NASA/TM—2010-215669

Development and Life Prediction of Erosion

Resistant Turbine Low Conductivity Thermal

Barrier Coatings

Dongming Zhu, Robert A. Miller, and Maria A. Kuczmarski

Glenn Research Center, Cleveland, Ohio

February 2010
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NASA/TM—2010-215669

Development and Life Prediction of Erosion

Resistant Turbine Low Conductivity Thermal

Barrier Coatings

Dongming Zhu, Robert A. Miller, and Maria A. Kuczmarski

Glenn Research Center, Cleveland, Ohio

Prepared for the

65th Annual Forum and Technology Display

sponsored by the American Helicopter Society

Grapevine, Texas, May 27–29, 2009

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

February 2010
Acknowledgments

The work was supported by NASA Fundamental Aeronautics Program (FAP) Subsonic Rotary Wing (SRW) project. The

authors are grateful to Mr. Michael Perez for performing the burner rig erosion tests.

This work was sponsored by the Fundamental Aeronautics Program

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management.

Available from

NASA Center for Aerospace Information National Technical Information Service

7115 Standard Drive 5301 Shawnee Road

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Available electronically at http://gltrs.grc.nasa.gov
Development and Life Prediction of Erosion Resistant Turbine

Low Conductivity Thermal Barrier Coatings

Dongming Zhu, Robert A. Miller, and Maria A. Kuczmarski

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Future rotorcraft propulsion systems are required to operate

under highly-loaded conditions and in harsh sand erosion

environments, thereby imposing significant material design

and durability issues. The incorporation of advanced thermal

barrier coatings (TBC) in high pressure turbine systems

enables engine designs with higher inlet temperatures, thus

improving the engine efficiency, power density and reliability.

The impact and erosion resistance of turbine thermal barrier

coating systems are crucial to the turbine coating technology

application, because a robust turbine blade TBC system is a

prerequisite for fully utilizing the potential coating technology

benefit in the rotorcraft propulsion. This paper describes the

turbine blade TBC development in addressing the coating

impact and erosion resistance. Advanced thermal barrier

coating systems with improved performance have also been

validated in laboratory simulated engine erosion and/or

thermal gradient environments. A preliminary life prediction

modeling approach to emphasize the turbine blade coating

erosion is also presented.

Introduction

Advanced thermal barrier coatings (TBC) are critical for

designing next generation rotorcraft turbine engines because of

their ability to allow increased engine gas temperatures and

reduced cooling requirements, thus helping to achieve improved

engine performance and durability. Multicomponent, oxidedefect-cluster based low conductivity thermal barrier coatings,

combined with improved single crystal Ni-base superalloys,

have demonstrated feasibility for increasing gas turbine engine

blade temperature capability and reducing cooling requirements

(Refs. 1 to 4). A recent systems assessment has shown that

advanced turbine thermal barrier coatings can provide

significant benefits in increasing engine efficiency (Ref. 5).

However, a prime-reliant coating design approach is necessary

to take full advantage of the thermal barrier coating technology

in future rotorcraft engine systems. In particular, turbine blade

thermal barrier coatings are especially susceptible to engine

ingested sand erosion and impact damages because of the high

velocity flow path nature of the blades and complex particulatecoating interactions under high heat-flux combustion turbine

environments. Although significant advances have been made

in thermal barrier coating technologies for reduced thermal

conductivity, increased temperature capability and improved

cyclic durability, erosion and impact resistance improvements

of the coating systems have remained one of the most

significant challenges for the turbine thermal barrier coating

developments and applications. Advanced turbine blade thermal

barrier coatings are currently being developed under the NASA

Rotary Wing Project aimed at increasing the coating toughness

and erosion resistance, in order to meet future engine design and

performance requirements. Extensive research efforts have been

made in developing new coating compositions, coating

processing, and erosion mechanism models in order to improve

the turbine coating erosion and impact durability (Refs. 6 and

7). The main objective of this effort is to develop higher

toughness turbine TBC systems with increased erosion and

impact resistance for rotorcraft propulsion engines. The

performance of advanced turbine blade thermal barrier coating

systems, designed based on nano-tetragonal phase toughening

approaches, will be discussed. These are based on the newly

established erosion burner rig testing in simulated sand

ingestion environments. Emphasis will also be placed on the

more engine relevant long-term, high-temperature smaller

particulate erosion conditions, so the erosion mechanisms can

be investigated for developing the turbine blade coating life

models.

Experimental Procedures

Advanced Thermal Barrier Coatings

Advanced ceramic thermal barrier coatings were designed

using a multicomponent defect-clustering approach to achieve

the required low conductivity, high-temperature stability and

improved durability (Refs. 1 to 3). In this study, the high

temperature erosion and impact performance of oxideclustering thermal barrier coating systems including ZrO2 -

Y2O3 -Gd2O3 -Yb2O3

 (t’ ZrYGdYb) and ZrO2-Y2O3 -Gd2O3 -

Yb2O3 -TiO2 -Ta2O5 (t’ ZrYGdYbTT) was investigated. The

advanced TBC systems, typically consisting of a 130 to

200 μm thick ceramic top coat and a 75 to 100 μm thick PtAl

intermediate bond coat, were electron beam-physical vapor

deposited (EB-PVD) on to 25.4 mm diameter disk Rene N5

nickel-base superalloy substrates. The coating compositions

were chosen based on nano-tetragonal phase toughening

design criteria and previous, preliminary phase

stability/durability test results, aiming to improve the erosion

resistance of the low conductivity thermal barrier coatings.

The advanced t’ coatings were also compared with the

NASA/TM—2010-215669
baseline ZrO2-7wt%Y2O3 (7YSZ), and cubic phased coatings

ZrO2-4mol%Y2O3 -3mol%Gd2O3 -3mol%Yb2O3 (cubic 433

ZrYGdYb), ZrO2-4mol%Y2O3 -3mol%Gd2O3 -3mol%Yb2O3 -

TiO2 -Ta2O5

 (cubic 433 ZrYGdYbTT), ZrO2 -4mol%Y2O3 -

4mol%Gd2O3 -4mol%Yb2O3 (cubic 444 ZrYGdYb), and

pyrochlore (Zr2Gd2O7) coatings. The EB-PVD coatings were

all deposited using pre-fabricated evaporation ingots made

using the designed compositions.

Erosion Burner Rig and Testing

A Mach 0.3 to 1.0 high velocity burner rig with erosion

capability was developed for advanced turbine blade thermal

barrier coating testing in a laboratory simulated turbine engine

relevant erosion environments. Early work has demonstrated

the burner rig effectiveness in evaluating the erosion behavior

of ceramic seal coats and thermal barrier coatings (Refs. 8 and

9). The rig development in this work has emphasized

increasing the erodent particle velocities at high temperatures

by increased gas mass flow rate, erodent flow uniformity with

improved feeding systems and a duct approach, and more

relevant thermal gradient turbine environments.

As can be seen in Figure 1, the high velocity burner erosion

rig consists of a Jet fuel combustor, a 19 mm (3/4 in.) diameter

highly efficient burner nozzle, a specimen holder fixture and

precision erodent feeder, and jet fuel and preheated

compressed air supplies. The burner nozzle inner surface

contour was designed based on an ANSI/ASME nozzle

standard to achieve better flame stability and uniformity

(Ref. 10), with the overall configuration modified for

accommodating increased burner mass flow and higher heat

flux environments. The nozzle was made of a single crystal

nickel-base superalloy turbine blade material to ensure high

temperature durability.

To achieve high erosion particle velocities, a computational

fluid dynamics modeling (CFD) approach in conjunction with

experimental investigations was employed to optimize the

burner and erodent injection design. The CFD model (Fluent)

was used to calculate the gas and particle velocities for the

burner rig temperatures and pressures of interest. The

modeling and experimental testing were conducted in relative

burner velocities ranging from Mach 0.3 to 0.9, representative

to turbine engine conditions. Al2O3 particles were used as

erodent and the particle sizes used were ranged from 27 to

560 μm in order to understand the coating broad erosion and

impact behavior. A modified burner rig configuration with an

extension duct (300 mm in length and 19 mm inner diameter)

was also explored to further accelerate the erodent particle and

increase the velocity.

The coating erosion evaluations was mostly conducted at

1800 °F (coating surface ~1900 °F) and 2000 °F (coating

surface 2100 °F), respectively, using 27 μm size Al 2O3

particles, but also compared with tests performed at 2200 °F

using 50 and 560 μm size Al2O3 particles. The erosion rate

was determined by interrupted coating erosion thicknessrecession and weight-loss measurements, and also by the

erodent amount required for the coating erosion penetration.

NASA/TM —2010-215669
Experimental Results and Discussion

Burner Rig Modeling and Characterization

Figure 2(a) shows the particle velocities modeled for a

27 μm erodent size using the CFD analysis in Mach 0.3 and

0.5 burner conditions, respectively. It can be seen that the

particle velocity increases with the distance from the burner

nozzle exit and burner gas velocity (Mach number). At 25 mm

(1 in.) distance, the particle velocity was estimated at 105 and

200 m/s, respectively, under the Mach 0.3 and Mach 0.5

burner gas flows. The maximum velocity is reached at 0.1 m

distance from the burner nozzle exit. The modeling is in a

good agreement with the experimental measurements, as show

in Figure 2(b) and (c).

For larger erodent particle size cases, lower particle

velocities are expected at a given burner Mach number and

longer acceleration distance is needed. As can be seen from

Figure 3 CFD modeling results, the velocities are

approximately 20 and 100 m/s at 25 mm distance from the

nozzle injector for the 50 and 560 μm particle sizes,

respectively under burner Mach 0.4 condition. The maximum

velocities can be reached in roughly 0.2 m from the nozzle

exit for the large particle size.

NASA/TM—2010-215669
NASA/TM —2010-215669
Thermal Barrier Coating Erosion Behavior

Figure 5 shows the typical erosion rate behavior of the

thermal barrier coatings at high temperatures. It can be seen

that the coatings usually have a higher initial erosion rate,

followed by a slower steady-state erosion rates. The decreased

steady-state erosion is due to some coating sintering and

densification of the very surface coating layers that have

higher erosion resistance. The coating erosion rates become

faster at the third stage when the coatings are at near the

failure stage.

The coating erosion behavior of various turbine thermal

barrier coatings are summarized in Figure 6, as expressed as

erosion rate (defined here as 25 μm coating thickness

recession per gram erodent required). It can be seen that

erosion rates of the coatings typically had some variability,

largely influenced by coating composition and processing

variations, and especially for the new multicomponent coating

systems. The baseline 7YSZ coatings had more consistent

erosion behavior.

In order to further increase the particle velocities and

uniformity, the burner rig was also modified with an

additional duct as shown in Figure 1 (b). The CFD model

predicts that the particle velocities can be increased from

150 m/s to over 250 m/s for the 27 μm particles under Mach

0.5 conditions. The particle velocities were also confirmed by

experiments measurement, as shown in Figure 4(b).

As compared to the baseline 7YSZ coating, the cubic

phased coating systems (cubic 433-ZrYGdYb, cubic 433-

ZrYGdYb 5-5TiTa, and pyrochlore based coatings) showed

the fastest erosion rates. Due to higher yield strength and

therefore lower toughness, it would be expected that the cubic

phase coatings had the lowest erosion rates. The lower phase

stability coating with t’ + monoclinic phase ZrRETT due to

the lower Gd, Yb dopant additions in the presence of Ti and

Ta also showed higher erosion rates. The well processed low

conductivity turbine t’ coating systems (both t’ ZrYGdYb and

ZyYGdYbTT) with reasonably controlled composition

homogeneity) generally showed improved erosion resistance

(lower erosion rates) as compared to the baseline 7YSZ in

both 27 and 50 μm size erosion and 560 μm impact tests.

Advanced processed low k coatings have shown significantly

improved erosion resistance.

The test results showed that the coating erosion rate

generally increases with erodent particle size tested in the

Mach 0.5 to 0.7 range. For the 27 μm size particles, the

coating erosion rates decreased dramatically in higher

temperature tests (coating temperature 2100 °F) compared to

the lower temperature (coating temperature 1900 °F). The

significantly increased plasticity of the t’ based coatings at the

higher temperature can contribute to the increased erosion

resistance.

NASA/TM —2010-215669
Figure 7 shows the particle energy dependence of the

erosion/impact rates, for 50 and 560 μm size erodent particles

tested at 2200 °F (1204 °C). The ln (erosion-impact rates) - ln

(particle energy) plot allows the determination of the energy

exponent of the erosion/impact rates. It can be seen that the

high temperature data showed relatively low energy

dependence. The average slope of the fitted lines for the

selected coating systems is found to be approximately

0.21 ±0.04. It should be noted that the two particle size cases

represent two distinct coating damage and failure mechanism

regimes (with the 50 μm size particles as the erosion dominant

mechanism and 560 μm size particles as the impact damage

mechanism). Nevertheless, it does provide some information

on the complexity of the coating general erosion behavior.

Figure 8 shows the erosion and impact surface morphologies

of low conductivity tetragonal phase t’ ZrYGdYb coatings. It

can be seen from Figure 8(a) that, in the 50 μm particle

erosion case, the ceramic coating is characterized by minor

surface plastic deformation and densification resulting from

the particle impingement. The coating erosion is primarily

caused by small-area and shallow-depth spalling under the

repetitive erodent particle loading. However, in the 560 μm

particle impact case as shown in Figure 8(b), the coating

surface typically experiences significant plastic deformation

and densification. It can be concluded that the high

temperature plastic deformation or plasticity can lead to

increased coating toughness and the energy absorption for the

560 μm size particles impact case, therefore leading to the low

energy exponent for the coating systems observed in this

study. Severe surface plasticity observed for the 560 μm size

coating impact testing case supports the hypothesis of the

increased coating toughness and reduced erosion/impact rates

for the large particles at the high temperatures. It should be

noted that large-area and deep-coating layer spallation can

occur especially at the late stage of the impact process due to

the increased coating delamination driving force and

accumulated damage under high impact energy conditions.

Thermal Barrier Coating Erosion Life Prediction

The erosion modeling is focused on the effect of the

underlying thermal barrier coating properties on the high

temperature erosion behavior. It has been shown that the TBC

erosion modeling and life models can be achieved by strain

amplitudes of various degradation mechanisms in conjunction

with fracture mechanics approaches.

The strain amplitude model is given by

E(AE p,eroson , AE p,creep , AE p,heat_flux , AE p, LCF , AE p, oxidation)

1/2

(1)

.Nf

 =Ec

Where AEp,erosion is the average repetitive particulate erosion

induced strain amplitude per particulate impact, AEp,creep,

AEp,heat_flux, AEp,LCF, and AEp,oxidation are coating sintering-creep,

NASA/TM—2010-215669
heat flux strain gradient and applied low cycle fatigue and

oxidation plastic strain induced strain amplitudes,

respectively. The thermal barrier coating erosion rate

Aep,erosion e p,erosion , defined as the mass loss from the coating

per unit mass erodent, can be determined from the erodent

particle kinetic energy and coating deformation volume based

on Equation (1)

e p,erosion = merodent' C' f(a )' Ve

n

rodent (2a)

p TBC ' p erodent '

e (2b ) p,erosion = K'

1/2 Vn

e

2

/2

c

 ' H(TY

Where merodent is the erodent mass, f(a) is the impingement

angle function, C, K are constants, n is the erodent velocity

exponent, typically n = 2–3 depending on the extent of

deformation occurred, p TBC and p erodent are the densities of the

thermal barrier coating and erodent, respectively, and V is

erodent velocity, H(T) is the temperature dependent dynamic

hardness of the coating. Given the coating defect size can be

typically assumed on the magnitude of the column width w,

the Equation (2b) can be rewritten in terms of the coating

toughness

e 

= K . p TBC ' w' ETBC ' p erodent 1/2 ' V

n

3 p,erosion 1

G ' H(T

)3/2 ( )

Where Gc is the critical strain energy release rate of the

coating. It can be seen that the erosion rate can be reduced by

increasing the coating toughness Gc, which can be achieved by

the coating compositional design and processing

improvements. Since the sintered coating hardness decreases

significantly with temperature based on our previous

experimental and modeling work, as shown in Figure 9, the

erosion rate temperature dependence will be governed by two

competing factors, that is, the erosion rate increases with the

overall coating softening (which increases the coating erosion

rate) and decreases with increased coating toughness due to

the increased plastic deformation (which reduces the coating

erosion rate). As mentioned above, the significantly reduced

erosion rates were observed at higher testing temperature

(2100 °F) as compared to those at lower testing temperature

(1900 °F). The reduction in erosion rates was attributed to the

predominant plasticity occurring at the higher temperature and

therefore increased coating toughness. The preliminary

coating development work demonstrated that the coating

toughness improvements through new compositional and

microstructural design and processing optimization that can

increase the erosion resistant up to 80 to 100 percent at high

temperatures. The coating erosion rate and life prediction will

focus on characterizing the effective coating toughness and

microstructural level coating properties to establish physicsbased models.

Concluding Remarks

An advanced high velocity burner rig based erosion test

approach has been established to evaluate the turbine thermal

barrier coatings in relevant engine environments. CFD models

have been established to understand gas and erodent flows for

the burner rig conditions.

A new series of t’ phase rare earth oxide (Gd2O3 and Yb2O3

codoped)- and TiO2/Ta2O5-alloyed, ZrO2-based thermal

barrier coatings were designed and processed. The advanced

turbine thermal barrier coatings developed for rotorcraft

engines will have combined low conductivity and high

toughness for improved thermal barrier performance and

erosion resistance. The coating systems demonstrated initial

improvements in the erosion resistance. The long-term

durability of the turbine airfoil thermal barrier coating systems

will be evaluated to develop comprehensive physics-based life

prediction models addressing the erosion, sintering and fatigue

interactions and strain damage accumulations, and to validate

the models for turbine blade components.

References

1. Dongming Zhu, and Robert A. Miller, “Low Conductivity

and Sintering Resistant Thermal Barrier Coatings,” US

Patent No. 6,812,176, 2004; US Patent No. 7,001,859,

2006; US Patent No. 7,186,466, 2007.

2. Dongming Zhu, and Robert A. Miller, “Thermal

Conductivity and Sintering Behavior of Advanced

NASA/TM —2010-215669
Thermal Barrier Coatings,” Ceram. Eng. Sci. Proc., vol.

23 (2002), pp. 457–468.

3. Dongming Zhu and Robert A. Miller, “Development of

Advanced Low Conductivity Thermal Barrier Coatings,”

International Journal of Applied Ceramic Technology,

vol. 1 (2004), pp. 86–94.

4. Dongming Zhu et al., “Thermal and Environmental

Barrier Coating Development for Advanced Propulsion

Engine Systems,” AIAA–2007 –2130, 48th

AIAA/ASME/ASCE/AHS/ASC Structures, Structural

Dynamics, and Materials Conference, 23–26 April 2007,

Honolulu, Hawaii.

5. Michael T. Tong, Scott M. Jones, and Philip C. Arcara,

Jr., “A Probabilistic Assessment of NASA Ultra-Efficient

Engine Technologies for a Large Subsonic Transport,”

Proceedings of Turbo Expo 2004, GT2004–53485.

6. Dongming Zhu and Robert A. Miller, “Low Conductivity

and High Toughness Tetragonal Phase Structured

Ceramic Thermal Barrier Coatings,” US Provisional

Patent Application No.: 60/712,613, August 26, 2005.

US Patent Application No.: 11/510,574, August 28, 2006.

7. Dongming Zhu, Maria A. Kuczmarski, Robert A. Miller,

and Michael D. Cuy, “Evaluation of Erosion Resistance

of Advanced Turbine Thermal Barrier Coatings,” The

31 st International Cocoa Beach Conference & Exposition

on Advanced Ceramics & Composites, Daytona Beach,

January 22 –26, 2007.

8. ANSI/ASME MFC-3M-2004, “Measurement of Fluid

Flow in Pipes Using Orifice, Nozzle, and Venturi,

Includes Addenda A,” American Society of Mechanical

Engineers/01-Jan-2004/92 pages ISBN: 0791829286.

9. Robert F. Handschuh, “High Temperature Erosion of

Plasma-Sprayed Yttria-Stabilized Zirconia in a Simulated

Turbine Environment,” NASA TP 2406; AVSCOM TR84-C-17, December 1984. Also AIAA–85–1219.

10. Robert W. Bruce et al., “Development of 1232 °C

(2250 °F) Erosion and Impact Tests for Thermal Barrier

Coatings,” Tribology Transactions, vol. 44 (1998),

3990410.

NASA/TM—2010-215669
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13. SUPPLEMENTARY NOTES

14. ABSTRACT

Future rotorcraft propulsion systems are required to operate under highly-loaded conditions and in harsh sand erosion environments, thereby

imposing significant material design and durability issues. The incorporation of advanced thermal barrier coatings (TBC) in high pressure

turbine systems enables engine designs with higher inlet temperatures, thus improving the engine efficiency, power density and reliability.

The impact and erosion resistance of turbine thermal barrier coating systems are crucial to the turbine coating technology application,

because a robust turbine blade TBC system is a prerequisite for fully utilizing the potential coating technology benefit in the rotorcraft

propulsion. This paper describes the turbine blade TBC development in addressing the coating impact and erosion resistance. Advanced

thermal barrier coating systems with improved performance have also been validated in laboratory simulated engine erosion and/or thermal

gradient environments. A preliminary life prediction modeling approach to emphasize the turbine blade coating erosion is also presented.

15. SUBJECT TERMS

Thermal barrier coatings; Erosion; Life prediction; Turbine engine; Computational fluid dynamics; Burner rig testing

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