Thermal barrier coatings are used to extend the life of metal components by creating a temperature drop across the coating, permitting the underlying metal to operate at a reduced temperature.
If you’ve flown on a commercial jet aircraft recently, it’s virtually certain that parts of its engine were protected by zirconia thermal barrier coatings (TBCs). These coatings are used to extend the life of metal components by creating a temperature drop across the coating, permitting the underlying metal to operate at a reduced temperature. Future gas turbines (GT) will use TBC technology to permit the simultaneous increase of turbine inlet temperature and the reduction of turbine cooling air, thereby increasing efficiency.
How TBCs Work
Superalloys used in GTs melt at temperatures between ~1200 and 1315°C. The combustion gases that flow through these engines are ~1350°C or higher. How do the engines run without melting? Large amounts of compressor air are used to cool the engine components, thereby avoiding melting, thermal fatigue and a variety of other potential failure modes. Providing this cooling air comes at the cost of decreasing engine performance and fuel economy. If less cooling air is required, fuel economy or other measures of performance can be increased. If cooling air and the temperature of the metal parts are simultaneously reduced, fuel economy and engine component lifetimes can be increased. This is what zirconia based TBCs do.
The properties of zirconia most critical for TBCs are a very low thermal conductivity (~1 W/mK) and a thermal expansion close to that of superalloys. If a thin layer of zirconia is coated on a cooled metal substrate, a significant DT can be supported across the layer. (If the substrate is not cooled, the DT will approach zero.) The zirconia coatings used in current engines can sustain a DT of ~165°C in airfoils, reduce specific fuel consumption by ~1%, increase thrust-to-weight ratios by ~5%, and significantly extend component life.
TBCs are a two-layer system composed of a zirconia layer ~0.254 mm thick, which faces the hot combustion gases, and an ~0.127 mm bond coat (typically, NiCoCrAlY alloy). The bond coat provides strong coating adherence and enhances the oxidation resistance of the substrate metals.
About 30 years ago TBCs of fully stabilized 22 wt % MgO/ZrO2
were introduced in the combustors of commercial aircraft. The coatings were applied by the plasma spray technique. In this process an electric arc ionizes an argon gas to form a plasma. Ceramic powders are injected into the plasma, heated to a “semiplastic” state, and accelerated toward the combustor surface. When the particles impact the target, a very complex interlocking microstructure, which is highly porous and microcracked, results. These coatings worked very well and extended combustor life as long as the temperature did not exceed ~980°C. Above this temperature, the MgO doped ZrO2
destabilized and failed by spallation.
A second-generation material, 7 wt % yttria, partially stabilized zirconia (7YPSZ) is now in use, which provides a fourfold increase in coating life at temperatures of ~1090°C. A process for making denser TBCs with highly columnar structures has facilitated the insertion of zirconia based coatings onto blades and vanes. This process, electron beam-physical vapor deposition (EB-PVD), relies on vaporizing the 7YPSZ with an electron beam and positioning the part so that the vapor will deposit where desired. Using this process, highly columnar grains of 7YPSZ have been deposited on airfoils. The columnar structure provides interfaces that are weakly bonded and thus separate at low stresses. This provides a coating with a high strain and thermal cycling tolerance. Airfoils with such coatings have been in airline service since the late 1980s.
EB-PVD TBCs have a thermal conductivity of ~1.8 W/mK, compared to ~1 W/mK for plasma sprayed coatings. One way of achieving a lower value of conductivity in the EB-PVD zirconia is to process it in a way that yields a multilayer substructure within each columnar grain. Such work is being actively pursued. Alternative materials under study include lanthanum phosphate and lanthanum hexaaluminate. The latter material may be useful in the 1100-1600°C range.
For Further Reading
1. T. Abraham, “A Cutting Edge in High-Performance Ceramic Coatings,” Ceramic Bulletin,
March 1999 pp. 69-71.
2. S M Meier, D K Gupta, and K D. Sheffler, “Ceramic Thermal Barrier Coatings for Commercial Gas Turbine Engines,” J. of Metals,
March 1991, pp. 50-53.
3. S M Meier and D K Gupta, “The Evolution of Thermal Barrier Coatings in Gas Turbine Engine Applications,” Trans. of the ASME,
vol. 116, January 1994, pp. 250-256.
4. J Kumpfert, M Peters, and W A Kayser, “Advanced Materials and Coatings for Future Gas Turbine Technology,” in the Proceedings of the NATO-RTO Symposium on, Gas Turbine Operation and Technology for Land, Sea and Air Propulsion and Power Systems,”
Ottawa, Canada, October 18-21, 1999.