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GE Aviation
GE Aircraft Engines
The Aircraft Engine Design Project
Fundamentals of Engine Cycles
Spring 2009
1
Ken Gould
Phil Weed
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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
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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)
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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
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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
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Methodology
gy Applies
pp
Statistical Analyses
y
to Center
Processes and Minimize Variation
17
General Electric Aircraft Engines
General Electric Aircraft Engines
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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
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