advanced core engine technology · 2015. 1. 30. · v2500 agtj100b cfm56-5 501f 9001g,h 7001g,h...
TRANSCRIPT
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Advanced Core Engine Technology
Dr. James D. Heidmann
Chief, Turbomachinery & Heat Transfer Branch
NASA Glenn Research Center
Green Aviation Summit
September 9, 2010
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NASA’s Subsonic Transport System Level Metrics…. technology for dramatically improving noise, emissions, & performance
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CORNERS OF
THE
TRADE SPACE
N+1 = 2015
Technology Benefits Relative
To a Single Aisle Reference
Configuration
N+2 = 2020
Technology Benefits Relative
To a Large Twin Aisle
Reference Configuration
N+3 = 2025
Technology Benefits
Noise
(cum below Stage -32 dB -42 dB -71 dB
4)LTO NO Emissionsx
(below CAEP 6)-60% -75% better than -75%
Performance:
Aircraft Fuel Burn-33% -50%** better than -70%
Performance:
Field Length-33% -50% exploit metro-plex* concepts
Goals are relative to reaching TRL 6 by the timeframe indicated
Better chance to reach all goals simultaneously for N+2 and N+3
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POTENTIAL REDUCTION IN FUEL CONSUMPTIONAdvanced N+2 Configurations
Advanced Configuration #1N+2 “tube-and-wing“
2025 EIS (TRL=6 in 2020)
Advanced EnginesΔ Fuel Burn = -15.3%
-120,300 lbs
(-43.0%)
Fuel Burn = 159,500 lbs
-43.0%
Advanced Configuration #2AN+2 HWB300
2025 EIS (TRL=6 in 2020)
Advanced EnginesΔ Fuel Burn = -18.5%
-139,400 lbs(-49.8%)
Fuel Burn = 140,400 lbs
-49.8%
Advanced Configuration #2BN+2 HWB300
2025 EIS (TRL=6 in 2020 assumingaccelerated technology development)
Advanced EnginesΔ Fuel Burn = -16.0%
-151,300 lbs(-54.1%)
Embedded Engines withBLI Inlets ∆ Fuel Burn = -3.2%
Fuel Burn = 128,500 lbs
-54.1%
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Propulsion Technology Enablers
Fuel Burn
• reduced SFC (increased BPR, OPR & turbine inlet
temperature, potential embedding benefit)
Velocity
TSFC
Lift
Dragln
Wfuel
WPL + WO=
•Aerodynamics • Empty Weight •Engine Fuel
Consumption
Aircraft
Range1 +
Velocity
TSFC
Lift
Dragln
Wfuel
WPL + WO=
•Aerodynamics • Empty Weight •Engine Fuel
Consumption
Aircraft
Range1 +
TSFC = Velocity / (ηoverall)(fuel energy per unit mass)
ηoverall = (ηthermal)(ηpropulsive)(ηtransmission)(ηcombustion)
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FY10 SIX MONTH REVIEW – INTEGRATED SYSTEM RESEAERCH
Click to edit Master title style
• Click to edit Master text styles
– Second level
• Third level
– Fourth level
» Fifth level
Propulsion system improvements require advances in propulsor
and core/combustor technologies
Propulsion Technology Opportunity
Core
Improvements
(direct impact on LTO NOx)
Propulsor
improvements
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Cycle Performance Improves with Temperature
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Engine Thermal Trends
500
600
700
800
900
1000
1100
1200
1300
1400
1500
1600
1700
1800
1930 1940 1950 1960 1970 1980 1990 2000
AERO-ENGINE
INDUSTRIAL GT
Temp. at Rotor Inlet
CFM56
E3
701F
TRENT
CF6-80
F100
F101
XF3-20 F3-30
F110
MF111
F404
PW2037
H-25 7001F
9001F TEPCO
V2500
AGTJ100BCFM56-5
501F
9001G,H
7001G,H
F100-PW-299
XG40
M-88HYPRA
HYPRB
PW4000GE90 701G
501G
CGTAGTJ100A
JR220
TF39
BBC4000kW1GO
BBC900kWKO-7
HeS3B NE-20
BMW003
W2/700
J69
J73
Dart10
Conway
T56-14
AvonJT3D
CT610 J3-7
T64
J79-17
JT8D
JR100
501AA
7001B
JT15D
TFE731
9001B
TF40
TF30
CF6-6CF6-50
RB199
FJR710/600
FJR710/20
FJR710-10
JT9D-3
RB211-22
7001E 9001E
701D
JT9D-7R4
501D
501B
W1
JUMO004-1J3
YEAR
TU
RB
INE
IN
LE
T
TE
MP
ER
AT
UR
E [℃
]
F.WhittlePAT
From Dr. Toyoaki Yoshida, National Aerospace Laboratory, Japan
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Materials and Cooling Improvements
Increase in operational temperature of
turbine components. After Schulz et al,
Aero. Sci. Techn.7:2003, p73-80.
approximately two-thirds of turbine temperature increase
has historically been enabled by cooling improvements
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Turbomachinery Aero Design-Based Tech Enablers
Highly-Loaded,
Multistage
Compressor
(higher efficiency
and operability)
Low-Shock
Design, High
Efficiency,
High Pressure
Turbine
Aspiration Flow
Controlled,
Highly-Loaded,
Low Pressure
Turbine
Low Pressure
Turbine
Plasma Flow
Control
FLOW
z
FLOW
z
UW
FLOW
UW
FLOW
Novel Turbine Cooling
Concepts
Compressor
Synthetic
Jet Flow
Control
(reduced
flow losses)
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Multi-Stage Axial Compressor (W7)Objective: To produce benchmark
quality validation test data on a state-of-
the-art multi-stage axial compressor
featuring swept axial rotors and stators.
The test will provide improved
understanding of issues relative to optimal
matching of highly loaded compressor
blade rows to achieve high efficiency and
surge margin.
ERB Test Cell W7
76B 3-Stage Axial Compressor
PR=4.7,TR=1.63, ha=88.5, Utip =1400 ft/s
Approach: Acquire and test a modern
axial compressor representative of the
first three stages of a commercial
engine high pressure compressor with
an engine company.
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UTRC NRA – High Efficiency Centrifugal Compressor (HECC)
m = 10.1 lbm/s
Opportunity for
improved rotary wing
vehicle engine
performance as well
as rear stages for
high OPR fixed wing
application
CFD prediction apply delta, CFD
to data
increase inlet
radius ratio
scale from rig to
engine
TT
po
lytr
op
ic e
ffic
ien
cy
Opportunity for
improved rotary wing
vehicle engine
performance as well
as rear stages for
high OPR fixed wing
application
86.0%
86.5%
87.0%
87.5%
88.0%
88.5%
89.0%
89.5%
90.0%
CFD prediction apply delta, CFD
to data
increase inlet
radius ratio
scale from rig to
engine
TT
po
lytr
op
ic e
ffic
ien
cy Engine
scale
polytropic
efficiency
is
estimated
as 87.9 -
88.9%
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Turbine Film Cooling Experiments
Objective: Fundamental study of heat transfer and
flow field of film cooled turbine components
Rationale: Investigate surface and flow interactions
between film cooling and core flow for various large
scale turbine vane models
Approach: Obtain detailed flow field and heat
transfer data and compare with CFD simulations
Trailing
Edge Film
Ejection:
IR images
Vane Heat Transfer:
Good agreement between
GlennHT and experiment
Large Scale Film Hole:
Film cooling jet
downstream of hole
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NASA/General Electric Highly-Loaded Turbine Tests
Conventional HPT Reduced Shock Design
Pressure Ratio (PTR/PS) = 3.25
Stage Pressure Ratio = 5.5
FLOW
FLO
W
COFFING
4 TON
HOIST
HPT: Reduced Shock Design
LPT: Flow-Controlled Stator
& Contoured Endwall
High and Low Pressure Turbine Testing in NASA
Glenn Single Spool Turbine Facility (W6)
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Anti-Vortex Film Cooling Concept
Top View
Flow Direction
Front View
Side View
Auxiliary holes (yellow) produce counter-vorticity to promote jet attachment
Advantages: Inexpensive due to use of only round holes, hole inlet area unchanged
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CMC Turbine Vanes
Objective: Reduce emissions in high bypass ratio turbine engines by utilizing
high temperature Ceramic Matrix Composites (CMCs) for high pressure
turbine vanes applications
Approach: Demonstrate fabricability of CMCs in vane shape using two or
more processing methods and evaluate current state of the art materials in
a relevant environment
Proposed Vane Shape
for Evaluations
SiC/SiC – Hi-Nic type S fibers
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Variable-Speed Power Turbine for Large Civil Tilt-Rotor
NASA FAP/SRW Large Civil Tilt-Rotor
Notional Vehicle
Need for variable-speed power-turbine with
range of operation 54% < NPT < 100%
Research addresses key technical challenges:
High efficiency at high stage work factor
Wide (50⁰) incidence range requirementLow Reynolds number operation
Large Civil Tilt-Rotor
TOGW 108k lbf
Payload 90 PAX
Engines 4 x 7500 SHP
Range > 1,000 nm
Cruise speed > 300 kn
Cruise altitude 28 – 30 kft
–
–
•
•
•
Goal – to enhance airport throughput
capacity by utilizing VTOL capable
rotary wing vehicles.
Principal challenge for LCTR –
required variability in main-rotor
speed:
650 ft/s at take-off
350 ft/s at Mn 0.5 cruise
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Core Engine Research Summary
Turbomachinery research directly impacts fuel burn reduction goals of NASA
Aeronautics projects
Compressor research focused on increasing overall pressure ratio while
maintaining or improving aerodynamic efficiency
Centrifugal compressor research can meet multiple project architecture needs
Turbine research focused on increased loading, reduced cooling flows, and
improved aerodynamic efficiency
Fundamental testing of turbine cooling flows to improve effectiveness
Low pressure turbine flow control and rotating testing
Computational fluid dynamic development and assessment across all
components, including advanced turbulence models such as LES and DNS
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