Thermal barrier coatings (TBCs) for turbine airfoils in high-performance engines represent an advanced materials technology with both performance and durability benefits. The foremost TBC benefit is the reduction of heat transferred into air-cooled components, which yields performance and durability benefits. This program focuses on predicting the lives of two types of strain-tolerant and oxidation-resistant TBC systems that are produced by commercial coating suppliers to the gas turbine industry. The plasma-sprayed TBC system, composed of a low-pressure plasma-spray (LPPS) or an argon shrouded plasma-spray (ASPS) applied oxidation resistant NiCrAlY (or CoNiCrAlY) bond coating and an air-plasma-sprayed yttria (8 percent) partially stabilized zirconia insulative layer, is applied by Chromalloy, Klock, and Union Carbide. The second type of TBC is applied by the electron beam-physical vapor deposition (EB-PVD) process by Temescal

Liu, A.

Neumann, J. F.

Strangman, T. E.

English

NASA Technical Reports Server

 8 -!29 0
THERMAL BARRIER COATING LIFE
PREDICTION MODEL DEVELOPMENT*
T.E. Strangman, J.F. Neumann and A. Liu
Garrett Turbine Engine Company (GTEC)
Thermal barrier coatings (TBCs) for turbine airfoils in high-performance
engines represent an advanced materials technology with both performance and
durability benefits. The foremost TBC benefit is the reduction of heat
transferred into air-cooled components, which yields performance and dura-
bility benefits (Figure i). To achieve these benefits, however, the TBC sys-
tem must be reliable. Mechanistic thermomechanical and thermochemical life
models are therefore required for the reliable exploitation of TBC benefits
on gas turbine airfoils. GTEC's NASA-HOST Program (NAS3-23945) goal is to
fulfill these requirements.
This program focuses on predicting the lives of two types of strain-
tolerant and oxidation-resistant TBC systems that are produced by commercial
coating suppliers to the gas turbine industry (Figure 2). The plasma-sprayed
TBC system, composed of a low-pressure plasma-spray (LPPS) or an argon
shrouded plasma-spray (ASPS) applied oxidation resistant NiCrAIY (or
CoNiCrAIY) bond coating and an air-plasma-sprayed yttria (8 percent)
partially stabilized zirconia insulative layer, is applied by Chromalloy
(Orangeburg, New York), Klock (Manchester, Connecticut), and Union Carbide
(Indianapolis, Indiana). The second type of TBC is applied by the electron
beam-physical vapor deposition (EB-PVD) process by Temescal (Berkeley,
California).
The overall objective of Phase I of this program is to develop mechanis-
tic mission-analysis-capable life prediction models for the predominant
environmental and thermomechanical TBC failure modes for preliminary design
analyses. Because the THC must be considered early in the component design
process in order to fully incorporate and exploit its benefits, an additional
model goal is to drive the preliminary TBC life model with component thermal
analysis data and simple snap acceleration-snap deceleration stress analysis
data. This approach permits the designer to economically include TBCs into
initial iterations of the blade and vane design process. More refined TBC
analyses for final design lives are the subject of Phase II.
A .comprehensive strategy to achieve this goal has been developed that
includes the analyses of the THC durability on both the TFE731-5 HP turbine
blades (Figure 3) and the burner rig test specimens. Due to the complex
nature of the problems involved, the finite element method is deemed to be
most effective in promoting the in-depth understanding of the essential over-
all thermal/mechanical behavior of the TBC systems as well as the inter-
actions between the individual material regimes and the interfacial condi-
tions in the TBC systems. This approach also interfaces efficiently with
existing airfoil design methods.
*Work done under NASA Contract NAS3-23945.
435
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'5'o ....
COOLINGAIR REQUIREMENT,PERCENT
ORIGINAL PAGE IS
OF POOR QUALITY
6_1_1
Figure I. TBCs Improve Creep Life and Reduce Cooling Air
Requirements for the GTEC High-Pressure Turbine
Blade.
TBC
80NO COAT
SUBSTRATE
PLASMA SPRAY
APS Y203 (8%l
STABILIZED ZrO2
LPPSNi-31Cr-11A1-O.5Y
MAR-M 247
SUPERALLOY
SUPPLIER • CHROMALLOY
• KLOCK
PLASMA SPRAY
APS Y203 (8%]
STABILIZED ZrO2
ASPS Co-32Ni-21Cr-
8Al-O.5Y
MAR-M 247
SUPERALLOY
• UNION CARBIDE
ELECTRONBEAM --
PHYSICAL VAPOR DEPOSITION
EB-PVD Y203 (20%)
STABILIZED ZrO2
EB-PVD
Ni-23Co.18Cr-12AI-O.3'f
MAR.M 247
SUPERALLOY
• TEMESCAL
05-195-2
Figure 2. Life Prediction Models are Being Developed for
Plasma-Sprayed and EB-PVD TBC Systems.
436
ENGINEDATA
• MISSION CYCLE
• TEST DATA
REFINED3D F.E.
SUB-MODEL
• THERMAL
• STRESS/STRAIN
• TOUGHNESS
G6_5_
COARSE30 F.E. MODEL
• THERMAL
• STRESS
• SNAP-CYCLE
STRESS/STRAIN
• IDENTIFYCRITICAL
LOCATIONS
I
TBC LIFE ANALYSIS
._ MATERIALS
PROPERTY
DATA
• SPALLINGSTRESS(KIc, T, _ ENGINEDATA
• OXIDATION(T, t) CORRELATION
• SALT FILM (T, t, ALT.)
B
• LINEARDAMAGE ___
Figure 3. Three TBC Damage Modes are Evaluated in TFE731
Blade Analysis.
TDC
Figure 4. TFE731 HP Turbine Blade F.E. Model Incorporates
Bond Coating and Zirconia Layers.
437
For computational efficiencies, typical preliminary design (PD) finite
element models were first constructed to analyze the bulk behavior of the TBC
systems on the airfoil component (Figure 4). Critical locations, in terms of
temperatures, stresses, and strains or their combinations, can be identified
from these PD models. Refined sub-models are then constructed for analysis
of critical locations. Detailed thermomechanical and thermochemical behavi-
ors of the TBC systems and the interactions between the individual material
regimes and the interfaclal conditions are being analyzed via these refined
sub-models. Major analysis work in this program is being performed with
ANSYS, a commercially available general purpose finite element code.
For preliminary component design analyses (Phase I), TBC lives are being
independently calculated for three operative damage modes:
o
o
o
Bond coating oxidation
Molten salt film damage, and
Thermomechanical stress induced spalling
Computation of the oxidation life of a TBC system in the preliminary
design model is driven by the component thermal analysis and engine power
requirements during a mission cycle.
Molten salt film damage life is calculated using the component thermal
analysis, engine power requirements during a mission, and aircraft altitude
(salt ingestion).
Zirconia spalling associated with thermomechanical stresses is calcu-
lated based on the analysis of the snap-cycle thermal transients as well as
the steady-state condition and the rotational loads in a mission cycle. Cal-
culated snap-cycle interracial tensile stresses and the largest pre-existing
flaw diameter (determined by NDE or calculated from bond strength tests) are
used to estimate a stress intensity factor that the coating must endure with-
out spalling. As indicated in subsequent paragraphs, the fracture toughness
of the zirconia or the bond coating-zirconia interface is dependent upon
exposure temperature and time. Time and temperature dependent changes in the
zirconia or interracial toughness are calculated based on the thermal analy-
sis results of the component and a linear cumulative damage model to account
for variations in a mission cycle.
Figure 5 is a schematic of the TBC life model that illustrates these
three failure modes and the respective temperatures regimes at which these
failure modes are likely to occur. Figure 6 illustrates parameters that
affect each of these three major failure modes.
The preliminary design TBC life is assessed via a linear damage rule,
composed of damages from these three modes during each of the mission cycles,
assuming no interactions between these failure modes; that is
Life = [(Lifeoxid) -I + (Lifesalt)-I + (Lifestress)-l] -I
438
TBC LIFE MODEL
D
,,--I
0
,,--,J
G5-195-5
SALT
SOLIDIFICATION
i/ "% EVAPORATION
: % % \ BOND
: %\ COATING
! MOLTI:,,- "_% OXIDATION
ZIRCONIA .,_
TOUGHNESS"'_ _
REDUCTION %
TEMPERATURE
Figure 5. TBC Life Model Has Three Failure Modes.
TBC
DEGRADATION
RATE
: F1 (MECHANICAL)
• COATINGSTRESSES
• TEMPERATURE
• MATERIALSYSTEM
• KIC
• FLAW SIZE
• ELASTICMODULUS
• SPALLINGSTRAIN
+ F2 (OXIDATION)
• TEMPERATURE
• CYCLESEGMENTLENGTH
• MATERIALSSYSTEM
+ F_ (SALT DEPOSITION)
• ALTITUDE(SALT INGESTION)
• TURBINEPRESSURE
• SALT EVAPORATION
• SALT SOLIDIFICATION
• TEMPERATURE
• GAS VELOCITY
• AIRCRAFTLOCATION
• MATERIALSSYSTEM
G5-195.3
Figure 6. TBC Life is a Function of Engine, Mission, and
Materials System Parameters.
439
However, the refined sub-models are being constructed to be sufficiently
flexible and detailed to analyze the interactions between these models. Sub-
sequent improvements in Phase II of the program will incorporate failure mech-
anism interactions into the life prediction model.
Burner rig and mechanical property data have been obtained to quantify
the capabilities of each of the TBC systems for each major mode of degrada-
tion. Burner rig test data are illustrated in Figures 7 and 8 for plasma-
sprayed and EB-PVD TBC coating systems. These data indicate that bond coat-
ing oxidation, high temperature zirconia densificiation (sintering), and
molten salt film damage at intermediate temperatures significantly affect TBC
life (Figure 7). The length of the heating cycle must also be considered
when computing a coating life (Figure 8).
Cohesive and interfacial toughness data have also been measured for
plasma-sprayed and EB-PVD TBC systems (Figure 9). It has been observed for
both types of coating systems that toughness is reduced by exposure at high
temperatures. A step transition in toughness, which is associated with sin-
tering shrinkage, is illustrated for plasma-sprayed TBC systems as a function
of exposure time at ll00C in Figure i0. The transition to lower toughness
levels correlates well with high temperature burner rig test data, as indi-
cated in Figure 7.
Lives of these TBC systems are being predicted for TFE731 high-pressure
turbine blades for factory engine test conditions, as well as business air-
craft mission. Thermal analysis of the turbine airfoil (Figure Ii) indicates
that the bond coating oxidation degradation mode results in minimum predicted
lives of approximately 7300 hours with the plasma-sprayed TBC system for the
business aircraft missions and I000 hours for the factory engine test condi-
tions. Lives for other failure modes for the plasma-sprayed as well as the
EB-PVD TBC systems are currently being analyzed.
This program is now in the third year of Phase I.
is provided in Figure 12.
The program schedule
440
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100,
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I
700
ORIGINAL PAGE tS
,_ OF pOOP, QUALITY ,o.00o.
%
%
%
%
%
%
%
MOLTER / L._
SALT FILM BeND COAT _'_y%
-O_,mAG_ OXIOATION _N _ ,00
i
A CRROMALLO_ / _
• KLOCK _ /" _ ,0,
ZIOCONIA
• uNioNCA.,DE
[] AVERAGETOUGHNESS AND BOND COAT
TRANSITIONTIME OXIOATIOR • w
%
%
%
%
MOLTEN
SALT FILM
DAMAGE
•TEMESCAL
• _NO COAT
INTERFACE
DEGRADATION
AND ZIDCONIA
OENSIFICATION
I , , I , I ' I l I 1
000 006 1000 1100 1200 ,3IX) 700 900 000 1000 ,100
TEMPERATURE.C TEMPERATURE.C
Figure 7. Three Degradation Modes Affect the Durability of
Plasma-Sprayed and EB-PVD TBC Systems.
| I
1200 1300
1000
J
© = 1020C TEMESCAL DATA
,_ = 1_C KLOCK DATA
• = IO_C UNION CARBIDE DATA
• = IO_C CHROMALLOY DATA
• = I1_C NASA DATA
! ; i WlV.,_ I l i itlll| i i | I ii|i I | i I I I,II
IO 100 I000 IO.OOG
HEATING CYCLE LENGTH. MIN
Figure 8. TBC Life Decreases with a Shorter Heating Cycle.
441
IMBEDDED
CONTAMINANT
_TT..;d..._FILM DOT
ZIRC0a IA--,=I==__-=-_-.41
J_E=Z NiCrAIY_,[ I
MAR-M
247
t
0
COHESIVE(INTERFACIAL]TOUGHNESSTEST
KIc - 2/vF_ orF v/C/-2
ZIRCONIA
FRACTURESURFACE
G5-0070-5
FIGURE 38
Figure 9. Cohesive and Interracial Toughness are Determined
with Modified Bond Strength Test.
ORIGINAL PAGE IS
OF_ pOOR QUALITY
442
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• UNION CARBIDEAFTER PCHT OF !1_C
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I
I o
I
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©
0
-2o : 0.41
u u i i i _m m ,it
1000
Figure i0. Fracture Toughness of Plasma-Sprayed Yttria (8 percent)
Stabilized Zirconia is Reduced After Long Exposures at
IIOOC.
443
ORIGINAL PAGE IS
OF POOR QUAI.n'Y
_1o
/
/
G6-IOG.152
Figure ii. Thermal Analysis of TBC-Coated Blade has been Conducted.
444
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OF POOR QUALITY
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DESIGN CAPABLE LIFE MOOELS
Figure 12. TBC Life Prediction Schedule.
445



Thermal barrier coating life prediction model development

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