Turbine Engine Hot Section Technology, 1985

Figure 18. STRESSES IN X - DlRECTKlN Part: 2C Unit: MPa Figure 19. STRESSES IN V - DIRECTKJN F i u r e 20. SHEARW STRESSES Figure 23. SHEARING STRESSES

Figure 24. STRESSES N X - DIRECTKM

Figure 25. STRESSES IN Y - DIRECTION

Part: 2C Ceramic Unit: MPa I F~gure 26. SHEARING STRESSES PROPOSED MECHANISM FOR OXIDATION INDUCED THERMAL BARRIER COATING FAILURE PROPOSED CERAMIC RESPONSE CALCULATED STRESS STATES Thermal Expans~on Mosmatch Radial Stress Microcracking and Microcrack Coalescence at Asperities ceramic ceramic - bond coat Crack Extens~on Leading t o Spalling Oxodatlon Growth Radaal Stress ceramic ceramic

- 0 f '\\ 0 f'

\ I \ o , = o / I ortde layer oxide layer bond coat bond coat Flgure 27. PROPOSED FAILURE MECHANISM I THERMAL BARRIER COATING LIFE PREDICTION I MODEL DEVELOPMENT T.E. Strangman, J.F. Neumann, and A. Tasooji Garrett Turbine Engine Company Phoenix, Arizona Thermal barrier coatings (TBC's) 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 (fig.

1 ) . To achieve these benefits, however, the TBC system must be reliable.

Mechanistic thermomechanical and thermochemical life models are therefore required for the reliable exploitation of TBC benefits on gas turbine airfoils. Garrett'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 (fig. 2 ) . The plasma-sprayed TBC system, composed of a low pressure plasma sprayed (LPPS) applied oxidation resistant NiCrAlY bond coating and an air plasma sprayed yttria (8 percent) partially stabilized zirconia insulative layer, is applied by both Chromalloy (Orangeburg, New York) and Klock (Manchester, Connecticut). The second type of TBC is applied by the electron beam-physical vapor deposition (EB-PVD) process by Temescal (Berkeley, California).

A viable model must predict TBC life on a turbine airfoil as a function of engine, mission, and materials system parameters. These parameters are incorporated into mechanical, oxidation, and salt deposition functions of TBC degradation as indicated in figure 3. The approach adopted in this program for developing a TBC life model is similar to that in use at Garrett for prediction of oxidation/hot corrosion lives of metallic coatings as a function of engine, mission, and materials system parameters (fig. 41. Similarities and differences in these models are illustrated in figure 5.

I A rapid computational capability is required for preliminary design and mission analyses of TBC lives. Substructure models are being developed to facilitate the rapid computation of TBC life as indicated in figure 6. TBC life analysis will be performed for each of the critical damage modes.

Burner rig and mechanical property data are being obtained to quantify the capabilities of each of the TBC systems for each critical mode of degradation.

Bilrner rig test data and zirconia fracture toughness data are illustrated on figures 7 and 8.

Lives of these TBC systems will be predicted for TFE731 high pressure turbine blades for factory engine test, business aircraft and maritime surveillance missions. Complementary engine validation tests are planned.

This program is now at the midpoint of Phase I. The program schedule is provided in figure 9.

I

I 'R!!XEDING PAGE BLANK NOT F I U D 125pm THICK ZlRCONlA COATING UNCOATED BLADE COOLING AIR REQUIREMENT. PERCENT Figure 1. TBCs Improve Creep Life and Reduce Cooling Air Requirements for the Garrett High-Pressure Turbine Blade.

ELECTRON BEAM -

PLASMA SPRAY PHYSICAL VAPOR DEPOSITION APS Y2O3 (8°/~] EB-PVD Y2O3 [20%] TBC

-

STABILIZED ZrOp STABILIZED ZrO2 EB-PVD BASE COAT LPPS Ni-31 Cr-11 A1 -0.5Y Ni-23Co-18Cr-11 A1 -0.3Y MAR-M 2 4 7 MAR-M 2 4 7 SUBSTRATE SUPERALLOY SUPERALLOY CHROMALLOY TEMESCAL KLOCK Figure 2. Life Prediction Models are Being Developed for Plasma-Sprayed and EB-PVD TBC Systems.

TBC = F1 [MECHANICAL) + F2 (OXIDATION) + Fg (SALT DEPOSITION)

ALTITUDE (SALT INGESTION) DEGRADAT1oN COATING STRESSES TEMPERATURE RATE TEMPERATURE CYCLE SEGMENT LENGTH TURBINE PRESSURE MATERIAL SYSTEM MATERIALS SYSTEM SALT EVAPORATION

' KIC SALT SOLIDIFICATION

FLAW SIZE TEMPERATURE ELASTIC MODULUS GAS VELOCITY SPALLING STRAIN AIRCRAFT LOCATION MATERIALS SYSTEM TBC Life is a Function of Engine, Mission and Figure 3.

Materials System Parameters.

ENGINE AND MISSION PARAMETERS

AIRCRAFT SULFUR TURBINE ALTITUDE CONTENT PRESSURE SULFUR IN FUEL = 0.2 PERCENT SULFUR IN FUEL = 0.2 PERCENT

ld j \ PRESSURE = 8 ATMOSPHERES

ALTITUDE = o [SALT IN AIR = 0.012 PPMJ PRESSURE = 25 ATMOSPHERES ALTITUDE = o [SALT IN AIR = 0.012 PPMJ SALT IN AIR AT SEA LEVEL = 0.012 PPM

-

i - 16, I E .

2 : N \ .

ua 0 -

s 104

w - Y -1

U -

$ I.

U TEMPERATURE. F TEMPERATURE. F TEMPERATURE. F

- - -

600 700 800 900 1000 1100 000 700 Boo 900 loo0 1100 MI0 700 800 900 l o o 0 1100 TEMPERATURE. C TEMPERATURE. C TEMPERATURE. C 65-1 954 Turbine Airfoil Coating Life Predicted by Computer Figure 4.

Model.

TBC LIFE MODEL

METALLIC COATING LIFE MODEL

\ SALT \

' SOLIDIFICATION

:\ SALT

~ I I D I F I C A T I O N SALT LY Y '\ EVAPORATION 2 \ m ca \ BOND C

s

OXIDATION I- COATING C1 '\ u \ d U

'

EVAPORATION d ZlRCONlA LOW TEMPERATURE INTERMEDIATE TOUGHNESS HOT CORROSION TEMPERATURE REDUCTION HOT CORROSION TEMPERATURE TEMPERATURE 65195.5 Figure 5. Approach to TBC Life Model Development is Facilitated by GTEC Metallic Coating Life Model.

SUBSTRUCTURE MOOEL ANALYSIS WILL BE PERFORMED FOR SELECTED ELEMENTS SUBSTRUCTURE MODEL FOR TBC STRESS ANALYSIS Figure 6. Substructure Models Facilitate the Rapid Computation of TBC Life Required for Preliminary Design and Mission Analysis.

600 -

400 -

200 -

"

a => = 100- 80- -1 60- 4: CL

" 40-

C \ W \ E \ i= 20- \ \ \ A KLOCK 0 TEMESCAL lo: 0 CHROMALLOY 1 r 1 1 1000 1050 1100 1150 65.195 1 TEMPERATURE. C Figure 7. TBC Spalling in Burner Rig Test is a Function of Temperature.

1.5- (-j- ll0OC 0 +

%

a 0- a

-

1000C/0.5 HOUR

: 1.0-

-

" 2

, J -- -&- - - 0 0

r -1 I C1

Q rr

0 'A L , , , d , - - 1150C I-

- AS-RECEIVED

a a

E 0.5 -

TOUGHNESS => I- u a LL 0 lO0OC 0 ll0OC A 1150C 1 1 I I I I I I I I I I I 1 2 4 6 8 1 0 20 40 60 80 100 200 400 600 800 C5 195 1 1 TIME AT TEMPERATURE. HOURS ' Figure 8. Zirconia Fracture Toughness of Chromalloy Plasma- Sprayed TBC System is a Step Function of Exposure Time and Temperature.

I 1984 I 1985 I 1986 I 1987 , 1988 , 1989 , PHASE I - FAILURE MODES ANALYSES AN0 MODEL DEVELOPMENT

------------

r-

I FAILURE m t c n A w s m s O ~ T E R I I I ~ T ~ I

I

ITASK

MUOR m o o t L l r t P~~OICTIOII

I

ITASK

m o M L YERlFlCATlOl I I EB no r o c 1

I lASK IV

l t r n R T l l G [ I

ITASK

FACTOR1 t l S l l t TESTS I 1 I

ITMK " -

I PHASE 11 [O~TION) -

1 OESlGN CAPABLE LIFE MOOELS

TASK VI THERMOMtCHAllCAL LIFE MOOtLS

I

I r n t n m o c n t n ~ c r ~ FAILURE l r o o t s l

TASK VII- I

I

TASK Vlll 1

C O I C M H t l S l M I l l 4 M A T t O I

urt n r o ~ r ~ a n 0

TASK I x I

LIFE P M m C T I O I m W t L d

TASK 1 I

FIELO E l 6 l l t TEST

I

L - - - - - - 8 Figure 9. TBC Life Prediction Schedule.

THERMAL BARRIER COATING LIFE PREDICTION M O D E L DEVELOPMENT* R . V . H i l l e r y , B.H. P i l s n e r , E . C . Duderstadt General E l e c t r i c A i r c r a f t Engine Business Group The o b j e c t i v e s of t h i s program a r e t o determine t h e predominant modes of d e g r a d a t i o n of a plasma sprayed thermal b a r r i e r c o a t i n g system, and t h e n t o develop and v e r i f y l i f e p r e d i c t i o n models accounting f o r t h e s e d e g r a d a t i o n modes. The program i s d i v i d e d i n t o two phases, e a c h c o n s i s t i n g of s e v e r a l t a s k s . The work i n Phase I i s aimed a t i d e n t i f y i n g t h e r e l a t i v e importance of t h e v a r i o u s f a i l u r e modes, and developing and v e r i f y i n g a l i f e p r e d i c t i o n model(s) f o r t h e predominant mode f o r a thermal b a r r i e r c o a t i n g system. Two p o s s i b l e predominant f a i l u r e mechanisms being e v a l u a t e d a r e bond c o a t o x i d a t i o n and bond c o a t creep. The work i n Phase I1 w i l l develop design-capable, c a u s a l , l i f e p r e d i c t i o n models f o r thermomechanical and thermochemical f a i l u r e modes, and f o r t h e e x c e p t i o n a l c o n d i t i o n s of f o r e i g n o b j e c t damage and e r o s i o n .

TBC SYSTEMS The primary TBC system c o n s i s t s of a low p r e s s u r e plasma-sprayed (LPPS) bond c o a t l a y e r of Ni-22Cr-lOAl-0.3Y, an a i r plasma sprayed y t t r i a p a r t i a l l y s t a b i l i z e d z i r c o n i a (Zr02-8%Y203) t o p c o a t , on a c o n v e n t i o n a l l y c a s t Rene '80 s u b s t r a t e a l l o y (Table 1 ) . This bond c o a t composition has been demonstrated t o p o s s e s s good o x i d a t i o n r e s i s t a n c e and has a l a r g e d a t a base a s a TBC bond c o a t . The Zr02-8%Y2D3 t o p c o a t was chosen s i n c e numerous s t u d i e s have shown t h a t z i r c o n i a p a r t i a l l y s t a b i l i z e d w i t h 6-8 w t . % Y2O3 i s t h e b e s t composition f o r plasma sprayed TBC's ( 1 ) . The Rene ' 8 0 s u b s t r a t e was chosen s i n c e a l a r g e TBC d a t a base i s p r e s e n t f o r t h i s s u b s t r a t e composition.

Four d i f f e r e n t TBC systems u t i l i z i n g f o u r d i f f e r e n t bond c o a t s a r e being e v a l u a t e d i n t h e experiment t o e v a l u a t e t h e e f f e c t of bond c o a t c r e e p s t r e n g t h on TBC t h e r m a l c y c l e l i f e (Table 2). These f o u r TBC systems a l s o u t i l i z e TBC system #1 has t h e Zr02-8%Y203 t o p c o a t s and Rene '80 s u b s t r a t e s . , same N i C r A l Y bond c o a t u t i l i z e d i n t h e primary TBC system. TBC systems #2, #3,and /I4 have modified N i C r A l Y bond c o a t s which have r e c e i v e d a l l o y a d d i t i o n s t o i n c r e a s e t h e bond c o a t c r e e p s t r e n g t h . An aluminide o v e r c o a t i s used i n each of t h e s e systems (1-4) t o reduce d i f f e r e n c e s i n o x i d a t i o n r e s i s t a n c e f o r t h e f o u r bond c o a t s . A comparison of t h e primary TBC system and i t s c o u n t e r p a r t w i t h an aluminide overcoat a s t h e bond c o a t i s shown i n F i g u r e 1.

THERMAL CYCLE TESTING Thermal c y c l e t e s t i n g i s being performed i n an automated Rapid Temperature Furnace ( F i g u r e 2). The thermal c y c l e s c o n s i s t s of t e n m i n u t e s h e a t up, a 45 minute exposure a t 1093OC (2000°F), and 1 5 minutes f o r c e d a i r c o o l i n g . T h i s f u r n a c e u t i l i z e s a l i f t which a u t o m a t i c a l l y c y c l e s t h e specimens from t h e upper f u r n a c e exposure zone t o t h e lower c o o l i n g compartment where a f a n p r o v i d e s f o r c e d a i r c o o l i n g .

*Work done under NASA C o n t r a c t NAS3-23943.

PRE-EXPOSURES I N A I R AND ARGON I n both t h e bond c o a t o x i d a t i o n and bond c o a t c r e e p e v a l u a t i o n s , pre-exposures i n a i r and argon have been u t i l i z e d t o e v a l u a t e t h e e f f e c t of bond c o a t o x i d a t i o n on thermal c y c l e l i f e . The specimens a f t e r e a c h of t h e pre-exposure c o n d i t i o n s (as-sprayed, argon, and a i r pre-exposures) a r e shown i n Figure 3. The specimens pre-exposed i n argon have t o p c o a t s t h a t a r e g r a y i n appearance which can be a t t r i b u t e d t o oxygen d e f i c i e n c y , w h i l e t h e bond c o a t ' s s i l v e r c o l o r i s i n d i c a t i v e of t h e l a c k of s i g n i f i c a n t o x i d a t i o n . The specimens pre-exposed i n a i r show t h a t t h e bond c o a t s have darkened due t o o x i d a t i o n , while t h e t o p c o a t ' s yellow appearance i s t y p i c a l of e l e v a t e d temperature a i r exposure.

BOND COAT OXIDATION EXPERIMENT I n t h e bond c o a t o x i d a t i o n experiment, pre-exposures i n a i r and argon f o r d i f f e r e n t pre-exposure times were u t i l i z e d . The g o a l of pre-exposures i n a i r was t o develop oxide s c a l e s p r i o r t o thermal c y c l i n g , w h i l e t h e g o a l of t h e pre-exposures i n argon was t o a l l o w t h e o t h e r t h e r m a l l y a c t i v a t e d phenomena p r e s e n t i n t h e a i r pre-exposures t o occur without developing t h e oxide s c a l e .

T h i s should a l l o w i s o l a t i o n of t h e e f f e c t s of bond c o a t o x i d a t i o n .

Pre-exposures u t i l i z e d were 10, 50, 100, and 500 hours both i n a i r and argon.

The oxide scale t h i c k n e s s e s of t h e a i r pre-exposures a r e p l o t t e d a s a f u n c t i o n of pre-exposure time i n Figure 4.

Unexpectingly , t h e specimens pre-exposed i n argon f a i l e d b e f o r e t h e

specimens pre-exposed i n a i r ( F i g u r e 5). The i n i t i a l h y p o t h e s i s f o r t h i s unexpected r e s u l t i s t h a t a phase d i s t r i b u t i o n change may be o c c u r r i n g i n t h e ceramic t o p c o a t . Another s u r p r i s i n g r e s u l t was t h e l o n g e r l i v e s f o r t h e specimens pre-exposed i n a i r f o r 1 0 hours. The l o n g e r l i v e s f o r t h e s e specimens than f o r t h e as-sprayed specimens may be due t o bond c o a t s i n t e r i n g and perhaps t o p c o a t s i n t e r i n g . The s i n t e r i n g r e s u l t s i n l e s s c o n s t r a i n t f o r t h e A1203 s c a l e , w h i l e providing b e t t e r t o p coatlbond c o a t adherence (improved chemical bonding).

F a i l u r e f o r t h e s e specimens was d e f i n e d a s when 10% ( s u r f a c e a r e a ) of t h e ceramic t o p c o a t has s p a l l e d ( F i g u r e 6). I n a l l c a s e s s p a l l i n g i n i t i a t e d a t t h e edges of t h e t o p c o a t . The f a i l u r e l o c a t i o n was t h e same f o r a l l pre-exposure c o n d i t i o n s and occurred i n t h e ceramic approximately 0.0025-0.0050 cm (0.001-0.002") from t h e bond c o a t / t o p c o a t i n t e r f a c e .

D i f f e r e n c e s due t o argon and a i r pre-exposures a r e c l e a r l y r e f l e c t e d i n t h e pre-exposed, specimen m i c r o s t r u c t u r e s ( F i g u r e 7). I n a l l c a s e s , a continuous A1203 f i l m formed a t t h e bond c o a t / t o p c o a t i n t e r f a c e f o r specimens pre-exposed i n a i r . The e f f e c t of o x i d a t i o n i s a l s o seen i n t h e bond c o a t m i c r o s t r u c t u r e where d e p l e t i o n of t h e high Al 6 phase i s observed a t t h e bond c o a t / t o p c o a t i n t e r f a c e . The e f f e c t of i n t e r d i f f u s i o n i n d e p l e t i o n of t h e high Al B phase i n t h e bond c o a t i s a l s o observed a t t h e bond c o a t s u b s t r a t e i n t e r f a c e . The argon pre-exposures, on t h e o t h e r hand, s i g n i f i c a n t l y r e t a r d e d bond c o a t o x i d a t i o n . The e f f e c t i v e n e s s i s demonstrated by t h e small degree of B d e p l e t i o n ( h i g h A 1 phase) i n t h e bond c o a t a t t h e bond c o a t / t o p c o a t i n t e r f a c e and t h e absence of A1203 s c a l e .

These photomicrographs a l s o show t h a t s i g n i f i c a n t i n t e r d i f f u s i o n t h a t has occurred between t h e s u b s t r a t e and bond c o a t .

D i f f e r e n c e s due t o argon and a i r pre-exposures a r e s t i l l r e f l e c t e d i n t h e pre-exposed, specimen m i c r o s t r u c t u r e a f t e r thermal c y c l i n g . The m i c r o s t r u c t u r e of t h e specimen pre-exposed i n argon a f t e r t h e r m a l c y c l i n g shows t h a t v e r y l i t t l e B d e p l e t i o n has occurred i n t h e bond c o a t a t t h e bond c o a t / t o p c o a t i n t e r f a c e . T h i s i s due t o t h e argon pre-exposure and t h e s h o r t thermal c y c l e l i v e s of t h e s e specimens ( i . e . very l i t t l e o x i d a t i o n has occurred). This i s c o n t r a s t e d w i t h t h e m i c r o s t r u c t u r e of t h e specimens pre-exposed i n a i r a f t e r thermal c y c l i n g where s i g n i f i c a n t l y more P d e p l e t i o n has occurred r e s u l t i n g i n a l a r g e r v l a y e r a t t h e bond c o a t / t o p c o a t i n t e r f a c e . This i l l u s t r a t e s t h e i n c r e a s e d demand t o c o n t i n u e t h e growth of t h e A1203 s c a l e . Something t h a t may be n e c e s s a r y t o i n c l u d e i n o u r model i s t h e e f f e c t t h a t a changing bond c o a t m i c r o s t r u c t u r e has on TBC i n t e g r i t y .

Oxide scale t h i c k n e s s measurements a f t e r thermal c y c l i n g i n d i c a t e t h a t bond c o a t o x i d a t i o n may be a s i g n i f i c a n t c o n t r i b u t i o n t o t h e f a i l u r e mechanism. The o x i d e s c a l e t h i c k n e s s a f t e r thermal c y c l i n g f o r t h e specimens pre-exposed i n a i r and t h e specimens r e c e i v i n g no-pre-exposure a r e p l o t t e d i n Figure 8. The p l o t shows t h a t t h e oxide s c a l e i s e s s e n t i a l l y t h e same f o r t h e specimens r e g a r d l e s s of pre-exposure t i m e w i t h t h e e x c e p t i o n of t h e 500 hr.

pre-exposure. T h i s i n d i c a t e s t h a t bond c o a t o x i d a t i o n may be important. The 500 hr. pre-exposure r e s u l t i n d i c a t e s t h i s 4 y m oxide s c a l e t h i c k n e s s i s n o t l a r g e enough t o c a u s e f a i l u r e i f o t h e r t h e r m a l l y a c t i v a t e d phenomena have n o t occurred. The oxide s c a l e s f o r t h e specimens pre-exposed i n argon were less t h a n 1 y m t h i c k .

BOND COAT CREEP EFFECT EXPERIMENT A s mentioned p r e v i o u s l y , t h e bond c o a t c r e e p e f f e c t experiment u t i l i z e d f o u r bond c o a t s a l l of which had r e c e i v e d aluminide over c o a t s t o reduce d i f f e r e n c e s i n o x i d a t i o n r e s i s t a n c e . Again, pre-exposures i n a i r and argon were u t i l i z e d (100 hr. pre-exposures). The r e s u l t s f o r a i r and argon pre-exposures a g a i n show t h a t pre-exposures i n argon a r e more d e t r i m e n t a l t h a n a i r pre-exposures ( F i g u r e 9). Also a s expected, t h e TBC's w i t h t h e N i C r A l Y bond c o a t , which has t h e lowest bond c o a t c r e e p s t r e n g t h , a l s o had t h e lowest thermal c y c l e l i f e . However, t h e s e p a r a t i o n between t h e o t h e r t h r e e bond c o a t c r e e p systems i s n o t c l e a r . These r e s u l t s a r e a l s o t r u e f o r t h e specimens t h a t r e c e i v e d no pre-exposure ( F i g u r e 10). It i s b e l i e v e d t h a t t h e bond c o a t c r e e p d i f f e r e n c e between t h e s e t h r e e p a r t i c u l a r systems ( 2 , 3, 4) i s n o t l a r g e enough t o o f f s e t t h e e f f e c t s of o t h e r c o n t r i b u t i o n s t o f a i l u r e . A new bond c o a t which has been developed a t G E which has a c r e e p s t r e n g t h lower than t h e t h r e e and c l o s e r t o t h e bond c o a t c r e e p s t r e n g t h of N i C r A l Y w i l l be used i n f u t u r e t e s t i n g t o demonstrate t h i s e f f e c t more d e f i n i t i v e l y .

KEY PROPERTY DETERMINATIONS Key p r o p e r t y d e t e r m i n a t i o n s of t h e bond c o a t and t h e t o p c o a t w i l l a l s o be made i n t h i s study. The methods and c o n d i t i o n s a r e l i s t e d i n F i g u r e 11.

Generally, c o n v e n t i o n a l t e s t i n g can be u t i l i z e d f o r t h e bond c o a t m a t e r i a l s , whereas s p e c i a l t e s t i n g i s needed f o r t h e ceramic c o a t i n g s .

TBC LIFE PREDICTION MODEL The f a i l u r e mechanism e v a l u a t i o n s and key p r o p e r t y d e t e r m i n a t i o n s w i l l then be coupled w i t h thermomechanical s t u d i e s t o develop t h e l i f e p r e d i c t i o n models. I n i t i a l s t u d i e s a r e aimed a t determining t h e magnitude of s t r a i n s p r e s e n t . The work w i l l be done u t i l i z i n g a t u b u l a r LCF b a r , p e r m i t t i n g i n t e r n a l c o o l i n g t o develop thermal g r a d i e n t s ( F i g u r e 1 2 ) . These i n i t i a l s t u d i e s w i l l be used t o set up a modified t e s t m a t r i x of experiments based on t h e i n i t i a l s t u d i e s t o g e n e r a t e t h e thermomechanical d a t a f o r t h e model.

1. S t e c u r a , S. " E f f e c t s of Compositional Changes on t h e Performance of a Thermal B a r r i e r Coating System," NASA T M 78976, 1979.

2. M i l l e r , R.A., G a r l i c k , R.G., and Smialek, J.L., "Phase D i s t r i b u t i o n s i n Plasma Sprayed Z i r c o n i a - Y t t r i a , " American Ceramic S o c i e t y B u l l e t i n , V.62, Dec. 1983, p. 1355-1358.

Tcble 1

Baseline Thermal Barrier Coating System

(Weight Percent) Substrate (Renef80): Ni-14Cr-9. ~ C O - ~ T ~ - ~ W - ~ M O - ~ A I - O . ~ ~ C - 0.321-0.0156 Ni-22Cr-10AI-0.3Y (Low Pressure Bond Coating: Plasma Spray)

Top Coating: Zr02 - 8Y203 (Air Plasma Spray)

Tcble 2

Bond Coat Creep Effect TBC Systems

Bond Coat Creep SubstratetBond Coating/Over (Larson/Miller Parameter Systems Coating/Top Coating at 3 ksi - rupture test) 39.0 Rene180/Bond Coating l'/Alumlnlde/ I Zr02-Y 2 0 , 45.7 2 Rene180/Bond Coating 2*/Alumlnlde/ Zr02-Y203 47.0 Rene180/Bohd Coating 3*/Aluminide/ Zr02-Y203 48.4 Rene180/Bond Coating 4*/Aluminidel Zr02-Y203 ' NI-22Cr-10AI-0.3Y Modifled NlCrAlY Bond Coats

As Sprayed TBC Microstructures

Figure 1 As-sprayed TBC microstructures Figure 2 Rapid-temperature furnace ORIGINAL PAGE IS @F POOR QUALITy Figure 3 Specimens prior to thermal cycle testing

Oxidizing Pre-Exposure at 2000" F

Oxlde Scale Thickness (Microns) O J 1 I I I I I I I I 1 1 0 50 100 150 200 250 300 350 400 450 500 Pre-Exposure Tlme - (Hours) Rapid Temperature Thermal Cycle Test at 2000°F

45 Minute Exposure - 15 Minute Cool Down

Cycles to Coating Failure Oxldlzlng Pre-Exposure

-

I I I I I T 0 I I 1 I 0 50 100 150 200 250 300 350 400 450 500 Pre-Exposure Time (Hours at 2000°F) Floure 5 t e s t i n g F i g u r e 6 F a i l e d s p e c i m e n s a f t e r t h e r m a l c y c l e

NiCrAlY Bond Coat

100 Hour Air Pre-Exposure

0 0

After Pre-Exposure After Thermal Cycllng q

z 8

100 Hour Argon Pre-Exposure 5 K

9 !a

After Pre-Exposure After Thermal Cycllng

Figure 7

A1203 Scale Thickness at Failure

After Thermal Cycle Testing

(NiCrAIY Bond Coat)

Average Oxide Scale Thickness at Failure (Microns) ,,MU

'

Pre-Exposure Thlckneaa

Pre-Exposure Tlme (Hours) -

Figure 8

100 Hour Pre-Exposure

Cycles to Failure 400 1 2 3 4 4 3 2 1 Argon Air 1 NI-22Cr-10AI-0.3Y 2,3,4 Modltied NlCrAlY Bond Coats Figure 9

As Sprayed

Cycles to Fsllure 1 NI-22Cr-10AI-0.3Y 2,3,4 ModHktd NlCrAlY Bond Coats Figure 10

Key Mechanical Properties

Bond TOP Coating Coating

1 Conventional

1 Tensile Strength

I Bend I

Resonance* Conventional Poisson's Ratio

1 Conventional 1 Conventional

1 Coefficient of Thermal Exposure

Conventional Resonance Dynamic Modulus Temperatures R.T., 1 OOO°F, 1 800°F, 1 900°F, 2000°F 'Also Strain Gauge at R.T.

Figure I 1

Thermomechanical Properties

Furnace Furnace cooiing Alr Determine Magnitude of Strains Present Utilize Test Matrix of LCF Testlng THERMAL BARRIER COATING LIFE PREDICTION MODEL DEVELOPMENT*

~

J.T. DeMasi and K.D. Sheffler United Technologies Corporation Pratt & Whitney The objective of this program is to develop an integrated life prediction model accounting for a1 1 potential 1 ife-1 imiting Thermal Barrier Coating (TBC) degradation and failure modes including spallation resulting from cyclic thermal stress, oxida-

tion degradation, hot corrosion, erosion and foreign object damage (FOD) . This

overall program objective will be accomplished in two phases. The goal of the first phase is to determine the mechanisms and relative importance of the various degra- dation and failure modes, and to develop and verify the methodology to predict predominant mode failure life in turbine airfoil applications. Phase I will develop an empirically-based correlative model relating coating life to parameterically expressed driving forces such as temperature and stress. The effort in this phase consists of three tasks: Failure Mechanism Determination (Task I), Modeling (Task

11) , and Substantiation Testing (Task 111). Phase I1 will experimentally verify

Phase I models and develop an integrated, mechanistically-based life prediction model including all relevant failure modes. The program is currently in the final stages of Task I; predominant failure modes have been identified and a preliminary life prediction model is being developed.

The two layer TBC system being investigated, designated PWA264, is currently in commercial aircraft revenue service, on turbine vane platforms in the JT9D and 2037 engines. It is also bill-of-material on turbine vane airfoils in the advanced PW4000 and IAE V2500 engines. The TBC consists of an inner low-pressure chamber plasma sprayed NiCoCrAlY metallic bond coat (4-6 mils) and an outer air plasma- sprayed 7 w/o Y203-Zr02 (8-12 mils) ceramic layer (figure 1). The composition and structure of this coating are based in part on effort conducted under previous NASA sponsored programs (ref. 1 and 2 ) .

PHASE I, TASK I - FAILURE MECHANISM DETERMINATION

I

A review of experimental and flight service components as well as laboratory test evaluations indicates that the predominant mode of TBC failure involves thermomechanical spallation of the ceramic coating layer. This ceramic spallation involves the formation of a dominant crack in the ceramic coating parallel to and closely adjacent to the topol~ogically complex metal ceramic interface (figure 2).

This cyclic "mechanical" failure mode clearly is influenced by thermal exposure effects as shown by results of experiments conducted to study thermal pre-exposure and thermal cycle-rate effects (ref. 3-6).

EXPERIMENTAL DESIGN AND TEST PLAN The Task I, "Failure Mechanisrr Determination" investigation was designed to

~

evaluate the relative importance of various thermomechanical and thermochemical "damage" nodes, focusing on thermal stress cycling, oxidative degradation and their potential interaction. The primary experimental method used in this investigation was cyclic burner rig testing. The cyclic tests were conducted with both clean and

i

*Work done under NASA Contract NAS3-23944.

contaminated fuels to assess the importance of hot corrosion induced ceramic

spallation (ref. 7 - 10). Static furnace tests also were performed to evaluate the

relative importance of oxidation and other thermal exposure effects. The test matrix (figure 3) was designed to study the influence of various "driving forces" such as temperature, thermal cycle frequency, environment, coating thickness and pre-burner rig test thermal exposure on TBC spalling life. To provide property data required for subsequently described thermal and stress analyses, physical and mechanical property tests are being conducted on monolithic ceramic and metallic specimens fabricated to simulate the composition and structure of the respective coating layers.

CRITICAL EXPERIMENT RESULTS

I

All burner rig and furnace test specimens exhibited the typical ceramic spalla- tion near the metal-ceramic interface, with a thin layer of ceramic remaining adherent after failure. Examination of the laboratory data clearly shows a strong temperature effect; comparison of these data with typical engine test conditions suggests that ''cyclic content", (i.e., relative frequency and severity of engine thermal cycling) also strongly influences TBC spallation life (figure 4). Oxidation damage occurring at the ceramic-metal interface for laboratory testing was found to be somewhat greater than that found for engine exposed failures. This is attributed to the relatively high interface temperature employed in the accelerated laboratory spallation life testing.

In the laboratory tests conducted to study environmental effects, results suggest that bond coat oxidation damage at the metal-ceramic interface contributes significantly to thermomechanical cracking in the ceramic layer. Low cycle rate furnace exposure in air versus exposure in Argon clearly shows a dramatic increase of spall ing life in the non-oxidizing environment (figure 5). The results of burner rig testing indicated that static thermal pre-exposure of burner rig test specimens in air causes a proportionate reduction of cyclic thertxal spalling life, whereas pre-test thermal exposure in Argon does not reduce cyclic thermal spalling life (figures 6 and 7). Typical respective pre-test microstructures for air and Argon pre-exposed specimens are shown in figures 8a and b .

I Laboratory testing was conducted in clean and contaminated (Na, S) fuel envi- ronments to evaluate the hot corrosion spallation resistance of the TBC. Corrodant induced failure was observed during cyclic hot corrosion testing at high corrodent I levels (35 ppm Na2S04) but not for low corrodent levels (10ppm Na2S04). The failure mode, which has not been observed on engine exposed components, involved "flaking" of small patches of ceramic above the typical failure location.

Testing was also conducted to evaluate the effects of ceramic thickness on TBC spalling life. Ceramic thickness was found to have an effect on coating durability (figure 9). Thick coatings were found to decrease TBC life while thin coatings increased it as compared to the "baseline" 1 0 mil thick ceramic.

PHASE I, TASK I PRELIMINARY LIFE PREDICTION MODEL DEVELOPMENT

The prel iminary 1 ife prediction model currently being developed focuses on the two major damage modes identified in the laboratory testing described above. The first of these modes involves a mechanical driving force, resulting from cyclic strains and stresses caused by thermally induced and externally imposed loads. The second is an environmental driving force which appears, based on the experimental results, to be related to "oxidation damage", most probably to the in-service growth of a NiCoCrAlY oxide scale at the metal-ceramic interface. Based on the apparently "mechanical" mode of ceramic failure, it is presumed that the growth of this oxide scale influences the intensity of the mechanical driving force. The mechanism(s) of this "interaction" are not presently understood, and no attempt is being made to incorporate interaction effects in the initial model, which will be based on linear damage summation. Interaction effects will be considered in the refined model to be developed in Task 2 of this program.

Mechanical failure of the ceramic layer is presumed to involve accumulation of fatigue "damage". Possible mechanisms for the accumulation of this damage might involve the initiation and propagation of a dominant crack in the ceramic, or possibly the subcritical growth and subsequent link-up of pre-existing microcracks in the ceramic structure. Metallographic examination of specimens removed from burner rig test prior to spallation failure presently is being conducted to identify specific mechanical damage accumulation rnechanism(sj.

Cyclic inelastic strain range in the ceramic layer will be used to represent the driving force for mechanical damage in the life model. Use of this parameter is based on results of mechanical (reversed bend) tests conducted on monolithic ceramic specimens having a porous, microcracked microstructure representative of the ceramic coating. These results have shown highly non-linear stress-strain behavior with

significant stress-strain hysteresis in reversed loading . Finite element calcula-

tions of ceramic inelastic strain range are being conducted for each of the Task I burner rig test conditions using transient thermal data obtained from thermocouple instrumented test specimens. It is presently planned to use a relatively simple empirical relationship such as Manson-Coffin to express the functional dependence of mechanical "damage" on ceramic inelastic strain range.

Based on the observation that thermal exposure damage appears to be related to oxidation effects, the relationship between thermal damage accumulation rate and primary exposure parameters (time and temperature) will be based on the accepted parabolic and exponential forms appropriate to oxidation kinetics.

, I A major shortcoming of the present model is the absence of any provision for

~

interaction between environmental and mechanical damage. The relatively coarse preliminary finite element break-up constructed to represent the substrate coating structure incorporates a planar metal-ceramic interface and predicts essentialy no change of stress level with growth of an interfacial oxide scale. Thus, oxidation effects will be "forced" in the preliminary model using the linear damage summation approach. One approach which will be evaluated in an effort to incorporate interac- tion effects in the refined Task 2 model will involve an attempt to represent, in a relatively simple geometric form such as that employed by G. C. Chang (ref. 11), the very complex (rough) topological form of the real physical interface shown in figure 1. Other changes to the relatively simple functional forms used in the preliminary Task I model undoubtedly will be suggested by ongoing microstructural damage interpretation and by testing this preliminary model against additional burner rig verification tests to be conducted at the conclusion of Task I.

I VERIFICATION TESTING - WORK PLANNED

To verify the preliminary Task I prediction model, additional burner rig tests

will be conducted using test parameters and methods which are different from those used to generate the data on which the model is based. The test method will involve exposure of a single rotating specimen located in the center of the burner rig spindle. This will improve and simp1 ify temperature measurement and control, and

i

I will eliminate circumferential thermal gradients which are inherent to the multiple specimen configuration used earlier in this task. To improve the simulation of air- foil conditions the specimen will be hollow and incorporate internal cooling, thus providing a steady state thermal gradient across the TBC. Three sets of test para- meters will be selected to simulate typical airfoil mission cycles.

1 REFERENCES 1. Sheffler, K.D.; Graziani, R.A.; and Sinko, G.C.: JT9D Thermal Barrier Coated Vanes. NASA CR 167964, April 1982.

Anderson, N.P.; and Sheffler, K.D.: Development of Strain Tolerant Thermal 2.

Barrier Coating Systems. Final Report, Contract MAS3-22548. 1982.

3. McDonald, G.; and Hendricks, R . C . : Effect of Thermal Cycling on Zr02-Y203 Thermal Barrier Coatings. Thin Solid Film. V.73, 1980, p.491.

Gedwill, M.A.; Burner Rig Evaluation of Thermal Barrier Coating Systems for 4.

Nickel-Base Alloys. NASA-TM 81685, February 1981.

5. Miller, R.A.; and Lowel, C . E . : Failure Mechanisms of Thermal Barrier Coatings Exposed to Elevated Temperatures. MASA TM 82905, April 1982.

6. Miller, R.A.; Oxidation-Based Model for Thermal Barrier Coating Life. Ceramic Journal. 1984, pp. 83-87.

7. Grisaffe, S.J.; and Levine, S.R.: Proceedings of First DOE/EPRI Conference on Advanced Materials for Alternative Fuel Capable Directly Fired Heat Engines.

Castine, ME. 1979, p . 680.

8. Bratton, R.J.; et. al.: Evaluation o f Present Day Thertxal Barrier Coatings for Industrial/Utility Applications. Thin Solid F i l m . 73, 1980, p. 429.

9. Hodge, P.E.; et. al.: Evaluation of the Hot Corrosion Behavior of Thermal Barrier Coatings. Thin Solid Films. 73, 1980, p. 447.

10. Bevan, C.E.; Development of Advanced Plasma Sprayed Ceramic Coatings for Industrial Gas Turbine Engines. Final Report, Contract B-A0747-A-Z. PWA 5906, July 1982.

Finite Element Analyses of Thermal Barrier 11. Chang G.C.; and Phucharoen W . : Coatings. Thermal Barrier Coatings Workshop, Op. cit . , May 1985, p . 1 1 1 - 126.

Figure 1 Light Photomicrograph Showing PWA264 ;4icrostructure 200X ORIGINAL PAGE IS OF POOR QUALITY Figure 2 Typical Thermal Barrier Coating Failure Mode T EXPOSURE STATIC CYCLIC FRACTIONAL EXPOSURE E M TEST BURNER RIG FURNACE BURNER RIG FURNACE P E ATMOSPHERE OXIDIZING NON-OXIDIZING OXIDIZING HOT CORROSION OXIDIZING NON-OXIDIZING OXIDIZING HOT CORROSION R CYCLE LENGTH SHORT LONG SHORT LONG T F S F S F S F S HEATING RATE A L A L A L A L S O S O S O S O CORRODENT LEVEL T w T w LOW HIGH T w T w LOW HIGH 0/ 1 6 5 0 J H K \ 2 0 0 0 2 1 0 0 A 1 2 2 0 0 B p.

Cn w CYCLIC OXIDATION BURNER RIG TEST SPECIMEN SET FOR CONDITIONS D l , D2, E & F-12 SPECIMENS PER TEST 4 1 0 MIL VIRGIN CERAMIC ("BASELINE" COATING) 2 5 MIL VIRGIN CERAMIC 2 1 5 MIL VIRGIN CERAMIC 2 1 0 MIL AIR PRE-EXPOSED FOR APPROXIMATELY '/2 ESTIMATED BURNER RIG HOT TlME LlFE 2 10 MIL Ar PRE-EXPOSED FOR APPROXIMATELY Y i ESTIMATED BURNER RIG HOT TlME LlFE CYCLE LENGTH

SHORT: 6 MINUTE CYCLE = 4 MINUTES IN THE FLAME + 2 MINUTES FORCE AIR COOLED

LONG: 6 0 MINUTE CYCLE = 57 MINUTES IN THE FLAME + 3 MINUTES FORCE AIR COOLED

CYCLE RATE FAST: NOMINAL 6 0 SECOND HEAT-UPTO MAXIMUM TEMPERATURE SLOW: NOMINAL 180 SECOND HEAT-UP TO MAXIMUM TEMPERATURE CORRODENT LEVEL LOW: 1 0 PPM Naz So4 HIGH: 3 5 PPM Naz SO, Figure 3 Task I Test Plan To Evaluate Thermal Barrier Coating Failure i i fe h furnace life Log time to coating failure A 10 hour cycle 1 hour cycle \ \ TBC life

\ '

high cyclic content component TBC life low cyclic content / component C lnterface temperature F i g u r e 4 T e s t Data Showing C o a t i n g L i f e Dependent on Temperature, " C y c l i c Content" TBC "Quasistatic" furnace exposure, argon (80 hour cycle) no failure Log time to coating failure lnterface temperature F i g u r e 5 T e s t Data Showing Thermal Exposure Atmosphere E f f e c t s on C o a t i n g Durabi l i ty pre-exposed in ARGON for approximately one-half the Log time to coating failure lnterface temperature F i g u r e 6 T e s t Data Showing "INERT" Pre-Exposure Does Not E f f e c t C o a t i n g Performance preexposed in AIR for approximately onehalf the anticipated burner rig hot time

,p\\\

Log time to , o', 0 Preexposure \ coating failure duration L lnterface temperature F i g u r e 7 T e s t Data Showing A i r Pre-Exposure Degrades Cycl i c L i f e O R I G I N S PAGE IS OF POOR QUALITY Microstructural Variations for Pre-Test Thermal Exposure Figure 8a and D Atmospheres, ( a ) Argon and ( b ) Air Log time to coating failure burner rig life Interface temperature F i g u r e 9 T e s t D a t a Showing Ceramic Thickness E f f e c t s CONCLUDING REMARKS: FOURTH ANNUAL WORKSHOP FOR THE HOST PROJECT Daniel E. Sokolowski NASA Lewis Research Center Cleveland, Ohio As reported at this workshop, the HOST Project activities are well underway and are producing substantial results. Many contractor reports have become available and I will continue in the future. Workshops such as this also will continue on an annual basis. The HOST Fifth Annual Workshop is tentatively scheduled for October 21-22, 1 1986.

The nature of the problem of durability requires not only the involvement of numerous disciplines, as discussed in the opening remarks, but also that the research itself be interdisciplinary. The HOST Project to date has been very successful. The success is due in part to unprecedented teamwork at Lewis, at the contractors, and between contractors and universities. Further, the HOST Project is recognized for the value of focused as well as interdependent research when compared with generic, independent, basic R&T activities. The problem being addressed has much influence in the advocation and successful implementation of such a project, however.

To date, $27.3 million has been invested in HOST. In FY 1986 another $5.1 million will be spent. The present plan for FY 1987-89 is for a total of $12.3 million to be spent.

Finally, I want to say "thank you" to the HOST Project Team for a job well done in conducting this workshop. In particular, I want to thank the contractor speakers; the Subprojet Managers for being session chairman; my Assistant Manager, Bob Ensign; and our Project Control Assistant, Barbara DiSanto, for helping to organize and coordinate the multitude of efforts required.

PRECEDING PAGE BLANK NOT FILM83

45 7 - - 3. Recipient's Catalog No.

2. Government Accession No.

1. Report No.

NASA CP-2405 5. Report Date 4. Title and Subtitle O c t o b e r 1985 T u r b i n e E n g i n e H o t S e c t i o n Technology 1985 6. Performing Organization Code 533-04-1 2 8. Performing Organization Report No.

7. Author(s) E-2727 10. Work Unit No.

i 9. Performing Organization Name and Address 11. Contract or Grant No.

N a t i o n a l A e r o n a u t i c s and Space A d m i n i s t r a t i o n Lewis Research C e n t e r

13. Type of Report and Period Covered i C l e v e l a n d , Ohio 44135

12. Sponsoring Agency Name and Address C o n f e r e n c e P u b l i c a t i o n N a t i o n a l A e r o n a u t i c s and Space A d m i n i s t r a t i o n 14. Sponsoring Agency Code Washington, D . C . 20546 15. Supplementary Notes 16. Abstract On O c t o b e r 22 and 23, 1985, t h e T u r b i n e E n g i n e S e c t i o n T e c h n o l o g y (HOST) P r o j e c t O f f i c e o f t h e NASA Lewis Research C e n t e r sponsored a workshop t o d i s c u s s c u r r e n t r e s e a r c h p e r t i n e n t t o t u r b i n e e n g i n e h o t s e c t i o n d u r a b i l i t y p r o b l e m s . P r e s e n t a - t i o n s were made c o n c e r n i n g t h e h o t - s e c t i o n e n v i r o n m e n t and t h e b e h a v i o r o f com- b u s t i o n l i n e r s , t u r b i n e b l a d e s , and t u r b i n e vanes. The p r e s e n t a t i o n s were d i v i d e d i n t o s i x s e s s i o n s e n t i t l e d I n s t r u m e n t a t i o n , Combustion, T u r b i n e Heat T r a n s f e r , S t r u c t u r a l A n a l y s i s , F a t i g u e and F r a c t u r e , and S u r f a c e P r o t e c t i o n . The p r i n c i p a l o b j e c t i v e o f each s e s s i o n was t o d i s s e m i n a t e r e s e a r c h r e s u l t s t o d a t e .

T o p i c s d i s c u s s e d i n c l u d e d m o d e l i n g o f t h e r m a l and f l u i d - f l o w phenomena, s t r u c - t u r a l a n a l y s i s , f a t i g u e and f r a c t u r e , s u r f a c e p r o t e c t i v e c o a t i n g s , c o n s t i t u t i v e b e h a v i o r o f m a t e r i a l s , s t r e s s - s t r a i n response, and l i f e - p r e d i c t i o n methods.

R e s e a r c h e r s f r o m i n d u s t r y , academia, and government p r e s e n t e d r e s u l t s o f t h e i r work sponsored by t h e HOST P r o j e c t . T h i s p u b l i c a t i o n c o n t a i n s e x t e n d e d a b s t r a c t s and v i s u a l m a t e r i a l p r e s e n t e d d u r i n g t h e workshop.

I / 117. Key Words (Suggested by Author@)) Turbine Engine Hot Section Technology (HOST); L i f e ; prediction; Canbustor 1 i ners; Turbine a i r f o i 1s; I Material behavior; A i r c r a f t engines; D u r a b i l i t y

1 analysis methods 1 STAR C a t e g o r y 39

I 22. Price* 21. No. o f pages 19. Security Classif. (of this report) 20. Security Classif. (of this page) A21 468 U n c l a s s l f l e d U n c l a s s i f i e d o or sale by the National Technical Information Service, Springfield, Virginia 22161