2.00 4.00 6.00 -1.00 1.00 3.00 5.00 7.00 Certainty is 92.50% from 0.00 to +Infinity Certainty is 95.05% from 0.00 to +Infinity D M D M Forecast: Margin:On Target-Low Variation 2,000 Trials O T FrequencyL V Chart 1 Outlier .052 T 104 Center LSL Reduce Process .039 78 Spread .026 52 .013 26 .000 0 LSLL S L -2.00 0.00 2.00 4.00 6.00 DM Understanding and Accounting for Process Variation Assures 18 Compliance with Design Limits.