High Temperature Coatings
Thermal Barrier Coatings Systems
Thermal barrier coatings (TBCs) are used extensively in both commercial and military gas turbine engines to increase component life and engine performance, Figure 1. TBCs are really a system of coatings consisting of a bond coat, a thermally grown oxide (TGO), and a thermally insulating ceramic top coat. In most applications, the bond coat is either a MCrAlY (where M=Ni or NiCo) or a Pt modified aluminide coating. A dense bond coat is required to provide protection of the superalloy substrate from oxidation and hot corrosion attack and to form an adherent, slow growing TGO on its surface. The TGO is formed by oxidation of the aluminum that is contained in the bond coat. The current (first generation) thermal insulatoin layer is composed of 7wt % yttria stabilized zirconia (7YSZ) with a typical thickness of 100-300 µm. Yttria stabilized zirconia is used due to its low thermal conductivity (2.6W/mK for fully dense material), relatively high coefficient of thermal expansion, and good high temperature stability. The electron beam directed vapor deposition (EB-DVD) process used to apply the TBC to turbine airfoils produces a columnar microstructure with several levels of porosity, Figure 2. The porosity between the columns is critical to providing strain tolerance (via a very low in-plane modulus), as it would otherwise spall on thermal cycling due to thermal expansion mismatch with the superalloy substrate. The porosity within the columns reduces the thermal conductivity of the coating.
Figure 1: A schematic illustration of a thermal barrier coating system applied to a hollow air cooled turbine blade
Figure 2: The outer surface and cross sections of 7YSZ coatings as depsoted (left) and after annealing at 1100oC for 50 hours.
The evolution of the allowable gas temperature in gas turbine engine turbines is shown in Figure 3. The use of multilayer TBC systems in future gas turbine engines that operate at higher temperatures will require improved TBC durability and an increasing resistance to high temperature and long time exposures in corrosive environments. New materials having improved high temperature properties must therefore be developed. Current top coat compositions are limited by a lack of phase and thermal stability at elevated temperatures and localized coating spallations caused by erosion/impact or CMAS damage.
Figure 3: The evolution of allowable gas temperature at the entry to the gas turbine and the contribution of superalloy development, film cooling technology, thermal barrier coatings and (in the future) ceramic matrix composite (CMC) air foils and perhaps novel cooling concepts.
Our groups’ research uses our novel directed vapor deposition processing approach for depositing compositional and morphologically controlled top coat and bond coat layers intended for use in next generation TBC systems.
Our recent work in this area consists of studying the properties of rare earth based zirconate materials (La, Gd, Sm and Yb etc.), which are promising candidate materials due to their high melting point, low thermal conductivity, high temperature phase stability and good sintering resistance. Specifically we are interested in the effect of zirconate compositions (for three, four and five component systems) on the phase stability, thermal conductivity and thermalchemical stability with alumina. Characterization techniques include HRTEN, SEM, XRD and direct thermal conductivity measurement based on the use of an IR camera.
A La2Zr2O7 coating is shown in Figure 4. The coating was reactively deposited from two different metal source rods (La and Zr) at 1050° using a helium-5.0 vol.% O2 gas jet
Figure 4: The surface a fractured cross section of a lanthanum zirconate (La2Zr2O7) thermal barrier coating deposited by EB-DVD.
A related samarium zirconate coating grown by evaporating a Sm2Zr2O7 source is shown in Figure 5. The coatings deposited in these EB-DVD processes can be manipulated by the jet flow conditions. Figure 6 shows the effects of this porosity upon conventional yttria stablized zirconia and samanum zirconate coatings. By combing a low intrinsic thermal conductivity with a high pore fraction we have been able to make coatings of exceptionally low thermal conductivity.
Figure 5: A fracture cross section of a samarium zirconate (Sm2Zr2O7) coating.
Figure 6: Effect of pore volume fraction upon the through thickness thermal conductivity of yttria stabilized zirconia and samarium zirconate coatings. The EB-PVD coatings can be grown with very high pore volume fractions and exceptionally low thermal conductivity.
As the gas temperatures in the high pressure turbine section of engines increases, the dust particles present in air (especially in desert regions) melts upon impact with TBC's. This liquid oxide containing varying quantities of calcium, aluminum, magnesium and silica (CMAS) then wicks into the porous coating causing it to fail on cooling. The sequence of events is summarized in Figure 7. Examples of some of the reactions products that are precipitated from a cooling CMAS melt that has dissolved a TBC are shown in Figure 8. There is great interest in finding solutions to this very serious problem.
Figure 7: Molten CMAS penetration through a surface heated coating containing inter-column pores. Penetration stops when CMAS reaches regions that are below its melting point. On cooling the loss of the strain compliance in the TBC layer and high CTE of the glass induces spallation of the infiltrated region.
Figure 8: Examples of the reactions that occur between CMAS and TBC’s.
Our group has explored one novel concept in which we deposit a thin layer of a nonreactive metal (such as platinum) near the coating surface, Figure 9. Preliminary research indicates the metal layer is able to arrest the CMAS infiltration confining it near the surface where it is much less damaging.
Figure 9: A cross-sectional image of a 7YSZ/Sr2Zr2O7 bi-layer sample containing a Pt layer in the Sm2Zr2O7 part of the coating after exposure to CMAS attack at 1250°C for 16hrs.
Environmental Barrier Coatings
Improvements to the fuel efficiency and specific thrust of future gas turbine engines will (i) require increases to the gas turbine inlet temperature, (ii) increase of the overall pressure ratio of the engine and (iii) incorporation of constant volume combustion concepts into novel engine designs. The combined effect will enable continued increases in both engine thermal and propulsive efficiencies. All three of these objectives are paced by the emergence of new generations of robust propulsion materials, protective surface coatings and their integration with novel material cooling (thermal management) concepts together with the development of manufacturing processes (and supply chains) that ensure reliable (and economical) component fabrication. Arguably the most difficult materials challenge will be encountered in the high pressure turbine section located immediately behind the combustors. Here, severely stressed rotating components will be subjected to impinging gas streams whose temperature will eventually approach the stoichiometric hydrocarbon-compressed air combustion temperature (~1800°C).
Increases of the turbine inlet temperature of advanced military engines, such as the F-135 engine, are achieved by the use of internally air cooled superalloy components protected against oxidation and hot corrosion by thermal barrier coatings (TBC’s). These coatings, when combined with internal air cooling, enable the surface temperature of the metallic component to be maintained below its maximum use temperature even as the gas temperatures significantly exceed it. This strategy is reaching its limit because the coatings are susceptible to sintering (leading to a loss of some thermal resistance and increased risk of delamination) and attack by molten droplets of calcium magnesium aluminum silicate (CMAS). Additionally, at the highest temperatures these coatings only weakly impede the radiative transport of the thermal flux to the metal surface. Even if these life limiting issues were solved, TBCs can be eroded by the impact of fine (dust) particles or chipped from the substrate by larger foreign object (particle) damage (FOD). The combination of these intrinsic and extrinsic damage processes results in a necessity for today’s TBC systems to be replaced three to four times over the design life of a turbine blade. Further significant increases in turbine inlet temperatures will require more frequent and costly coating replacement. This will either effectively end the use of metallic hardware in the highest temperature regions of future engines or stall future increases in engine operating temperature.
The maturation of thermal protection concepts for metallic components in the most advanced gas turbines has stimulated efforts to develop turbine components from ceramic materials with much higher maximum use temperatures. Monolithic ceramics such as those based on silicon nitride have insufficient resistance to flaw growth under engine thermal and mechanical stress conditions, and so the focus has concentrated upon more damage tolerant fiber reinforced ceramic matrix composites (CMCs) with weak fiber/matrix interfaces. Figure 1 shows the past evolution of turbine inlet gas flow and material temperatures for the most advanced engines and anticipates the needs (and our view of the most promising candidate materials and thermal management strategies) of the future. The most promising CMC’s are based either upon (i) woven fabrics of either aluminum oxide (Nextel 720) fibers and tape cast alumina matrices or (ii) boron nitride coated SiC fibers (such as Hi-Nicalon S and Sylramic fibers) and SiC matrices incorporated by chemical vapor infiltration with residual pores filled by silicon slurry infiltration (followed by carburization). While the oxide CMC systems are chemically inert in oxidizing environments, the current fibers they have low creep rupture strengths at engine operating temperatures and so future engines are likely to utilize SiC CMC systems for the most highly loaded components.
Figure 1: The past and (predicted) future evolution of propulsion materials, coatings, cooling concepts and turbine inlet T4.1 gas temperatures with year of entry.
Unfortunately, silicon containing ceramics react with oxygen and water vapor in combustion environments to form SiO2 scales, and these then react with water vapor to form one of the gaseous silicon hydroxides by one or a combination of the following reactions:
SiO2(s) + 2H2O(g) = Si(OH)4(g)
SiO2(s) + H2O(g)= SiO(OH)2(g)
SiO2(s) + 1/2H2O(g) = SiO(OH)(g) + 1/4O2(g)
The rate of SiC volatilization depends upon the temperature, the incident water vapor flux (engine pressure) and the effectiveness with which the silicon hydroxide reactions occur. SiC recession rates significantly greater than 1μ/hr can occur at engine gas temperatures in the 1300-1350°C temperature range where pressures of 5-20 atmospheres (or higher) exist (depending upon the engine’s altitude of operation and overall pressure ratio). Since engine components are normally expected to survive for thousands of hours of operation, these composite components must be coated by materials that impede the diffusion of oxygen and water vapor to the composite surface, thereby inhibiting reactions. The development of these environmental barrier coatings (EBCs) is likely to pace the future use of silicon-based CMC’s in the hot sections of military engines until the discovery and development of other (less vulnerable) ceramic systems (such as higher creep strength fibers perhaps based upon oxide-oxide eutectic composites) shown on Figure 1.
A typical advanced EBC system is schematically illustrated in Figure 2. It consists of a silicon bond coat applied directly to a SiC based CMC, a layer of mullite (which impedes diffusion of oxygen to the silicon bond coat) and a low volatilization rate material (in this case a rare earth silicate) that protects the mullite and silicon layers from volatilization by water vapor. To achieve robust performance the EBC must be designed and fabricated from materials that provide complete aerial protection from oxygen and water vapor penetration for thousands of hours of operation at temperatures approaching (and eventually exceeding) 1500°C. The EBC must also not fail by coating fracture or delamination during repeated thermal cycling; it has to be able to survive impact by small and large particles (exhibit erosion and FOD resistance) and it must be able to survive exposure to molten CMAS and salts that are present in fuel and the naval operating environment.
Figure 2:The architecture of a candidate environmental barrier coatings system for protecting silicon carbide composite components.
Multilayer EBC systems are usually deposited by thermal spray processes and an example of a coating made this way using a thermal spray facility at NASA Glenn is shown in Figure 3. The as deposited coatings suffer from many problems including “mud-cracking”, void incorporation, internal oxidation and evaporation of some elements in the rare earth silicate coatings. The back scattered electron image of the top coat in Figure 3 reveals many of these issues. The variation in contrast arises from compositional differences in the powder particles that have been deposited. The small particles appear to have lost a significant fraction of their silicon content. Upon annealing multiple phases form in the silicate layer, Figure 4.
Figure 3: Micrographs of an ytterbium monosilicate topcoat trilayer EBC in the as-fabricated state.
Figure 4: Exampled of a plasma sprayed ytterbium monosilicate coating after annealing at 1200°C
Our group has developed an improved plasma spray process and installed a full scale system in our high temperature coatings laboratory. A schematic illustration is shown in Figure 5. Up to four different powders can be melted in this process under very controlled conditions and then sprayed on to components. In future years we will begin a systematic exploration of the relation between processing conditions and defects in the coatings and explore the role of these defects during high temperature exposure in steam environments.
Figure 5: A schematic diagram of the air plasma spray system developed by our group for EBC coating deposition.
Morphology and Thermal Conducitivity of Yittria Stabilized Zirconia Coatings, Hengbei Zhao, Fengling Yu, Ted D. Bennett, Haydn N.G. Wadley, Acta Materialia, 54, p. 5195-5207, 2006.
Gas Jet Assisted Vapor Deposition of Yttria Stabilized Zirconia, D.D. Hass, H.N.G. Wadley, Journal of Vacuum Science and Technology A, 27, p. 404-414, 2009.
Vapor Deposited Samarium Zirconate Thermal Barrier Coatings, Hengbei Zhao, Carlos G. Levi, Haydn N.G. Wadley, Surface and Coating Technology, 203, p. 3157-3167, 2009.
Pore Evolution During High Pressure Atomic Vapor Deposition, D.D. Hass, Y.Y Yang, H.N.G. Wadley, Journal of Porous Materials, 17, p. 27-38, 2010.
The Influence of Coating Compliance on the Delamination of Thermal Barrier Coatings, Hengbei Zhao, Zhuo Yu, Haydn N.G. Wadley, Surface and Coatings Technology, 204, p. 2432-2441, 2010.
Multi-scale Pore Morphology in Vapor Deposited Yttria-Stabilized Zirconia Coatings, D.D. Hass, H. Zhao, T. Dobbins, A.J. Allen, A.J. Slifka, H.N.G. Wadley, Materials Science and Engineering A, 527, p. 6270-6282, 2010.
The Vapor Deposition and Oxidation of Pt/YSZ Multilayers, Zhuo Yu, Hengbei Zhao, Haydn N.G. Wadley, Journal of American Ceramic Society, 94, p. 2671-2679, 2011.
Delamination of Ceramic Coatings with Embedded Metal Layers, Matthew R. Begley, Haydn N.G. Wadley, Journal of American Ceramic Society, 94, p. 96-103, 2011.
Reaction, Transformation and Delamination of Samarium Zirconate Thermal Barrier Coatings, Hengbei Zhao, Matthew R. Begley, Arthur Heuer, Reza Sharghi-Moshtaghin, Haydn Wadley, Surface and Coating Technology, 205, p. 4355-4365, 2011.
Delamination Resistance of Thremal Barrier Coatings Containing Embedded Ductile Layers, Matthew R. Begley, Haydn N.G. Wadley, Acta Materialia, 60, p. 2497-2508, 2012.