g
GE Aviation
GE Aircraft Engines
The Aircraft Engine Design Project Fundamentals of Engine Cycles
Spring 2009 1
Ken Gould Phil Weed
g
GE Aviation Technical History
GE Aircraft Engines
U.S. jet engine U.S. turboprop engine V i bl stator Variable t t engine i Mach 2 fighter engine Mach 3 bomber engine High bypass engine
I-A - First U.S. jet engine (Developed in Lynn, MA, 1941)
GE90 on test
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
2
T
L
I K F
O
P
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: Cost C and Schedule S
FAA/JAA Safety/reliability
Engines Systems: Module requirements Technical: Pressure ratio, efficiency, ff weight, life f Program: NRE, part cost, schedule, validation plan
3
Design & Validation
Noise/emissions
Qualified Product
g
GE Aircraft Engines
The Aircraft Engine Design Project Fundamentals of Engine Cycles Combustor
HPT
Compressor
Exhaust
airflow Inlet 4
T b j t Engine Turbojet E i
g
Turbojet Stations
GE Aircraft Engines
Engine Modules and Components Compressor HPT Combustor Inlet Exit HP Spool
2
0
3
4
5
Turbojet j Engine g Cross-Section Multi-stage compressor module powered by a single stage turbine 5
9
g
Ideal Brayton Cycle: T-S Representation
GE Aircraft Engines
HP Turbine Inlet
4
Expansion Turbine Exit Pressure
5
T
5
Combustor Inlet
Δ pressure available for expansion across Exhaust Nozzle Ambient Pressure
3
Compression Lines of Constant Pressure
2
Compressor Inlet
S 6
P = Δ21 h0 W
Note: 1) Flight Mach = 0 2) Pt2 = Pamb 3) P = power 4) W = mass flow rate 5) h0 = total enthalpy
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
T
Combustor Inlet
Ambient Pressure
3
Impact of Real Efficiencies: Decreased Thrust @ if T4 is maintained
Compression Lines of Constant Pressure
2
Or Increase Temp (fuel flow) to maintain thrust!
Compressor Inlet
S 7
P = Δ21 h0 W
g
Jet Engine Cycle Analysis
GE Aircraft Engines
Engine Inlet • Flow capacity (flow function relationship) Starting g with the conservation of mass and substituting g 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) Pt is total pressure (psia) W is airflow f (lbm/sec) ( / ) Ae is effective area (in2) gc is gravitational constant =32.17 lbm ft/(sec2 lbf) γ is ratio of specific heats R is gas constant (ft-lbf)/(lbm-deg R) 8
Turbojet Compressor
Exit
Inlet
0
Combustor HPT
HP Spool
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
Exit
Inlet
9
0
Combustor HPT
HP Spool
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*CpcombustorTinlet 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 BTU/(lbm deg R) Can express WF/W as fuel to air ratio (FAR)
0
Compressor
Combustor HPT
Exit
Inlet
10
Turbojet
HP Spool
2
3
4
5
9
g
Jet Engine Cycle Analysis
GE Aircraft Engines
Turbine • From efficiency relationship ηturbine = Actual Work/Ideal Work
= Cp*(Tinlet Cp (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 energy Turbine Work = Compressor Work + Losses (W+ WF)* Cp C turb* (Ti (Tinlet l t - Texit)| T it)|turb = W * Cp C compressor* (T (Texit it - Tinlet)| Ti l t)|comp Turbojet
where Pexit is turbine exit total p pressure (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 compressor BTU/(lbm-deg R) 11
Compressor
Exit
Inlet
0
Combustor HPT
HP Spool
2
3
4
5
9
g
Jet Engine Cycle Analysis
GE Aircraft Engines
Nozzle • Isentropic p relationship, p can determine exhaust p properties p 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) Ts is static temp (deg R) gc is gravitational constant =32.17 lbm ft/(sec2 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
Turbojet Compressor
Exit
Inlet
0
Combustor HPT
HP Spool
2
3
4
5
9
g
Jet Engine Cycle Analysis
GE Aircraft Engines
E i Performance Engine P f • Thrust relationship: from conservation of momentum Fnet = W9 V9/ gc - W0 V0/ gc + (Ps9-Ps0) A9 If flight Mach number is 0, v0 = 0 p to ambient,, PS9=Ps0 and and if nozzle expands
Fnet = W9 V9/ gc where gc is gravitational constant
• Specific Fuel Consumption (SFC) Turbojet
SFC = Wf/ Fnet (lbm/hr/ lbf) (l (lower SFC is i better) b tt )
Compressor
Exit
Inlet
13
0
Combustor HPT
HP Spool
2
3
4
5
9
g
Modern Afterburning Turbofan Engine Single-stage HPT module
GE Aircraft Engines
A/B w// V Variable i bl Exhaust E h t Nozzle N l
3-stage g fan module
Single Stage LPT module
Annular Combustor multi-stage compressor module
14
Terms: blade vane stage PLA
rotating airfoil static airfoil rotor/stator pair pilot’s throttle
Typical Operating Parameters: OPR 25 1 25:1 BPR 0.34 ITT 2520oF Airflow 142 lbm/sec Thrust Class 16K-22K lbf
g
Thermodynamic Station Representation Wf AB Wf_AB Wf_comb
7
GE Aircraft Engines
8
9
5 4.5
FN
4 3 2.5
2
Nozzle Expansion
A/B Temp Rise LP Turbine expansion
W2
Fan Pr (P25/P2)
15
HPC Pr (P25/P2)
Overall Pressure Ratio (P3/P2)
Comb Temp Rise
HP Turbine expansion
A8 (nozzle area)
g
GE Aircraft Engines
Fan
Compressor Bypass Flow
Inlet
Air Flow
HPT LPT Combustor
Afterburner Exit
HP Spool LP Spool
Augmented Turbofan Engine Cross-Section
16
General Electric Aircraft Engines
g
GE Aircraft Engines
Design Considerations Considerations- Process Centering and Variation Off-Target
Variation
X XXX X X XXX X X X
X X
X X X
X
X
X X X
On-Target
X
Center Process
XX XXX X XX X XXX
X
Reduce Spread
Six Sigma g Methodology gy Applies pp Statistical Analyses y to Center Processes and Minimize Variation 17 General Electric Aircraft Engines
General Electric Aircraft Engines
g
GE Aircraft Engines
Probabilistic Design g Techniques q Account for Process Variation Forecast: Margin-: Average Off Target 2,000 Trials
D
Forecast: Margin: High Variation
MFrequencyO Chart T
0 Outliers
.054
2,000 Trials
108
LSL
.041 041
T
D
M
Frequency ChartH
V
49 Outliers
.023
45
T
LSL
81
.017
.027
54
.011
22.5
.014
27
.006
11.25
.000 000
0
.000 000
-2.00
0.00 2.00 4.00 Certainty is 92.50% from 0.00 to +Infinity
D
6.00
33.75
0 -1.00
M
1.00 3.00 5.00 Certainty is 95.05% from 0.00 to +Infinity
D
7.00
M
Forecast: Margin:On Target-Low Variation 2,000 Trials
Center Process
O T
L V Chart Frequency
.052
Reduce Spread
78
.026
52
.013
26
.000
0 -2.00
0.00
2.00
D 18
104
T
LSL
.039
1 Outlier
4.00
LSL L
S
6.00
M
Understanding and Accounting for Process Variation Assures Compliance with Design Limits
L
