Thermal and Environmental Barrier Coating

Development for Advanced Propulsion

Engine Systems

NASA/TM—2008-215040

January 2008

Dongming Zhu

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

Robert A. Miller and Dennis S. Fox

Glenn Research Center, Cleveland, Ohio

ARL–TR–4368

AIAA–2007–2130 U.S. ARMY

RESEARCH LABORATORY
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Thermal and Environmental Barrier Coating

Development for Advanced Propulsion

Engine Systems

NASA/TM—2008-215040

January 2008

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Dongming Zhu

U.S. Army Research Laboratory, Glenn Research Center, Cleveland, Ohio

Robert A. Miller and Dennis S. Fox

Glenn Research Center, Cleveland, Ohio

ARL–TR–4368

AIAA–2007–2130 U.S. ARMY

RESEARCH LABORATORY

Prepared for the

48th Structures, Structural Dynamics, and Materials Conference

sponsored by the AIAA, ASME, ASCE, AHS, and ASC

Waikiki, Hawaii, April 23–26, 2007
Acknowledgments

This work was supported by NASA Ultra-Efficient Engine Technology (UEET) Program and NASA Fundamental Aeronautics

(FA) Program Supersonics Project.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22161

Available electronically at http://gltrs.grc.nasa.gov

Level of Review: This material has been technically reviewed by technical management.
NASA/TM—2008-215040 1

Thermal and Environmental Barrier Coating Development for 

Advanced Propulsion Engine Systems 

Dongming Zhu 

U.S. Army Research Laboratory 

Glenn Research Center 

Cleveland, Ohio 44135 

Robert A. Miller and Dennis S. Fox 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to 

protect engine hot-section components in the harsh combustion environments, and extend component 

lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, better thermal stability 

and higher toughness than current coatings will be beneficial for future low emission and high 

performance propulsion engine systems. In this paper, ceramic coating design and testing considerations 

will be described for turbine engine high temperature and high-heat-flux applications. Thermal barrier 

coatings for metallic turbine airfoils and thermal/environmental barrier coatings for SiC/SiC ceramic 

matrix composite (CMC) components for future supersonic aircraft propulsion engines will be 

emphasized. Further coating capability and durability improvements for the engine hot-section component 

applications can be expected by utilizing advanced modeling and design tools. 

Introduction 

Ceramic TEBCs are considered technologically important because of their ability to increase gas 

turbine engine operating temperatures and reduce cooling requirements, thus help achieving engine 

performance goals. A recent system assessment has shown that the low conductivity turbine coatings 

provide significant benefits in increasing engine efficiency and reducing NOX emissions for various 

classes of vehicles among the major engine technologies considered (refs. 1 and 2). The advantages of 

using TEBCs include higher engine efficiency by increasing gas temperatures, and improved engine 

reliability by reducing engine hot-section component temperatures. The development of advanced 

ceramic barrier coatings is currently aimed at significantly increasing engine operating temperature and 

simultaneously reducing air cooling, in order to meet future engine lower emission, higher efficiency and 

improved reliability goals. 

For the supersonic vehicles currently envisioned in the NASA Fundamental Aeronautics Program 

(FA), advanced gas turbine engines will be used to provide high power density thrust during the extended 

supersonic flight of the aircraft, while meeting stringent emission requirements. Combustor exit 

temperatures during the supersonic cruise will lead to turbine inlet temperatures around 167 °C (300 °F) 

higher than the current state-of-the-art for subsonic engines. Because of the extended supersonic cruise 

time with very high inlet air temperature, the engine components will experience long-duration operation 

at high temperatures throughout the cruise portion of the mission. Improved ceramic coating systems are 

critical to the performance, life and durability of the hot-section components in the engine systems. The 

development of advanced ceramic thermal and environmental barrier coatings will enable high 

temperature capable, lightweight engine hot-section components for increased efficiency, reduced 

emission and long-term durability required by supersonics propulsion systems. 
NASA/TM—2008-215040 2

To take full advantage of the TEBC capability potential, an aggressive design approach—allowing 

greater temperature reductions through the coating systems and less cooling air to the components—is 

required whenever possible. The ceramic coating systems must be designed with increased high-temperature 

stability, lower thermal conductivity, greater toughness and improved thermal stress and erosion resistance. 

To achieve these performance characteristics, an advanced defect cluster coating design approach will be 

beneficial (refs. 3 to 7). As shown in figure 1, the development of revolutionary ceramic coatings, combined 

with improved single crystal Ni-base superalloy and CMC substrates, will result in a step increase in gas 

turbine engine blade, vane and combustor component temperature capability. 

Higher surface temperatures and larger thermal gradients as compared to conventional coating 

systems are required in advanced TEBC systems. Figure 2 is a schematic diagram showing how advanced 

ceramic barrier coatings with lower thermal conductivity and better temperature stability will allow the 

use of thinner coatings to achieve a larger temperature reduction in the higher operating temperature 

engines. 

2400°F (1316°C)

3000°F+ (1650°C+)

Gen I

Temperature 

Capability (T/EBC) surface

Gen II – Current commercial

Gen III

Gen. IV

Increase in ΔT 

across T/EBC

Single Crystal Superalloy

Year

Ceramic Matrix Composite

Gen I

Temperature 

Capability (T/EBC) surface

Gen II – Current commercial

Gen III

Gen. IV

Increase in ΔT 

across T/EBC

Single Crystal Superalloy

Year

Ceramic Matrix Composite

2700°F (1482°C)

2000°F (1093°C)

Step increase in temperature capability

3100°F SiC/SiC

CMC coatings

2700°F SiC/SiC CMC 

and Si3N4 coatings

2800oF 

combustor 

TBCs

2500oF 

Turbine 

TBCs

Figure 1.—The development of high temperature stability and low thermal conductivity ceramic coatings, 

will result in a step increase in the temperature capability of gas turbine hot-section components. The 

temperature design limits for various classes of materials are shown. 

Ceramic 

coating

Ceramic 

coating

Bond Substrate Substrate

coat

Bond 

coat

Tgas

Tsurface

Tgas

Tsurface

Tback

Tback

Ceramic 

coating

Ceramic 

coating

Bond Substrate Substrate

coat

Bond 

coat

Tgas

Tsurface

Tgas

Tsurface

Tback

Tback

Ceramic 

coating

Ceramic 

coating

Bond Substrate Substrate

coat

Bond 

coat

Tgas

Tsurface

Tgas

Tsurface

Tback

Tback

(a) (b) 

Figure 2.—Advanced thermal and environmental barrier coatings can significantly increase gas temperatures, reduce 

cooling requirements, and improve engine fuel efficiency and reliability. (a) Current TEBCs; (b) Advanced TEBCs. 
NASA/TM—2008-215040 3

Combustor Vane Blade Ceramic nozzles

Coated CMC 

combustor liner 

and vane

Figure 3.—Ceramic thermal and environmental barrier coatings are critical 

technologies in future gas turbine engine hot-sections for combustor, vane 

and turbine airfoil applications. 

Figure 3 shows examples of ceramic barrier coatings that are typically used for combustor, vane and 

turbine airfoil applications. The low conductivity coatings will have a significant advantage over the 

conventional ZrO2-(7-8)wt%Y2O3 coatings especially for rotating engine components (such as turbine 

blades), where a reduced weight is highly desirable. The multi-component doped, oxide alloy defect 

clustered ceramic TEBC has been shown to offer the low conductivity and high stability required for both 

high temperature subsonic and supersonic engine applications (refs. 3 to 7). Figure 3 shows the engine 

hot-section component (combustors, vanes and turbine airfoils) applications for the ceramic thermal and 

environmental barrier coatings. 

In this paper, ceramic coating design and testing considerations are described for engine hot-section 

component applications in high temperature and high-heat-flux, oxidizing and water vapor-containing 

combustion environments. Both thermal barrier coatings (TBCs) for metallic turbine blades, and TEBCs 

for SiC/SiC CMC combustor and turbine applications, will be emphasized especially for supersonic 

engines. Coating performance of advanced systems will be discussed. Further improvements will be 

obtained using new structural designs and composition optimization, in conjunction with modeling and 

design tools. 

Design Approach of Ceramic Defect-Clustering Coatings 

Advanced ZrO2- and HfO2-based oxide TBCs were designed using a multi-component defectclustering scheme (refs. 3 to 7). The advanced oxide coatings were designed by incorporating multicomponent, paired-cluster dopants in conventional zirconia- and hafnia-yttria oxides. The dopant oxides 

were selected based on the cation-anion interatomic and chemical potentials, lattice elastic strain energy, 

polarization and the electro-neutrality of the oxides. Since defect clusters can attenuate and scatter lattice 

phonon waves as well as radiative photon waves at a wide spectrum of frequencies, one can expect 

significant reduction in the oxide intrinsic lattice and radiation thermal conductivity. The creation of 

thermodynamically stable, highly defective lattice structures with controlled ranges of defect cluster sizes 

has been confirmed by high resolution transmission electron microscopy (TEM) (ref. 8). The presence of 

the highly distorted lattice structures, with essentially immobile defect clusters and/or nanoscale ordered 
NASA/TM—2008-215040 4

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

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

(a) 

(b) 

Figure 4.—High-resolution TEM lattice image showing the defect clusters by moiré patterns in ZrO2-based and 

HfO2-based top coating systems. (a) ZrO2 coating systems; (b) HfO2 coating system. 

phases, effectively reduce the mobile defect concentration and suppress the atomic mobility and mass 

transport, thus significantly improving the oxide sintering-creep resistance and mechanical properties. 

The defect clustering design is essential in reducing coating lattice and radiative thermal conductivity, and 

improving sintering-creep resistance. Improvements in mechanical properties such as fracture toughness 

and cyclic durability have been observed by the formation of complex nanoscale defect clusters and 

nanophase composites. Since a temperature gradient of 110 to 394 °C per 100 μm (50 to 100 °F per 

0.001 in.) coating thickness is important to advanced engines, the coatings must be designed with the 

optimal combination of performance characteristics required for the extended supersonics cruise mission. 

Figure 4 contains high resolution TEM lattice images of the defect clusters which form in the multi-oxide 

component ZrO2- and HfO2-based coatings. 

Thermal and environmental barrier coatings are being designed to protect SiC/SiC CMC combustor 

and turbine components by extending the CMC temperature capability to 1650 °C (3000 °F) in oxidizing 

and water vapor containing combustion environments (refs. 9 and 10). To achieve the engine performance 

goals, the coatings must provide the thermal and environmental protection of Si-based components at the 

gas temperatures in excess of 1760 °C (3200 °F), while maintaining the CMC substrate below its 

temperature limit, which is below 1316 °C (2400 °F). The coating consists of a multi-component zirconiaor hafnia-based oxide defect cluster top layer, a strain tolerant interlayer, a mullite/rare-earth silicate 

environmental barrier layer, and a ceramic oxide/rare earth silicate-mullite(+Si) bond coat layer. The 

thermal and environmental barrier coating system provides an ultimate protective coating solution for 

silicon-based ceramic components exposed to very high temperatures and/or high flow gas velocities in 

water vapor-containing combustion and erosion environments. The advanced defect cluster oxide top 

layer possesses low intrinsic thermal conductivity and an extraordinarily high thermal stability at 

temperatures up to 3100 °F (1704 °C). The nanoscale oxide defect clustering and the thermodynamicallystable interfaces induce phonon scattering, resulting in a reduction in thermal radiative thermal 

conductivity within the coating system. Furthermore, the advanced interlayers impart a strain-tolerant 
NASA/TM—2008-215040 5

High temperature capability 

thermal and radiation barrier

Energy dissipation and chemical 

barrier interlayer

Second radiation barrier and 

thermal control

Environmental barrier and bond coat

Ceramic matrix composite (CMC)

Tsurface 1650°C (3000°F)

Tcoating/CMC interface <1316°C (2400°F)

High temperature capability 

thermal and radiation barrier

Energy dissipation and chemical 

barrier interlayer

Second radiation barrier and 

thermal control

Environmental barrier and bond coat

Ceramic matrix composite (CMC)

High temperature capability 

thermal and radiation barrier

Energy dissipation and chemical 

barrier interlayer

Second radiation barrier and 

thermal control

Environmental barrier and bond coat

Ceramic matrix composite (CMC)

Tsurface 1650°C (3000°F)

Tcoating/CMC interface <1316°C (2400°F)

Tsurface 1650°C (3000°F)

Tcoating/CMC interface <1316°C (2400°F)

Figure 5.—Innovative multi-layer architecture which incorporates a high-temperature top 

coat, an advanced environmental barrier coat and a ceramic bond coat. 

buffer between the top layer and the substrate (which can accommodate large thermal gradient strains 

through the coating system), enabling long-term coating and component durability. Figure 5 illustrates the 

revolutionary multilayer coating structure concept used to provide protection of Si-based ceramic and 

ceramic matrix composite engine components. 

Testing of Ceramic Coatings 

The development of low conductivity and high temperature stable ceramic barrier coatings requires 

testing techniques that can accurately evaluate thermal conductivity at high surface temperatures in the 

range of 1300 to 1400 °C for thermal barrier coatings, and up to 1650 °C (3000 °F) for the more advanced 

applications. A unique steady-state CO2 laser (wavelength 10.6 μm) heat-flux test rig is used to determine 

the thermal conductivity and cyclic durability of TEBC systems at high temperatures and under large 

thermal gradients across the coating. A schematic diagram showing the laser thermal conductivity rig is 

shown in figure 6. The laboratory high-power CO2 laser heat-flux rig effectively simulates the engine 

heat-fluxes and thermal gradients. The laser heat-flux based testing approach can also be used to monitor 

the coating real-time thermal conductivity change kinetics, which is a critical aspect for the coating 

developments (refs. 6 and 7, and 9 to 15). The high velocity and high pressure burner rigs provide 

simulated engine gas flow, pressure and combustion environment conditions to evaluate the coating 

systems. The simulated combustion water-vapor environment, thermomechnical fatigue and erosion 

testing capabilities have also recently been incorporated into the laser heat-flux and burner rigs, as shown 

in figure 7. 
NASA/TM—2008-215040 6

reflectometer 10.6 micron 

laser beam

cooling air

specimen

3. 5 KW CO2 High Power Laser

ceramic coating

substrate

platinum flat coils

aluminum back plate

air gap

cooling air tube

TBC coated

back aluminum

plate edge

pyrometers

bond coat

camera

Figure 6.—Schematic diagram showing the laser heat-flux rig used to determine 

the thermal conductivity of ceramic TEBCs. During the test, the ceramic surface 

and the metal backside temperatures are measured by infrared pyrometers. The 

interfacial temperatures, and the actual heat-flux passing through the thermal 

barrier coating system, are therefore determined under the steady-state laser 

heating conditions by a one-dimensional (1-D) heat transfer model. 

Specimen under testing

Steam jets Steam jets

Specimen holder and water vapor jets

Laser beam delivery 

and optic system

Infrared 

pyrometer

Laser heat flux water vapor test rig

(a) 

Erosion nozzle

Laser beam Pyrometer

Strain gauge HCF rig

Specimen

Erosion nozzle

Laser beam Pyrometer

Strain gauge HCF rig

Specimen

(b) 

IR image 

showing the 

flame core

(c) (d) 

Figure 7.—Laser high-heat-flux and burner rig test rigs for simulated thermal gradient testing of thermal and environmental 

barrier coatings. (a) The laser heat-flux water vapor environment rig for advanced TEBC development; (b) Combined 

laser high heat flux, high cycle fatigue and erosion rig; (c) High velocity (Mach 0.3 to 0.7) burner erosion test rig. (d) High

pressure burner rig with 6 to 10 atm pressure and up to 3000 °F combustion gas environments. 
NASA/TM—2008-215040 7

Advanced Ceramic Coatings Development 

Ceramic Thermal Barrier Coatings 

Advanced ceramic barrier coatings were designed using a multi-component defect-clustering 

approach to achieve the required low conductivity, high-temperature stability and improved durability. 

Selected oxide-cluster thermal barrier coating systems including ZrO2-Y2O3-Nd2O3(Gd2O3,Sm2O3)-

Yb2O3(Sc2O3) were synthesized, and their thermal conductivity, sintering behavior and cyclic durability 

were investigated at high temperatures. The advanced TBC systems, typically consisting of a 127 to 

250 μm ceramic top coat and a 75 to 120 μm NiCrAlY or PtAl intermediate bond coat, were either 

plasma-sprayed or electron beam-physical vapor deposited (EB-PVD) on to 25.4 mm diameter disk 

substrates. The plasma-sprayed coatings were processed using pre-alloyed powders. The ceramic powders 

of selected compositions were first spray-dried, then plasma-reacted and spheroidized, and finally plasmasprayed as the coating form. The advanced EB-PVD coatings were deposited using pre-fabricated 

evaporation ingots made using the designed compositions. 

Due to the relatively porous nature of the thermal barrier coating, thermal conductivity can increase 

significantly due to coating sintering and/or phase structure changes after a long-term thermal exposure. 

Therefore, evaluation of the initial and post-exposure thermal conductivities, and the rate-of-conductivityincrease is crucial in characterizing the coating’s performance. Figure 8 shows the thermal conductivity 

kinetics of plasma-sprayed baseline ZrO2-8wt%Y2O3 thermal barrier coating, determined using the 

steady-state laser heat-flux testing. The coating shows faster initial conductivity increase, due to the fast 

microcrack sintering rate at the initial stage. 

Figure 8.—Thermal conductivity change kinetics of plasmasprayed baseline ZrO2-8wt%Y2O3 thermal barrier coatings, 

determined using steady-state laser heat-flux testing at 

moderate surface temperatures (pass-through heat flux 

64 W/cm2

). Because a thermal conductivity gradient is 

expected through thickness of the ceramic coating under the 

thermal gradient conditions, the observed ceramic thermal 

conductivity increase reflects the overall effect of the 

conductivity change in the coating. 
NASA/TM—2008-215040 8

After long-term testing, the coating conductivity reaches a plateau in the measuring conductivity. The 

coating thermal conductivity change kinetics for the baseline coating can be expressed as 

⎭

⎬

⎫

⎩

⎨

⎧

⎥

⎦

⎤ ⎢

⎣

⎡

τ

⎟ − − ⎠

⎞ ⎜

⎝

⎛ − = ⋅ − t

RT

k k . c c 1 exp

68228 102 2 exp 0 (1) 

⎟

⎠

⎞ ⎜

⎝

⎛

τ = ⋅ RT . 41710 572 5 exp (2) 

where kc is the coating thermal conductivity at any given time t, 0

c k and infc k are ceramic coating thermal 

conductivity values at the initial time and at infinitely long times, respectively, R is gas constant, and τ is 

relaxation time.

Figure 9 shows the sintering and creep strain kinetics of a baseline turbine airfoil ZrO2-7wt%Y2O3

thermal barrier coating, derived from the surface wedge crack-opening displacements after testing for 

various times with temperature gradients (ref. 13). It can be seen that the sintering and creep strains are on 

the magnitude 2 percent approximately at 1316 °C (2400 °F) following 1300 hr exposure. The low 

resistant to sintering and creep of the baseline coating can result in accelerated cracking and delamination 

during the cyclic operation condition. 

Figure 9.—Surface sintering and creep strains (%) of an EB-PVD turbine airfoil ZrO2-7wt%Y2O3 coating derived 

from laser high-heat-flux tests, using a heat flux of 220 W/cm2

 as functions of temperature and time. 
NASA/TM—2008-215040 9

Figure 10 illustrates thermal conductivity behavior of advanced multi-component plasma-sprayed and 

EB-PVD thermal barrier coatings as a function of test time at high temperatures. The advanced oxide 

coatings were comprised of primarily ZrO2-Y2O3, and co-doped with additional paired rare earth oxides 

Nd2O3-Yb2O3 or Gd2O3-Yb2O3 (i.e., ZrO2-(Y,Nd,Yb)2O3 and ZrO2-(Y,Gd,Yb)2O3 coating systems). These 

coatings were found to possess much lower thermal conductivity and better temperature stability than the 

conventional baseline ZrO2-7-8wt%Y2O3 coatings, due to the paired rare earth dopant addition and the 

defect cluster formation. From figure 11, it can be seen that depending on the test temperature and coating 

composition, a conductivity reduction of 50 to 66 percent is achieved in the advanced coating systems as 

compared to the ZrO2-8wt%Y2O3 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

Thermal conductivity, W/m-K

Time, hours

ZrO2-4.55mol%Y2O3 (ZrO2-8wt%Y2O3 )

ZrO2-13.5mol%(Y,Nd,Yb)2O3

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%Y2O3 (ZrO2-7wt%Y2O3 )

ZrO2-4mol%Y2O3 (ZrO2-7wt%Y2O3 )

Low conductivity ZrO2-10mol%(Y,Gd,Yb)2O3 coating

Thermal conductivity, 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

1316°C

0.0

0.5

1.0

1.5

2.0

2.5

0 5 10 15 20 25

ZrO2-4mol%Y2O3 (ZrO2-7wt%Y2O3 )

ZrO2-4mol%Y2O3 (ZrO2-7wt%Y2O3 )

Low conductivity ZrO2-10mol%(Y,Gd,Yb)2O3 coating

Thermal conductivity, 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

0.0

0.5

1.0

1.5

2.0

2.5

0 5 10 15 20 25

ZrO2-4mol%Y2O3 (ZrO2-7wt%Y2O3 )

ZrO2-4mol%Y2O3 (ZrO2-7wt%Y2O3 )

Low conductivity ZrO2-10mol%(Y,Gd,Yb)2O3 coating

Thermal conductivity, 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

1316°C

(a) (b) 

Figure 10.—Thermal conductivity of advanced multi-component plasma-sprayed and EB-PVD oxide cluster thermal 

barrier coatings, determined by laser heat-flux technique. (a) Plasma-sprayed coatings; (b) EB-PVD coatings. (a) 

Plasma-sprayed coating; (b) EB-PVD coating. 

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)

Thermal conductivity, W/m-K

Total dopant concentration, mol%

Plasma-sprayed 

coatings

Advanced

ZrO2-

based 

coatings

k0

k20

k0

k20

k0

k20

k0

k20

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)

Thermal conductivity, W/m-K

Total dopant concentration, mol%

Plasma-sprayed 

coatings

Advanced

ZrO2-

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)

Thermal conductivity, 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)

Thermal conductivity, W/m-K

Total dopant concentration, mol%

Plasma-sprayed 

coatings

Advanced

ZrO2-

based 

coatings

k0

k20

k0

k20

k0

k20

k0

k20

Figure 11.—Thermal conductivity of oxide-cluster thermal barrier coatings 

tested at high temperatures. Both tetragonal t’ phase coatings and cubic 

phase ZrO2 cluster coatings showed significantly lower thermal 

conductivity than did the baseline ZrO2-8wt%Y2O3 at higher 

temperatures. The k0 and k20 denote the conductivity values under asprocessed and 20 hr steady-state tested conditions, respectively. 
NASA/TM—2008-215040 10

The advanced low conductivity plasma-sprayed and EB-PVD coatings have demonstrated excellent 

cyclic durability and erosion resistance (ref. 11) with compositional and coating structural-design 

optimization. Figure 12 shows that both the advanced tetragonal phase composition coatings, and the 

cubic phase composition coatings with an interlayer system, achieved furnace cyclic oxidation durability 

comparable or better than the baseline ZrO2-8wt%Y2O3 coatings. Figure 13 shows the advanced low 

conductivity turbine airfoil and combustor thermal barrier coating demonstrated long-term simulated 

engine high-heat-flux, high-thermal-gradient cyclic durability. 

0

200

400

600

800

1000

4 5 6 7 8 9 10 11 12

Cubic phase low k TBC

Tetrogonal t' phase low k TBCs

8YSZ

Cycles to failure

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 TBC

Tetrogonal t' phase low k TBCs

8YSZ

Cycles to failure

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 TBC

Tetrogonal t' phase low k TBCs

8YSZ

Cycles to failure

Total dopant concentration, mol%

2075°F (1135°C)

Low-k t’-phase 

region Low-k cubic 

phase region

With the 

interlayer

1135°C

Figure 12.—Furnace cyclic oxidation behavior of plasmasprayed ZrO2-(Y,Gd,Yb)2O3 thermal barrier coatings. The 

cyclic tests were conducted using 45 min hot time with 15 min 

cooling in air at 1138 °C. The tetragonal t’-phase composition 

low conductivity coatings achieved at least the baseline ZrO2-

(7-8)wtY2O3 life with a single layer architecture. The cubicphase composition low conductivity TBC durability, although 

inferior to the baseline coatings in a single layer configuration, 

can be significantly improved by an initial ZrO2-8wtY2O3 or 

low k tetragonal t’-phase interlayer coating. 

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

Thermal conductivity, 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

(a) 

0.0

0.5

1.0

1.5

0 50 100 150 200 250

Thermal conductivity, 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 plasmasprayed combustor coating

(b) 

Figure 13.—An advanced low conductivity EB-PVD turbine airfoil and a plasma-sprayed combustor thermal barrier 

coatings demonstrated long-term cyclic durability in simulated engine high-heat-flux, high thermal gradient 

conditions at 1360 °C (2480 °F), 1650 °C (3000 °F), respectively. (a) EB-PVD low conductivity turbine airfoil 

thermal barrier coating; (b) Plasma-sprayed low conductivity combustor thermal barrier coating. 
NASA/TM—2008-215040 11

Ceramic Thermal and Environmental Barrier Coatings 

Multi-layerd and multi-component thermal and environmental barrier coatings based on HfO2 and 

rare earth mullite systems were developed for SiC/SiC CMC combustor liner and vane applications. The 

coating systems have shown improved phase stability, lower lattice and radiation thermal conductivity, 

and enhanced sintering and thermal stress resistance under the simulated engine heat-flux and thermal 

cycling conditions. This is largely due to the defect clustering structures which have been used to 

effectively promote the creation of the thermodynamically-stable oxide defect clusters and/or nano-phases 

within the coating systems (ref. 16), as compared to the baseline barium strontium aluminosilicate 

(BSAS) based environmental barrier coatings, i.e., BSAS/mullite+BSAS/Si, which has a temperature 

capability below 2400 °F (refs. 17 and 18). The multi-component coatings have also demonstrated longterm cyclic durability and good water vapor stability. Figure 14 illustrates the durability exhibited by the 

multi-component HfO2-18mol%(Y,Gd,Yb)2O3/Gd, Yb and Hf modified mullite/mulite+20%BSAS/Si 

EBC system, during the 1650 °C thermal gradient cycling testing. The coating system showed relatively 

low overall thermal conductivity, with no visual damage after the testing, demonstrating 300-hr long-term 

1650 °C sintering stability and cyclic durability on SiC/SiC ceramic matrix composite substrate under the 

thermal gradient cyclic testing conditions. As illustrated in figure 15, the 6 atm high pressure burner rig 

testing results indicated the excellent water vapor stability of the coating system in a simulated 

combustion environment. Figure 16 shows that a hybrid EB-PVD and plasma sprayed multi-component 

HfO2-based mullite-BSAS model turbine CMC coating system successfully performed at 3100 °F and 

100 thermal cycle laser high-heat-flux test (using 1 hr cycles). 

0.5

1.0

1.5

2.0

0 50 100 150 200 250 300 350

Thermal conductivity, W/m-K

Time, hours

Tsurface=3000°F

Tinterface=2400°F

Figure 14.—Multi-component HfO2-18mol%(Y,Gd,Yb)2O3/rare earth 

mullite/HfO2-based thermal and environmental barrier coating 

systems demonstrated 300 hr long-term 3000 °F sintering and cyclic 

durability on a SiC/SiC ceramic matrix composite substrate under 

thermal gradient cyclic testing. 
NASA/TM—2008-215040 12

-2.0

-1.0

0.0

1.0

2.0

3.0

0 10 20 30 40 50 60

Specific weigh change, mg/cm2

Time, hours

High Pressure Burner Rig Tested 

Gas temperature 2650°F, 6 atm

Gas velocity 25 m/sec

Specimen after testing 

Figure 15.—Multi-component, multi-layerd HfO2-(Y,Gd,Yb)2O3/rare earth oxide doped mullite/Si thermal and 

environmental barrier coating systems demonstrated excellent water vapor stability and durability in the high 

pressure burner rig testing environments. 

0.0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

0

400

800

1200

1600

0 20 40 60 80 100 120

kcera

Tsurface

Tinterface

Tback

Thermal conductivity, W/m-K

Temperature, °C

Time, hours

3100°F coatings

Interlayer

Top layer

EBC

Si bond coat

Interlayer

Top layer

EBC

Si bond coat

Cross-section micrograph 

Specimen after testing 

Figure 16.—An EB-PVD-plasma sprayed hybrid multi-component HfO2-based mullite-BSAS model turbine vane coating 

system successfully completed the total 100-1 hr cycle laser heat flux testing at 3100 °F (60 min hot, 3 min cool). 
NASA/TM—2008-215040 13

Concluding Remarks 

Thermal and environmental barrier coatings will play a crucial role in advanced gas turbine engines 

because of their ability to further increase engine operating temperature and reduce cooling requirements, 

thus help contributing to achieve engine emission and efficiency goals. Future coating systems must be 

designed with increased phase stability, lower thermal conductivity, and improved sintering and thermal 

stress resistance to effectively protect engine hot-section metallic and ceramic components. Advanced 

low conductivity TEBCs have been developed by incorporating multi-component oxide dopants into 

zirconia-yttria or hafnia-yttria to promote the formation of thermodynamically stable defect clusters 

within the coating structures. The advanced TEBC systems have demonstrated cyclic durability at 

temperatures up to 1650 °C which is far beyond the current state-of-the-art baseline ZrO2-Y2O3 thermal 

barrier and BSAS/mullite+BSAS environmental barrier coating capabilities. Modeling efforts are being 

pursued to understand the coating conductivity, sintering, and fracture, which can contribute to 

establishing design tools for the coating developments. 

References 

1. 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. 

2. Clayton L. Mayer, private communication. 

3. 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. 

4. Dongming Zhu, and Robert A. Miller, “Multi-functionally graded environmental barrier coatings for 

Si-based ceramic components,” Provisional patent application number: 60/712,605, August 26, 2005; 

US Patent Application No. 11/510 573, August 28, 2006. 

5. 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. 

6. Dongming Zhu, and Robert A. Miller, “Thermal Conductivity and Sintering Behavior of Advanced 

Thermal Barrier Coatings,” Ceramic Eng. Sci. Proc., vol. 23 (2002), pp. 457–468. 

7. 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. 

8. Dongming Zhu, Yuan L. Chen and Robert A. Miller, “Defect Clustering and Nano-phase Structure 

Characterization of Multicomponent Rare-Earth Oxide Doped Zirconia-Yttria Thermal Barrier 

Coatings,” Ceramic Eng. Sci. Proc., vol. 24 (2003). Also NASA TM-212480. 

9. Dongming Zhu, and Robert A. Miller, “Hafnia-Based Materials Development for Advanced 

Thermal/Environmental Barrier Coating Applications,” in Research and Technology 2003, NASA 

Glenn Research Center, NASA/TM—2004-212729, pp. 41–43. 

10. Dongming Zhu, Narottam P. Bansal, and Robert A. Miller, “Thermal conductivity and Stability of 

Hafnia- and Zirconate-Based Materials for 1650 °C Thermal/Environmental Barrier Coating 

Applications,” Ceramic Transactions, vol. 153 (2003). Also NASA TM-212544, 2003. 

11. Dongming Zhu, and Robert A. Miller, “Thermal Conductivity Change Kinetics of Ceramic Thermal 

Barrier Coatings Determined by the Steady-State Laser Heat Flux Technique,” in Research and 

Technology 1999, NASA Glenn Research Center, NASA TM-209639, pp. 29–31, March 2000. 

12. Dongming Zhu, and Robert A. Miller, “Thermal Conductivity and Elastic Modulus Evolution of 

Thermal Barrier Coatings Under High Heat Flux Conditions,” Journal of Thermal Spray Technology, 

vol. 9 (2000), pp. 175–180. 
NASA/TM—2008-215040 14

13. Dongming Zhu, Robert A. Miller, Ben A. Nagaraj, and Robert W. Bruce, “Thermal Conductivity of 

EB-PVD Thermal Barrier Coatings Evaluated by a Steady-State Laser Heat Flux Technique,” Surface 

and Coatings Technology, vol. 138 (2001), pp. 1–8. 

14. Dongming Zhu, Narottam P. Bansal, Kang N. Lee, and Robert A. Miller, “Thermal Conductivity of 

Ceramic Coating Materials Determined by a Laser Heat Flux Technique,” High Temperature Ceramic 

Matrix Composites IV, Proc. the 4th High Temperature Ceramic Matrix Composites Conference (HTCMC 4), Munich, Germany, Oct. 1–3 (2001), Edited by W. Krenkel, R. Naslain, and H. Schneider, 

Wiley-VCH, Verlag GmBH, Germany, 2001, pp. 262–267. Also NASA TM-211122, NASA Glenn 

Research Center, Cleveland, 2001. 

15. Dongming Zhu, and Robert A. Miller, “Thermal Barrier Coatings for Advanced Gas Turbines,” MRS 

Bulletin, vol. 25 (7) 2000, pp. 43–44. 

16. Dongming Zhu, and Robert A. Miller, “Multicomponent Hafnia-Based Oxide Systems Developed, 

Characterized, and Evaluated for Advanced Ceramic-Matrix-Composite Barrier Coating 

Applications,” in Research and Technology 2005, NASA Glenn Research Center, NASA/TM—2006-

214016, pp. 122–124. 

17. Kang N. Lee, Dennis S. Fox, Jeffrey I. Eldridge, Dongming Zhu, Raymond C. Robinson, Narottam P. 

Bansal, and Robert A. Miller, “Upper Temperature Limit of Environmental Barrier Coatings Based 

on Mullite and BSAS,” Journal of the American Ceramic Society, 86 [8] (2003), pp. 1299–1306. 

18. Dongming Zhu, Sung R. Choi, Jeffrey I. Eldridge, Kang N. Lee, and Robert A. Miller, “Surface 

Cracking and Interface Reaction Associated Delamination Failure of Thermal and Environmental 

Barrier Coatings,” NASA TM-212318, 2003. 
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Thermal and Environmental Barrier Coating Development for Advanced Propulsion Engine 

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14. ABSTRACT

Ceramic thermal and environmental barrier coatings (TEBCs) are used in gas turbine engines to protect engine hot-section components in 

the harsh combustion environments, and extend component lifetimes. Advanced TEBCs that have significantly lower thermal conductivity, 

better thermal stability and higher toughness than current coatings will be beneficial for future low emission and high performance 

propulsion engine systems. In this paper, ceramic coating design and testing considerations will be described for turbine engine high 

temperature and high-heat-flux applications. Thermal barrier coatings for metallic turbine airfoils and thermal/environmental barrier 

coatings for SiC/SiC ceramic matrix composite (CMC) components for future supersonic aircraft propulsion engines will be emphasized. 

Further coating capability and durability improvements for the engine hot-section component applications can be expected by utilizing 

advanced modeling and design tools. 

15. SUBJECT TERMS

Thermal barrier coatings; Environmental barrier coatings; Defect clustering; High stability oxides; Thermal conductivity; Sintering; 

Creep; High heat flux; Erosion 

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