The Aircraft Engine Design Project Fundamentals of Engine Cycles

GE Aviation The Aircraft Engine Design Project Fundamentals of Engine Cycles Spring 2009 Ken Gould Phil Weed 1

GE Aviation Technical History I-A - First U.S. jet engine (Developed in Lynn, MA, 1941) U.S. jet engine U.S. turboprop engine Variable stator t engine Mach 2 fighter engine Mach 3 bomber engine High bypass engine Variable cycle turbofan engine Unducted fan engine GE90 on test 30:1 pressure ratio engine Demonstration of 100k+ engine thrust Certified double annular combustor engine First U.S. turboprop powered aircraft, Dec. 1945 2 T L I K O P F

Flowdown of Requirements 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: Cost and Schedule Engines Systems: Module requirements Technical: Pressure ratio, efficiency, weight, life Program: NRE, part cost, schedule, validation plan FAA/JAA Safety/reliability Noise/emissions Qualified Product 3 Design & Validation

The Aircraft Engine Design Project Fundamentals of Engine Cycles Compressor Combustor HPT Exhaust airflow 4 Turbojet Engine

Turbojet Stations Engine Modules and Components Compressor Combustor HP Spool HPT Exit 0 2 3 4 5 9 Turbojet Engine Cross-Section Multi-stage compressor module powered by a single stage turbine 5

Ideal Brayton Cycle: T-S Representation HP Turbine 4 Expansion Turbine Exit Pressure T Combustor 5 5 Ambient Pressure 3 W Δ pressure available for expansion across Exhaust Nozzle P W 2 = Δ1h 0 Compression 2 Compressor Lines of Constant Pressure Note: 1) Flight Mach = 0 2) P t2 =P amb 3) P = power 4) W = mass flow rate 5) h 0 = total enthalpy 6 S

Real Brayton Cycle: T-S Representation HP Turbine 4 Expansion 5 Turbine Exit Pressure T Combustor 3 Ambient Pressure Δ pressure available for expansion across Exhaust Nozzle P W 2 = Δ1h 0 Compression Impact of Real Efficiencies: Decreased Thrust @ if T4 is maintained Lines of Constant Pressure Or 2 Compressor Increase Temp (fuel flow) to maintain thrust! 7 S

Jet Engine Cycle Analysis Engine Flow capacity (flow function relationship) Starting 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(g c *γ/r) Pt* Ae [1+ ((γ-1)/2)*m 2 ] (γ+1)/[2*(γ-1)] 8 where M is Mach number Tt is total temperature (deg R) Pt is total pressure (psia) W is airflow (lbm/sec) Ae is effective area (in 2 ) gc is gravitational constant =32.17 lbm ft/(sec 2 lbf) γ is ratio of specific heats R is gas constant (ft-lbf)/(lbm-deg R) 0 Turbojet Compressor Combustor HPT Exit HP Spool 2 3 4 5 9

Jet Engine Cycle Analysis Compressor From adiabatic efficiency relationship η compressor = 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 HP Spool Exit 9 0 2 3 4 5 9

Jet Engine Cycle Analysis Combustor From Energy balance/ Combustor efficiency relationship: η combustor = Actual Enthalpy Rise/ Ideal Enthalpy Rise = (WF + W)*Cp combustor Texit W*Cp combustor Tinlet 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) Cp is combustor specific heat BTU/(lbm-deg R) Can express WF/W as fuel to air ratio (FAR) Compressor Combustor HPT Turbojet HP Spool Exit 10 0 2 3 4 5 9

Jet Engine Cycle Analysis Turbine From efficiency relationship η turbine = Actual Work/Ideal Work = Cp*(Tinlet Texit) Cp*(Tinlet Texit ) = 1 - (Texit/Tinlet) 1 - (Pexit/Pinlet) t) (γ-1)/γ Work Balance: From conservation of energy Turbine Work = Compressor Work + Losses (W+ WF)* Cp turb * (Tinlet - Texit) turb = W*C Cp compressor *(T (Texit - Tinlet) comp where Pexit is turbine exit total pressure (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 compressor BTU/(lbm-deg R) 11 0 Turbojet Compressor Combustor HPT Exit HP Spool 2 3 4 5 9

Jet Engine Cycle Analysis Nozzle Isentropic relationship, can determine exhaust propertiesp Tt/Ts= (Pt/Ps) (γ-1)/γ = 1 + ((γ -1)/2) * M 2 From Mach number relationship can determine exhaust velocity v= M*a where a, speed of sound= sqrt(γ*g c *R*Ts) where Tt is total temperature (deg R) Pt is total pressure (psia) Ps is static pressure (psia) Ts is static temp (deg R) g c is gravitational constant =32.17 lbm ft/(sec 2 lbf) γ is ratio of specific heats R is gas constant (ft-lbf)/(lbm-deg R) v is flow velocity y( (ft/sec) a is speed of sound (ft/sec) M is Mach number 12 0 Compressor Combustor HPT HP Spool Turbojet Exit 2 3 4 5 9

Jet Engine Cycle Analysis Engine Performance Thrust relationship: from conservation of momentum Fnet = W9 V9/ g c -W0 V0/ g c + (Ps9-Ps0) A9 If flight Mach number is 0, v0 = 0 and if nozzle expands to ambient, PS9=Ps0 and Fnet = W9 V9/ g c where g c is gravitational constant Specific Fuel Consumption (SFC) SFC = Wf/ Fnet (lbm/hr/ lbf) (lower SFC is better) Compressor Combustor HPT Turbojet HP Spool Exit 13 0 2 3 4 5 9

Modern Afterburning Turbofan Engine Single-stage HPT module A/B w/ Variable Exhaust Nozzle 3-stage fan module Single Stage LPT module 14 Annular Combustor multi-stage compressor module Terms: blade rotating airfoil vane static airfoil stage rotor/stator pair PLA pilot s throttle Typical Operating Parameters: OPR 25:1 BPR 0.34 ITT 2520 o F Airflow 142 lbm/sec Thrust Class 16K-22K lbf

Thermodynamic Station Representation Wf_comb Wf_AB 5 4.5 7 8 9 FN 2 2.5 3 4 Nozzle Expansion W2 Fan Pr (P25/P2) HPC Pr (P25/P2) Comb Temp Rise LP Turbine expansion HP Turbine expansion A/B Temp Rise A8 (nozzle area) 15 Overall Pressure Ratio (P3/P2)

Fan Compressor Air Flow Bypass Flow Combustor HP Spool HPT LPT Afterburner Exit LP Spool Augmented Turbofan Engine Cross-Section 16 General Electric Aircraft Engines

Design Considerations- Process Centering and Variation Off-Target Variation On-Target Center Process Reduce Spread 17 Six Sigma Methodology Applies Statistical Analyses to Center Processes and Minimize Variation General Electric Aircraft Engines

General Electric Aircraft Engines Probabilistic Design Techniques Account for Process Variation 2,000 Trials Frequency Chart 0 Outliers.054.041 Forecast: Margin-: Average Off Target D M O T LSL T 108 81 Forecast: Margin: High Variation 2,000 Trials Frequency Chart 49 Outliers.023 45.017 LSL D M H V T 33.75.027 54.011 22.5.014 27.006 11.25.000 0-2.00 0.00 2.00 4.00 6.00 Certainty is 92.50% from 0.00 to +Infinity D M.000 0-1.00 1.00 3.00 5.00 7.00 Certainty is 95.05% from 0.00 to +Infinity D M Center Process Forecast: Margin:On Target-Low Variation 2,000 Trials Frequency Chart 1 Outlier.052.039.026 O T L V T LSL 104 78 52 Reduce Spread 18.013.000-2.00 0.00 2.00 4.00 6.00 D M 26 0 LSLL S L Understanding and Accounting for Process Variation Assures Compliance with Design Limits