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Advanced Low Conductivity Thermal Barrier Coatings: Performance and Future Directions (Invited paper)
Dongming Zhu and Robert A. Miller
NASA Glenn Research Center 21000 Brookpark Road, Cleveland, Ohio 44135
Thermal barrier coatings will be more aggressively designed to protect gas turbine
engine hot-section components in order to meet future engine higher fuel efficiency and lower emission goals. In this presentation, thermal barrier coating development considerations and performance will be emphasized. Advanced thermal barrier coatings have been developed using a multi-component defect clustering approach, and shown to have improved thermal stability and lower conductivity. The coating systems have been demonstrated for high temperature combustor applications. For thermal barrier coatings designed for turbine airfoil applications, further improved erosion and impact resistance are crucial for engine performance and durability. Erosion resistant thermal barrier coatings are being developed, with a current emphasis on the toughness improvements using a combined rare earth- and transition metal-oxide doping approach. The performance of the toughened thermal barrier coatings has been evaluated in burner rig and laser heat-flux rig simulated engine erosion and thermal gradient environments. The results have shown that the coating composition optimizations can effectively improve the erosion and impact resistance of the coating systems, while maintaining low thermal conductivity and cyclic durability. The erosion, impact and high heat-flux damage mechanisms of the thermal barrier coatings will also be described.
https://ntrs.nasa.gov/search.jsp?R=20100042394 2018-05-15T00:39:04+00:00Z
National Aeronautics and Space Administration
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Advanced Low Conductivity Thermal Barrier Coatings: Performance and Future Directions Advanced Low Conductivity Thermal Barrier Coatings: Performance and Future Directions
Dongming Zhu and Robert A. Miller
Durability and Protective Coatings Branch, Structures and Materials DivisionNASA John H. Glenn Research Center
Cleveland, Ohio 44135, USA
35th International Conference On Metallurgical Coatings And Thin Films (ICMCTF 2008)San Diego, California, April 27-May 2, 2008
Contact: Dr. Dongming Zhu(216) 433-5422
National Aeronautics and Space Administration
www.nasa.gov
Acknowledgments
This work was supported by NASA Fundamental Aeronautics (FA) Program Supersonics and Subsonic Rotary Wing Projects.
Howmet CoatingsHoneywell EnginesUCSBDirect Vapor Technol.
GE AviationPratt and WhitneyRolls Royce-Liberty WorksSUNY/Mesoscibe Tech.
Collaborators
National Aeronautics and Space Administration
www.nasa.gov
Motivation— Thermal barrier coatings (TBCs) can significantly increase gas
temperatures, reduce cooling requirements, and improve engine fuel efficiency and reliability
(a) Current TBCs (b) Advanced TBCs
Ceramic coating
Ceramic coating
SubstrateSubstrateBond coat
Bond coat
Tgas
Tsurface
Tgas
Tsurface
TbackTback
Ceramic coating
Ceramic coating
SubstrateSubstrateBond coat
Bond coat
Tgas
Tsurface
Tgas
Tsurface
TbackTback
Ceramic coating
Ceramic coating
SubstrateSubstrateBond coat
Bond coat
Tgas
Tsurface
Tgas
Tsurface
TbackTback
Turbine
Combustor Vane Blade Ceramic nozzles
CMC combustor liner and vane
Turbine
Combustor Vane Blade Ceramic nozzles
CMC combustor liner and vaneCMC combustor liner and vaneCMC combustor liner and vane
National Aeronautics and Space Administration
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NASA Ceramic Coating Development Goals— Meet engine temperature and performance requirements
- improved engine efficiency - reduced emission- increase long-term durability
— Improve technology readiness— The programs require a step-increase in coating capability— Reliability critical
2400°F (1316°C)
3000°F+ (1650°C+)
Gen I
Temperature Capability (T/EBC) surface
Gen II – Current commercialGen III
Gen. IV
Increase in ∆T across T/EBC
Single Crystal Superalloy
Year
Ceramic Matrix Composite
Gen I
Temperature Capability (T/EBC) surface
Gen II – Current commercialGen III
Gen. IV
Increase in ∆T across T/EBC
Single Crystal Superalloy
Year
Ceramic Matrix Composite
2700°F (1482°C)
2000°F (1093°C)
Step increase in temperature capability3000°F SiC/SiC CMC coatings2700°F SiC/SiC CMC and Si3N4
coatings
2800ºF combustor TBCs
2500ºF Turbine TBCs
National Aeronautics and Space Administration
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Outline─ Simulated high-heat-flux testing approaches
• Laser high heat flux• Burner and laser high temperature erosion• High pressure burner and high heat-flux capability
─ Low conductivity thermal barrier coating developments• Low conductivity TBC design requirements • Performance of low k four-component TBC systems
Conductivity, and cyclic durability• High toughness Low k four- and six-component turbine airfoil
TBC development – erosion resistance• CMAS interaction testing
─ Future directions
─ Summary
National Aeronautics and Space Administration
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High Heat-Flux Test Approaches– High-heat-flux tests crucial for turbine TBC developments
• CO2 laser simulated turbine engine high-heat-flux rig• Atmospheric burner rig simulated heat flux testing• High pressure burner rig simulated engine heat flux and pressure environments
High pressure burner rig combustion flow seen from viewport
TC
Atmospheric burner rig
High power CO2 Laser high heat flux rig
Tsurface2400-2500°F
(1316-1371°C)Tinterface
1950°F(1066°C)
- ∆T ~450°F (250°C) across 5mil coating- Heat flux up to 400 W/cm2
Turbine blade TBC testing requirements
T
Dis
tanc
e fr
om s
urfa
ce
Heat flux
cooling hc=0.4 W/cm2-K max
TBC
National Aeronautics and Space Administration
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High Velocity Burner Erosion Rig and Laser high Heat Flux Erosion Test Rig for Turbine TBC Testing
Erosion jet
High precision particle feeder system
Specimens under testing
Mach 0.3-1.0 burner erosion rig Laser heat flux erosion rig
National Aeronautics and Space Administration
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ZrO2-(7-8) wt%Y2O3Thermal Barrier Coating Systems
Relatively low intrinsic thermal conductivity ~2.5 W/m-K High thermal expansion to better match superalloy substrates Good high temperature stability and mechanical properties Additional conductivity reduction by micro-porosity
100 µm
Ceramic coating
Bond coat
(a) Plasma-sprayed coating
25 µm
Ceramic coating
Bond coat
(b) EB-PVD coating
National Aeronautics and Space Administration
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0.0
0.5
1.0
1.5
2.0
2.5
3.0
Plasma-sprayed TBC EB-PVD TBC
Ther
mal
con
duct
ivity
, W/m
-K
Coating Type
Conductivity reduction by microcracks and microporosity
— Significant conductivity increase at high temperature due to sintering— Accelerated failure due to phase stability and reduced strain tolerance
Intrinsic ZrO2-Y2O3
conductivity
As received conductivity(EB-PVD)As received conductivity(Plasma Coating )
20-hr rise at 1316°C
20-hr rise at 1316°C
20-hr rise at 1400°C
20-hr rise at 1371°C
Sintering and Conductivity Increase of ZrO2-(7-8) wt%Y2O3
National Aeronautics and Space Administration
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Sintering Kinetics of Plasma-Sprayed ZrO2-8wt%Y2O3Coatings
−−
−⋅=
−
−τt
RTkkkk
cc
cc exp168228exp2.1020inf
0
⋅=RT
41710exp5.572τ
Zhu & Miller, Surf. Coat. Technol., 1998; MRS Bulletin, 2000
National Aeronautics and Space Administration
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Sintering Cracks and Delaminations
— High heat flux surface sintering cracking and resulting coating delaminations
Tsurface=1280°CTinterface=1095°CThickness=130 µm
Zhu et al, Surf. Coat. Tech., 138 (2001), 1-8
surface vertical cracks
Delamination cracks
50 µm
National Aeronautics and Space Administration
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Sintering Cracks and Delaminations- continued
— Sintering strain corresponding to the thermal gradient across the coating (Tsurface=1280°C, Tinterface=1095°F)
0.00
0.50
1.00
1.50
2.00
0 50 100 150 200 250
0 hr46 hr200 hrmean strain
Surf
ace
open
ing
stra
in, %
Time, hours
strain (in%) = 0.40757 + 0.41 · t (in hr) 0.2
100 µm
National Aeronautics and Space Administration
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Low Conductivity and Sintering Resistant Thermal Barrier Coating Design Requirements
— Low conductivity (“1/2” of the baseline) retained at 2400°F— Improved sintering resistance and phase stability (up to 3000°F)— Excellent durability and mechanical properties
• Cyclic life• Toughness• Erosion/impact resistance• CMAS and corrosion resistance• Compatibility with the substrate/TGO
— Processing capability using existing infrastructure and alternative coating systems
— Other design considerations• Favorable optical properties• Potentially suitable for various metal and ceramic components • Affordable and safe
National Aeronautics and Space Administration
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Low Conductivity Thermal Barrier Coating Design Approaches
Efforts on modifying coating microstructures and porosity, composite TBCs, or alternative oxide compounds
Emphasize ZrO2- or HfO2-based alloy systems – defect cluster approach for toughness consideration
Advantages of defect cluster approach
• Advanced design approach: design of the clustering
• Better thermal stability: point defects are thermodynamically stable
• Improved sintering resistance: effective defect concentration reduced and activation energies increased by clustering
• Easy to fabricate: plasma-sprayed or EB-PVD processes
National Aeronautics and Space Administration
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Development of Advanced Defect Cluster Low Conductivity Thermal Barrier Coatings
— Multi-component oxide defect clustering approach (Zhu and Miller, US Patents No. 6,812,176, No.7,001,859, and No. 7,186,466)
— Defect clusters associated with dopant segregation— The nanometer sized clusters for reduced thermal conductivity, improved
stability, and mechanical properties
EELS elemental maps of EB-PVD ZrO2-(Y, Gd,Yb)2O3
Plasma-sprayed ZrO2-(Y, Nd,Yb)2O3
EB-PVD ZrO2-(Y, Nd,Yb)2O3
e.g.: ZrO2-Y2O3-Nd2O3(Gd2O3,Sm2O3)-Yb2O3(Sc2O3) systemsPrimary stabilizer
Oxide cluster dopants with distinctive ionic sizes
Zhu et al, Ceram. Eng. Sci. Proc., 2003
National Aeronautics and Space Administration
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Defect Clusters in a Plasma-Sprayed Y2O3, Nd2O3 and Yb2O3 Co-Doped ZrO2-Thermal Barrier Coating
— Yb, Nd rich regions consisting of small clusters with size of 5 to 20 nm
Energy (keV)
Cou
nts
20151050
1000
800
600
400
200
0
Cu
Cu
Cu
OAl
Mo
Mo
Zr
Zr
Y
Y
Yb
YbYb
Yb
YbNd
Nd
Nd
Nd
5 nm5 nm
Yb, Nd rich region clusters
Energy (keV)
Cou
nts
151050
1200
1000
800
600
400
200
0
Mo
Mo
Cu
Cu
Cu
Zr
Zr
Zr
ZrZr
Nd
NdNd
NdNd
Y
Y
Y
Y
ScSc
Yb
YbYb
Yb
Yb
O
Overall EDS
Yb rich region EDS
~40Å
(111)
~11.4Å
9.07Å
National Aeronautics and Space Administration
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Low Conductivity Defect Cluster Coatings Demonstrated Improved Thermal Stability
(a) Plasma-sprayed coatings
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
0 5 10 15 20 25
Ther
mal
con
duct
ivity
, W/m
-K
Time, hours
ZrO2-4.55mol%Y
2O
3 (ZrO
2-8wt%Y
2O
3 )
ZrO2-13.5mol%(Y,Nd,Yb)
2O
3
rate increase: 2.65×10-6 W/m-K-s
rate increase: 2.9×10-7 W/m-K-s
1316°C
0.0
0.5
1.0
1.5
2.0
2.5
0 5 10 15 20 25
ZrO2-4mol%Y
2O
3 (ZrO
2-7wt%Y
2O
3 )
ZrO2-4mol%Y2O
3 (ZrO
2-7wt%Y
2O
3 )
Low conductivity ZrO2-10mol%(Y,Gd,Yb)
2O
3 coating
Ther
mal
con
duct
ivity
, W/m
-K
Time, hours
rate increase: 2.2-3.8×10-6 W/m-K-s
rate increase: 6.0×10-7 W/m-K-s
1371°C
(b) EB-PVD coatings
— Thermal conductivity significantly reduced at high temperatures for the low conductivity TBCs
National Aeronautics and Space Administration
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Thermal Conductivity of Defect Cluster Thermal Barrier Coatings
(k0, k5 and k20 are the initial thermal conductivity, and the conductivity at 5 and20 hours, respectively)EB-PVD coatings
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
0 5 10 15 20 25 30
k2 0
-YSZ-Nd-Yb
k20
-YSZ-Gd-Yb
k20
-YSZ-Sm-Yb
k20
-YSZ
k5-YSZ-Gd-Yb
k5-YSZ-Gd-Yb-Sc
k5-YSZ-Nd-Yb-Sc
k0-4.55YSZ
k5-4.55YSZ
The
rmal
con
duct
ivity
, W/m
-K
Total dopant concentration, mol%
YSZ coatings4.55YSZ k
0
4.55YSZ k20
cluster oxide coatings
4.55YSZ k5
t’ phase region Cubic phase region
National Aeronautics and Space Administration
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Thermal Conductivity of Defect Cluster Thermal Barrier Coatings
— Thermal conductivity benefit of oxide defect cluster thermal barrier coatings demonstrated
Plasma-sprayed coatings(k0, and k20 are the initial thermal conductivity, and the conductivity at 5 and20 hours, respectively)
t’ phase region Cubic phase region
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
2 4 6 8 10 12 14
8YSZ k0 (2500F)8YSZ k20 (2500F)8YSZ k0 (2600F)8YSZ k20 (2600F)Refractron k0 (~2500F)Refractron k20 (~2500F)Refractron k0 (~2700F)Refractron k20 (~2700F)Praxair k0 (2500F)Praxair k20 (2500F)NASA k0 (2500F)NASA k20 (2500F)
Ther
mal
con
duct
ivity
, W/m
-K
Total dopant concentration, mol%
Plasma-sprayed coatings
AdvancedZrO2-based
coatings
k0k20k0k20k0k20k0k20
t’ phase region Cubic phase region
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
2 4 6 8 10 12 14
8YSZ k0 (2500F)8YSZ k20 (2500F)8YSZ k0 (2600F)8YSZ k20 (2600F)Refractron k0 (~2500F)Refractron k20 (~2500F)Refractron k0 (~2700F)Refractron k20 (~2700F)Praxair k0 (2500F)Praxair k20 (2500F)NASA k0 (2500F)NASA k20 (2500F)
Ther
mal
con
duct
ivity
, W/m
-K
Total dopant concentration, mol%
Plasma-sprayed coatings
AdvancedZrO2-based
coatings
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
2 4 6 8 10 12 14
8YSZ k0 (2500F)8YSZ k20 (2500F)8YSZ k0 (2600F)8YSZ k20 (2600F)Refractron k0 (~2500F)Refractron k20 (~2500F)Refractron k0 (~2700F)Refractron k20 (~2700F)Praxair k0 (2500F)Praxair k20 (2500F)NASA k0 (2500F)NASA k20 (2500F)
Ther
mal
con
duct
ivity
, W/m
-K
Total dopant concentration, mol%
Plasma-sprayed coatings
0.4
0.6
0.8
1.0
1.2
1.4
1.6
1.8
2.0
2 4 6 8 10 12 14
8YSZ k0 (2500F)8YSZ k20 (2500F)8YSZ k0 (2600F)8YSZ k20 (2600F)Refractron k0 (~2500F)Refractron k20 (~2500F)Refractron k0 (~2700F)Refractron k20 (~2700F)Praxair k0 (2500F)Praxair k20 (2500F)NASA k0 (2500F)NASA k20 (2500F)
Ther
mal
con
duct
ivity
, W/m
-K
Total dopant concentration, mol%
Plasma-sprayed coatings
AdvancedZrO2-based
coatings
k0k20k0k20k0k20k0k20
National Aeronautics and Space Administration
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Furnace Cyclic Behavior of ZrO2-(Y,Gd,Yb)2O3Thermal Barrier Coatings
― t’ low k TBCs had good cyclic durability― The cubic-phase low conductivity TBC durability needed improvements
Zhu and Miller, Ceram. Sci. Eng. Proc., 2002
EB-PVD
0
100
200
300
400
500
600
0 5 10 15 20 25 30 35
YSZ-Nd-YbYSZ-Gd-YbYSZ-Sm-YbYSZ
Cyc
les
to fa
ilure
Total dopant concentration, mol%
YSZ
National Aeronautics and Space Administration
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Furnace Cyclic Behavior of ZrO2-(Y,Gd,Yb)2O3Thermal Barrier Coatings - Continued
― t’ low k TBCs had good cyclic durability― The cubic-phase low conductivity TBC durability initially improved by an
7YSZ or low k t’-phase interlayer
0
200
400
600
800
1000
4 5 6 7 8 9 10 11 12
Cubic phase low k TBCTetrogonal t' phase low k TBCs8YSZ
Cyc
les t
o fa
ilure
Total dopant concentration, mol%
2075°F (1135°C)
Low-k t’-phase region
Low-k cubic phase region
With the interlayer
1135°C
Without interlayer
0
200
400
600
800
1000
4 5 6 7 8 9 10 11 12
Cubic phase low k TBCTetrogonal t' phase low k TBCs8YSZ
Cyc
les t
o fa
ilure
Total dopant concentration, mol%
2075°F (1135°C)
Low-k t’-phase region
Low-k cubic phase region
With the interlayer
1135°C
0
200
400
600
800
1000
4 5 6 7 8 9 10 11 12
Cubic phase low k TBCTetrogonal t' phase low k TBCs8YSZ
Cyc
les t
o fa
ilure
Total dopant concentration, mol%
2075°F (1135°C)
Low-k t’-phase region
Low-k cubic phase region
With the interlayer
0
200
400
600
800
1000
4 5 6 7 8 9 10 11 12
Cubic phase low k TBCTetrogonal t' phase low k TBCs8YSZ
Cyc
les t
o fa
ilure
Total dopant concentration, mol%
2075°F (1135°C)
Low-k t’-phase region
Low-k cubic phase region
With the interlayer
1135°C
Without interlayer
EB-PVD low k coatings
National Aeronautics and Space Administration
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Advanced Low Conductivity TBC Showed Excellent Cyclic Durability
― Coating validated for down-selected low conductivity coating systems
0.4
0.6
0.8
1.0
1.2
1.4
1.6
0 50 100 150 200 250 300 350
0 500 1000 1500 2000 2500
Ther
mal
con
duct
ivity
, W/m
-K
Time, hours
Cycle number
Tsurface=2480°F (1360°C)Tinterface=2020°F(1104°C)
6 min heating, 2 min cooling cycles
Low conductivity EB-PVD turbine airfoil coating
0.0
0.5
1.0
1.5
0 50 100 150 200 250
Ther
mal
con
duct
ivity
, W/m
-K
Time, hours
Tsurface=3000°F(1650°C)Tinterface=1670°F(910°C)
3030, 3min heating & 1.5min cooling cycles completed
Low conductivity plasma-sprayed combustor coating
Laser heat flux tests
Burner rig tests
National Aeronautics and Space Administration
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Advanced Low Conductivity Combustor Thermal Barrier Coating Developments
― Low k TBC coated components demonstrated in simulated engine environments
― Low k TBC being incorporated in advanced engine development programs
Low conductivity TBC combustor liner demonstration in Combustor rig
Low conductivity TBC Propulsion 21 flame tube and deflector
demonstrations
Low conductivity TBC flame tube and combustor deflector demos in Advanced Subsonic Combustion Rig (ASCR)
Flame tube
Low conductivity TBC: combustor liner demonstration
Outer liner
Inner liner
National Aeronautics and Space Administration
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Erosion and Impact Resistant Turbine TBC Development
— Multi-component ZrO2 low k coatings showed promise in improving erosion and impact resistance
0
50
100
150
200
250
300
350
ErosionImpact
Ero
sion
and
impa
ct re
sist
ance
(spe
cific
ero
dent
wei
ght r
equi
red
to p
enet
rate
coa
ting)
, m
g/pe
r mil
coat
ing
thic
knes
s
Coating type
ZrO2-7wt%Y
2O
3Cubic-phase
multi-component coating
High toughness, tetragonal multi-component coating
EB-PVD coatings Erosion and impact resistance, measured as the erodent Al2O3 weight required to penetrate unit thickness coating
2200°F burner rig erosionZhu & Miller, NASA R&T, 2004
National Aeronautics and Space Administration
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Advanced Multi-Component Erosion Resistant Turbine Blade Thermal Barrier Coating Development
─ Rare earth (RE) and transition metal oxide defect clustering approach (US Patents No. 6,812,176, No.7,001,859, and 7,186,466; US patent application 11/510,574 ) specifically by additions of RE2O3 , TiO2 and Ta2O5
─ Significantly improved toughness, cyclic durability and erosion resistance while maintaining low thermal conductivity
─ Improved thermal stability due to reduced diffusion at high temperature
ZrO2-Y2O3- RE1 {e.g.,Gd2O3,Sm2O3}-RE2 {e.g.,Yb2O3,Sc2O3} – TT{TiO2+Ta2O5} systemsPrimary stabilizer
Oxide cluster dopants with distinctive ionic sizesToughening dopants
National Aeronautics and Space Administration
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Furnace Cyclic Test Lifetime and Thermal Conductivity of TiO2 Doped Thermal Barrier Coatings
─ Unpublished work 2003
0
100
200
300
400
500
0.0
0.5
1.0
1.5
2.0
2.5
3.0
Cubic low k 433 FCT lifeCubic 433+10TiO
2 low k FCT life
Cubic low k 433 low k thermal conductivityCubic 433+10TiO
2 low k thermal conductivity
FCT
cycl
es to
failu
re, h
r
Ther
mal
con
duct
ivity
, W/m
-K
Cub
ic 4
-com
pone
nt
Cub
ic 5
-com
pone
nt (R
e 2O3+T
iO2)
Cub
ic 4
-com
pone
nt
Cub
ic 5
-com
pone
nt (R
e 2O3+T
iO2)
EB-PVD
0
100
200
300
400
500
600
0 5 10 15 20 25 30 35
YSZ-Nd-YbYSZ-Gd-YbYSZ-Sm-YbYSZ
Cyc
les
to fa
ilure
Total dopant concentration, mol%
YSZ
National Aeronautics and Space Administration
www.nasa.gov
Furnace Cyclic Lifetime of Advanced Turbine Thermal Barrier Coatings
─ Furnace cyclic life can be optimized with RE2O3 and TT additions─ Stability and volatility with too high TT concentrations
0
100
200
300
400
500
2 4 6 8 10 12 14
Four-component low k7YSZ
Cyc
les t
o fa
ilure
, hr
Totoal RE2O
3 dopant concentration
2125°F (1163°C)
Six-componnet low k (with TT addtions)
TT concentration
National Aeronautics and Space Administration
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Cyclic Life of Four-Component Thermal Barrier Coatings
3.5
4.0
4.5
5.0
5.5
6.0
6.5
7.0
0.00069 0.00070 0.00071 0.00072 0.00073 0.00074 0.00075
Zr2.5Y0.75Gd0.75Yb (ln)Zr2.0Y1.5Gd1.5Yb (ln)Zr1.6Y1.2Gd1.2Yb (ln)Zr2.5Y0.75Gd0.75Yb (Ln)Zr2.0Y1.5Gd1.5Yb (Ln)Zr1.6Y1.2Gd1.2Yb (Ln)Zr2.5Y0.75Gd0.75Yb (2h) (Ln)Zr2.0Y1.5Gd1.5Yb (2h) (Ln)Zr1.6Y1.2Gd1.2Yb (2h) (Ln)7YSZ (ln)7YSZ (Ln)7YSZ (2h) (Ln)7YSZ7YSZ Laser heat flux7YSZ burner rig
ln (C
ycle
tim
e to
failu
re),
hour
s
1/T, 1/K
1436 1423 1408 1338
Temperature, K
─ Furnace and high heat flux cyclic life being optimized for long-term durability
National Aeronautics and Space Administration
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Thermal Conductivity of Selected Low k Thermal Barrier Coatings
0.0
0.5
1.0
1.5
2.0
2.5
4.5 5.0 5.5 6.0 6.5 7.0
ZrO2-8wt%Y
2O
3 k20
ZrO2-8wt%Y
2O
3 anti-sin tering k20
Refarctron k0Refactron k20Praxair k0Praxair k20NASA k0NASA k20PVD-ZrO
2-7wt%Y
2O
3 k20
PVD t' low k k20PVD cubic low k k20
Ther
mal
con
duct
ivity
, W/m
-K
1/T ·104, K -1
120013001500 14001600Tem perature, °C
Plasma-sprayed and EB-PVD coatings
AdvancedLow k TBCs
National Aeronautics and Space Administration
www.nasa.gov
Impact Resistance of Advanced Multi-component Low Conductivity Thermal Barrier Coatings
― Improved impact/erosion resistance observed for advanced low conductivity six-component coatings
0.0
0.5
1.0
1.5
2.0
2.5Zr
YG
dYb
t' -1
Zr
YG
dYb
t' -2
Zr
YG
dYb
t' -3
Zr
YG
dYb
t' -4
Zr
YG
dYb
Cub
ic m
ultil
ayer
Zr
YG
dYb
Cub
ic +
TT
Zr
O2-4
Y (b
asel
ine)
Advanced coatingsBaseline
Eros
ion
resi
stan
ce(E
roda
nt a
mou
nt re
quire
dpe
r mg
coat
ing
rem
oval
, g/m
g) 2200°F
>100%Improvements
650 micron Al2O3
National Aeronautics and Space Administration
www.nasa.gov
Erosion Resistance of Advanced Multi-component Low Conductivity Thermal Barrier Coatings
50 µm Al2O3
100% increase
0
5
10
15
20
25
30
ZrY
GdY
b t'
ZrY
GdY
b t'+
TT
ZrY
GdY
b C
ubic
+TT
ZrY
GdY
b C
ubic
ZrO
2-4Y
(bas
elin
e)
Advanced coatingsBaseline
Eros
ion
resi
stan
ce(E
roda
nt a
mou
nt re
quire
dpe
r mil
coat
ing
rem
oval
, g/m
il) 2175°F
─ The original cubic low k coating showed significant increase in erosion resistance due to the incorporation of TiO2 and Ta2O5
National Aeronautics and Space Administration
www.nasa.gov
Tetragonality of Multi-Component ZrO2 being Evaluated and Correlated to Coating Performance
Area detector x-ray diffractometer used for EB-PVD coatings
1 7 F 1 A 5 p o w d e r - In s tru m e n t C o m p a r is o n
A r e a D e te c to r P V DA r e a D e te c to r P o w d e r
Sqr
t (C
ps)
1
1 0
1 00
1 00 0
2 00
3 00
4 00
5 00
6 00
2 00 0
2 - T h e ta - S c a le7 2 . 5 73 7 4 7 5
─ Multi-component TiO2/Ta2O5 and rare earth dopants increase the tetragonality (c/a ratio)
─ Current efforts in optimizing the dopant composition ranges
1.000
1.010
1.020
1.030
1.040
1.050
2 4 6 8 10 12
7YSZZrO
2-MultiRE doped
ZrO2-RETiTa doped
c/a
ratio
RE dopant concentration
Higher TT dopant conc.
lower TT dopant conc.
National Aeronautics and Space Administration
www.nasa.gov
Impact Failure of Advanced Multi-Component Low Conductivity Thermal Barrier Coatings
30 um30 um
Surface sintering and impact densification zone
SEM micrographs of advanced thermal barrier coating after impact/erosion damage
Backscattered electron imageSecondary electron image
― Surface sintering and impact densification zones observed, with subsequent spallation under the erodent further impacts
― Toughened structures observed
National Aeronautics and Space Administration
www.nasa.gov
Impact Failure of Advanced Multi-Component Low Conductivity Thermal Barrier Coatings
Coating 1Coating 2
Eros
ion
rate
Time
Near failure
Initial rate
desified region
steady-state
fast failure region
30 um
─ Effect of erosion parameters will be modeled and validated
National Aeronautics and Space Administration
www.nasa.gov
High Heat Flux Testing of Turbine EB-PVD Thermal Barrier Coatings to Study CMAS Effect
─ Specimens typically tested at Tsurface ~2400ºF, Tinterface 2000°F─ Heat flux up to 250-300 W/cm2, cooling heat transfer coefficient up to hc
0.32 W/cm2-K─ Accelerated failure observed with CMAS interactions ─ Advanced multi-component coatings completed 50 hr testing
Specimen under the rig test
(a) Upon initial heating
(b) After testing
Combustor TBC
Baseline coating
Turbine TBCs
Advanced coatings50 hrs
National Aeronautics and Space Administration
www.nasa.gov
Future Directions for Low Conductivity TBC Development
— Emphasize high heat flux durability and erosion resistance
- Optimize high toughness erosion resistant turbine coatings
- Improve turbine airfoil TBCs with up to 3x erosion resistance
- Emphasize creep, fatigue, erosion, and CMAS interactions
- Develop multilayered damping and erosion coatings
- Develop turbine blade TBC life prediction model
National Aeronautics and Space Administration
www.nasa.gov
High temperature capability thermal and radiation barrier
Energy dissipation and chemical barrier interlayer
Environmental barrier
Advanced bond coat
Ceramic matrix composite (CMC)
Tsurface >1482°C (2700°F)
Tcoating/CMC interface <1316°C (2400°F)
High temperature capability thermal and radiation barrier
Energy dissipation and chemical barrier interlayer
Environmental barrier
Advanced bond coat
Ceramic matrix composite (CMC)
High temperature capability thermal and radiation barrier
Energy dissipation and chemical barrier interlayer
Environmental barrier
Advanced bond coat
Ceramic matrix composite (CMC)
Tsurface >1482°C (2700°F)
Tcoating/CMC interface <1316°C (2400°F)
Tsurface >1482°C (2700°F)
Tcoating/CMC interface <1316°C (2400°F)
127 µm(0.005”)
CMC Turbine Blade coatings
Future Directions for Low Conductivity TBC Development
CMC combustor liner and vane
— Emphasize thin ceramic matrix composite turbine coating processing- Advanced processing for integrated TEBCs
- Ceramic nanocomposite and nanotube-based TEBCs for improved durability and optical properties
- Embedded sensors
- Life prediction methodology and design tool development
National Aeronautics and Space Administration
www.nasa.gov
Summary
• Four-component low k TBC systems developed for low k combustor applications
• Advanced turbine airfoil TBCs being developed with combined low conductivity and high toughness
• Improved erosion/impact resistance observed for the multi-component coating t’ and t’/cubic nano-composite systems
• Coatings being optimized for cyclic life, thermal conductivity and erosion/impact and CMAS resistance
• High heat flux durability, multifunctional coatings and lifingmodels being emphasized in the current research programs