ceramic top coat by the same process.

The primary purpose for use of thermal barrier coatings is to continually or momentarily decrease the rate of flow of heat to the underlying substrate. For example, it has been calculated that a ten mil coating of zirconium oxide can reduce the steady state temperature of the leading edge of an advanced design, air cooled turbine blade by 200°F (1). Because of the insulating properties, the coatings can also reduce the magnitude of transient thermal strains on cooled or uncooled parts and therefore delay the onset of life limiting thermal fatigue cracking.

Figure 2. Typical microstructure of electron beam vapor deposited zirconium oxide (1000X).

In gas turbines, the insulating properties of the coatings can be used to either extend the life of hot section airfoils by lowering metal temperature and thermal stresses, or to increase fuel efficiency by decreasing the amount of cooling air used. It has also been anticipated that the ceramic coatings would be more resistant than metallic coatings to hot corrosion caused by low grade fuels, but this has proved to be a somewhat simplistic viewpoint; more will be said on this subject later in this paper.

In diesel engines, the major intended use of thermal barrier coatings has been to Insulate the combustion zone components to reduce heat .loss to the cooling systems and thereby increase fuel efficiency by more effectively using this saved energy by, for example, turbocharging or turbo-compounding. Again It was assumed that the coatings, would also improve component life with respect to hot corrosion caused by low grade fuels. Secondary objectives have included alleviation of thermal fatigue cracking of, for example, piston crowns.

History of Development and Applications

Early Applications of Thermal Sprayed Ceramic Coatings

Although there may be others, the earliest ideas on uses of thermal barrier coatings include some work by the NACA on gas turbine airfoils in the 1945 to 1958 time period. References can be found in a paper by Miller at NASA (2). It must be concluded that while the ideas were sound, coatings available would not withstand the magnitudes of thermal stresses imposed in even these early designs.

There is some anecdotal evidence that USSR technologists were working on some rather detailed designs of thermal barrier coated gas turbine airfoils during the 1950 to 1955 time period but hard evidence for this has eluded the author.

During the decade of the sixties there was a good deal of activity involving the practical use of ceramic coatings to extend the life of rocket nozzles. Ste-cura of NASA has provided a brief summary of related historical events (3). It appears that it was this activity that prompted some of the pioneering work at NASA Lewis on applications of thermal barrier coatings to cooled gas turbine airfoils.

Plasma Sprayed Ceramic Coatings In Gas Turbines

As previously mentioned, the first production application of a thermal barrier coating in a gas turbine engine known to this author involved the use of the coating to extend the life of a hot section sheet metal transition duct. The duct, fabricated from dispersion strengthened nickel, served to guide hot gases from the engine burners to the first stage vanes. It was initially coated only with a diffusion aluminide coating to protect it from excessive oxidative degradation. The available burner pattern factor in the engine was such that the duct was subject to sufficiently severe hot streaking to render its life impractical-ly short, even though it was cooled by means of a louvered design.

In an attempt to alleviate the problem, the long shot of application of a thermal barrier coating was tried. The coating consisted of a plasma sprayed nickel-aluminum bond coat of about two mils, followed by a topcoat of about fifteen mils of magnesia stabilized zirconia. The materials were off—the—shelf from a thermal spray powder vendor and there was at least sufficient technology involved at that time to have determined that it was desirable to attempt to stabilize the cubic crystal form of zirconia to mitigate the effects of the destructive phase transformation between the cubic and monoclinic forms as the practically mandated operating temperatures were varied.

This application was an unqualified success in that it more than doubled the life of the part and Its durability was deemed to be quite satisfactory for the projected application of the engine. While detailed temperature measurements were not made available, it was estimated that the coating lowered the peak metal temperature of the cooled part from about 2100°F to about 1900°F.

It was immediately obvious that thermal barrier coatings could be beneficially used in many other areas of the hot sections of gas turbine engines, provided that economically adequate coating durability could be achieved.

The next successful application of the coatings was on louver cooled sheet metal burner assemblies of the type illustrated in Figure 3. Again, the first, coatings tested were of the two layered nickel-aluminum or nickel-chromium bond coat - magnesia stabilized topcoat variety. Although these were moderately successful, and were used in the then modern high bypass commercial and military engines, occasional spalling of the coatings prompted more extensive laboratory programs aimed at developing longer lived, more reproducible systems.

It was soon discovered that a major cause of coating failure was oxidative degradation of the metallic bond coat. Coincidentally, the so-called MCrAlY turbine airfoil coatings were the subject of intensive development programs at that time and it seemed logical to test them as improved bond coat materials.

It was found that substantial improvements in coating life could be achieved with the use of several CoCrAlY and/or NiCoCrAlY compositions as bond coats (4). There is now little doubt that this was attributable to the superior cyclic oxidation resistance of these coating alloys compared to those based soley on nickel-chromium and nickel-aluminum alloys.

In empirically based attempts to further improve coating life, a considerable amount of work was expended in developing so—called graded coating systems. It was anticipated that grading from the metallic bond coat through mixtures containing less and less metal, to pure zirconia topcoats would increase resistance to spalling caused by the significant thermal expansion mismatch between the metallic substrate and the ceramic coating. Unfortunately, supporting design technology to predict the temperature distribution in the coating for difficult high temperature applications was not available. At considerable expense it was subsequently learned that in many applications the coating temperature was high enough to cause extensive oxidation of the metallic particles in the grading layers. The compressive stresses generated by the volume change from metal to oxide were large enough to cause destructive distortion in nickelbase sheet metal burner structures. Conversely, on more rigid structures, such as airfoils, the same stresses caused spalling of the coatings after very short exposures at practical temperature levels. Graded coating structures were not to reappear until design technology allowed sufficiently accurate temperature prediction to preclude use of the coatings in impractical forms or applications.

Thus, two layer coatings came into wide use on gas turbine burner components and are still the predominant type in use for this application. The best of these coatings consisted of a bond layer of an MCrAlY composition and a ceramic top layer of magnesia stabilized zirconia. Exploratory work on more advanced ceramics, e.g. yttrium oxide (yt-tria) stabilized zirconia was reserved for advanced development.

In about 1973 an experimental test of the capability of then available coatings for protection of gas turbine airfoils was executed. Several types of coatings, two layered, three layered and graded, were applied by plasma spraying to first stage vanes of a large, high-bypass air craft gas turbine. The vanes were subjected to about one-hundred-fifty hours of cyclic endurance testing. A representative airfoil after test is shown in Figure A. Even though the best coating spalled in high heat flux areas, such as the leading edges of the vane airfoils, the test was considered to be moderately successful in that the peak metal temperature in these areas was about 2100°F without coating, and, of course, considerably higher at the surface of the ceramic coating.

In 1976 a landmark eneine test was run

Figure 3. Plasma spray thermal barrier coated aircraft gas turbine inner combustion chamber.

at NASA Lewis with a rotor of first stage cooled blades coated with a two layer thermal barrier system - a NiCrAlY bond coat and a stabilized zirconia topcoat (5). Although the engine was run at relatively low pressures, the gas turbine community was sufficiently impressed to prompt an explosive Increase In development funds and programs to attempt to achieve practical utilization of the coatings on turbine airfoils.

Soon after the public announcement of the NASA test, Pratt and Whitney and NASA agreed to run an endurance test of two layer coatings on JT9D cooled first stage blades (6). The blades were coated at both NASA and Pratt and Whitney and run in a cyclic endurance test in 1977. The results, shown in a representative illus tration in Figure 5, indicated that while the coatings had considerable promise, further development would be required; failure of the coatings had again occurred in high heat flux areas of the blade airfoils. As in all coating failures noted during that period, separation occurred in the ceramic layer just above the bond coat as shown in Figure 6. It was tentatively concluded, as has been now shown by finite element analysis, that the failure was due to excessive compressive, rather than tensile stresses .

Over the next ten years there followed a large number of privately and publicly funded development programs aimed at improving thermal barrier coating properties to the point where they could be practically used in both aircraft and terrestrial gas turbines. Summaries of some of the results of these programs are available (2,6,7).

To provide a better context for describing the results of some of these programs it is useful to digress to a brief description of the state of the art of the plasma spray coating process. In this process, powder particles of the desired composition are injected into a hot, high velocity, partially ionized stream of an inert gas such as argon. The particles are heated to a plastic or molten state and are then allowed to impinge on the article to be coated. The coating structure is thus built up in layered fashion from a myriad of particle "splats" which have been rapidly cooled from the plastic or molten state.

It is immediately apparent that the process, while seemingly quite simple, is very complicated. The net result of this complexity is a coating vith properties which fall within a relatively wide scatter band compared to those of fully dense metallurgically sound systems. However, increasingly tighter control of the major variables of the process, for example, powder particle size and homogeneity, powder feed rate, standoff distance, substrate temperature, and equipment wear, has led to coatings with sufficient reliability to allow applications on moderately critical aircraft engine and diesel engine parts. The development of the capability to plasma spray coatings in a low pressure chamber, or with plasma streams protectively shrouded with argon,' are major advances which have enabled the application of bond coats, of virtually full density and resultant improved oxi dation resistance. Robotically controlled spraying, either in air or at low pressure, has also been of great aid in achieving this level of reliability.

Aided by these improved processes and another major contribution by NASA Lewis - the discovery that a partially stabilized zirconia (approximately 8% yttria) had the highest resistance to thermal stress cracking (3) - development progressed to the point where thermal barrier coatings are routinely applied to hot section vane platforms of commercial aircraft gas turbines as shown in Figure 7 (8,9). In these applications the MCrAlY bond coats are applied either by low pressure plasma spraying or argon shielded torch plasma spraying.

It is also rumored in the industry that at least one advanced model of a large commercial aircraft engine has thermal

Figure 4. Typical condition of plasma spray thermal barrier coated aircraft gas turbine vane airfoils after engine test.

barrier coatings as bill-of-material on first and second hot section air cooled vane airfoils.

In these applications, however, the coatings serve principally to extend the life of the parts; only a relatively modest saving in cooling air, and thereby fuel saving, is achieved. Conditions are such that if the coating fails, the parts will suffer only a shorter life and not immediate failure. Even given the advances that have been made in stress analysis and coating reliability, prime reliant coating-airfoil systems, that is those that are designed to only save cooling air, have apparently not yet been achieved. In such systems there would be some risk of part failure in unacceptably short times if the coating were to fail prematurely.

On a more positive note, capabilities for stress analysis of these coating systems have progressed sufficiently to aid in the development of metal-ceramic multi-layered systems. For example, an apparently workable turbine outer air ■seal comprised of several layers of mixtures of increasing zirconia content and decreasing metal content, ending with an outer layer of pure ceramic, has been described in the patent literature (10).

With regard to use of thermal barrier coatings on airfoils in land based or marine gas turbines, the early promises of Increased efficiency have not yet been fulfilled. Results of publicly supported programs indicate that while the coatings could have satisfactory durability in clean fuel and air environments, exposure to the combustion products of fuels containing relatively large amounts of vanadium and sulfur, would cause early failure of the coatings (11). Although the subject is still somewhat controversial, laboratory studies on the possible mechanisms of hot corrosion of various stabilized zirconia ceramics (12) tend to support the contention that systems with better resistance to such corrosion will be necessary to allow practical applications in turbines burning low grade fuels.

Electron Beam Vapor Deposited Coatings In Gas Turbines

As previously described, stabilized zirconia thermal barrier coatings can also be deposited by physical vapor deposition methods, the most well developed technique being that of electron beam evaporation. Early thermal stress tests on yttria stabilized zirconia, vapor deposited with a columnar grain structure over a vapor deposited NiCoCrAlY bond coat, indicated an order of magnitude life improvement over plasma sprayed coatings (13,14).

Processing methods for these coatings have been refined to the point where such coated blades have been reliably tested In test stand (15) and field service test engines (16). If prime reliant coating systems are to become practical, it would appear that electron beam vapor deposited zirconia is the leading candidate.

Plasma Sprayed Thermal Barrier Coatings In Diesel Engines

The first widely publicized application and testing of thermal barrier coatings on diesel engine components was the work of Kvernes (17). The initial results showed promise of increasing the hot cor-

Figure 5. Typical condition of plasma spray thermal barrier coated aircraft gas turbine 1st stage blades after 1977 engine test.

coatings on combustion area components (valves, pistons, fire decks) of medium speed diesels. The early test results were sufficiently promising to cause coating of the same components for controlled testing in a tugboat engine application. The coating technology was drawn directly from aircraft engine experience; that is 2 mils of a CoNiCrAlY bond coat and 15 mils of partially stabilized (8% yttria, pre-alloyed) zirconia, both deposited by robotized air plasma spraying. The spray process and typical parts are illustrated in Figures 8, 9 and 10.

Figure 6. Typical failure mode of plasma sprayed zirconia thermal barrier coating (500X).

rosion resistance -of large pistons in low speed diesels; apparently thermal barrier effects in terms of Increased efficiency were not determined or noted in this irork. Kvernes' work did spark world-wide interest in the potential of the coatings as a stepping stone between all-metallic to all-ceramic components on the way to development of so-called adiabatic diesel engines.

In the U.S., the Maritime Administration and the Association of American Railroads sponsored programs involving testing of plasma sprayed thermal barrier

As has been partially reported (18) and will be more fully reported at this conference (19), the coatings endured 14,000 hours of running without failure. While the coatings were not sufficiently thick to cause significant thermal barrier effects, such as an increase in exhaust gas temperature, other benefits, including reduced fuel consumption, reduced ignition delay, increased maximum power and smoother overall operation, were documented.

The above results have prompted funding of additional development programs by

DOE-NASA aimed at providing durable thicker coatings to attempt to achieve true thermal barrier effects and the potential further increases in fuel efficiency (20,21,22).

Suggestions For Future Development

While this keynote paper has not attempted to present a comprehensive review and bibliography of all past and present research and development programs on thermal barrier coatings for heat engines, it to the routine measurement of thermal conductivity of production coatings should be accelerated. Such methods should allow determination of bond quality and porosity which are the two most important properties of thermal barrier coatings not currently under adequate production control.

Upgrading of powder technology should be continued to achieve a better tradeoff between uniform coating properties and powder costs.

The scope of research on hot corrosion of ceramics should be broadened to pro-

Figure 8. Robot controlled plasma spraying of thermal barrier coatings on diesel engine piston crowns.

Figure 9. Plasma sprayed thermal barrier coatings on diesel engine piston crowns.

Figure 8. Robot controlled plasma spraying of thermal barrier coatings on diesel engine piston crowns.

does provide a context for the more detailed reports which follow and perhaps also for some suggestions for work which may not be adequately funded at this time. A few such suggestions follow.

Research should be continued to more completely characterize the two critical interfaces in thermal barrier coatings, that is, the bond coat-thermally grown alumina and the alumina-zirconia interfaces. Coating integrity is completely dependent on these interfaces and understanding of MCrAlX-alumlna adherence, for example, is still far from complete.

Development of non-destructive inspection methods, particularly with regard

Figure 10. Plasma sprayed thermal barrier coatings on diesel engine fire deck and valves.

vide better support for the technology required for increased use of low grade fuels.

Research should be initiated to understand the improvement in diesel engine efficiency observed when the ceramic coatings in combustion zones are so thin that true thermal barrier insulation effects are negligible.


1. Sevcik, W.R. and Stoner, B.L., An Analytical Study of Thermal Barrier Coated First Stage Blades in a JT9D Engine; NASA CR-135360, 1978.

2. Miller, R.A., Current Status of Thermal Barrier Coatings - an Overview; Surface and Coatings Technology 30 1-11 (1987).

3. Stecura, S., Effects of Compositional Changes on the Performance of a Thermal Barrier Coating System; NASA TM-78976, 1979.

4. Goward, G.W., Grey, D.A., and Krutenat, R.C., Thermal Barrier Coating for Nickel and Cobalt Base Superalloys; U.S. Patent 4,248,940, 1981.

5. Liebert, C.H., Jacobs, R.E., Stecura, S., and Morse, C.R., Durability of Zirconia Thermal Barrier Ceramic Coatings on Air Cooled Turbine Blades in Cyclic Jet Engine Operation; NASA TM-X-3410, 1976.

6. Grlsaffe, S.J. and Levine, S.R., Review of NASA Thermal Barrier Coating Programs for Aircraft Engines; Proceedings of the First Conference on Advanced Materials for Alternative Fuel Capable Directly Fired Heat Engines, Eds. Fairbanks, J.W. and Stringer, J., DOE-EPRI, 1979, pp 680-703.

7. Goward, G.W., Thermal Barrier Coatings for Gas Turbines - A Review of Development and Production Capability; Proceedings of the First NATO Advanced Workshop on Coatings for Heat Engines, Eds. Clarke, R.L. et al., Acquafredda di Maratea, Italy, 1984, pp 195-217.

8. Gaffin, W.O., NASA ECI Programs: Benefits to Pratt and Whitney Engines; Paper 82-GT-272, American Society of Mechanical Engineers, 1982.

9. Bennett, A., Toriz F.C., and Thakker, A.B., A Philosophy for Thermal Barrier Coating Design and Its Corroboration by 10,000 Hour Service Experience on RB211 Nozzle Guide Vanes, Surface and Coatings Technology, 32 359-375 1987.

10. Bosshart, G.S. and Matarese, A.P., Method of Applying Ceramic Coatings on a Metallic Substrate; U.S. Patent 4,481,237, 1984.

11. Bratton, R.J., Lau, S.K., and Lee, S.Y., Evaluation of ;Present Day Thermal Barrier Coatings for Potential Service in Electric Utility Gas Turbines; NASA CR-165545, July, 1982.

12. Jones, R.L., Low Quality Fuel Problems with Advanced Engine Materials, This Conference.

13. Strangman, T.E., Columnar Grained Ceramic Thermal Barrier Coatings; U.S. Patent 4,321,331, 1982.

14. Ulion, N.E. and Ruckle, D.L., Columnar Grained Thermal Barrier Coatings on Polished Substrates: U.S. Patent 4,321,310, 1982.

15. Strangman, T.E., Development and Performance of Physical Vapor Deposition Thermal Barrier Coating Systems; This Conference.

16. Shankar, R., Electron-Beam Physical Vapor Deposition Development of Zirconia Coatings; This Conference.

17. Kvernes, I., Farturn, P., and Henrik-sen, R., Characterization of Microstructures and Measurements of Thermal Effects of Ceramic Coatings in Service; Metallurgical Coatings 1980, American Vacuum Society, Elsevier Sequoia, Lausanne and New York, Volume 2, p478.

18. Levy, A. and MacAdam, S., The Behavior of Ceramic Thermal Barrier Coatings on Diesel Engine Combustion Zone Components; Surface and Coatings Technology 30 51-61 (1987).

19. Winkler, M., Durability Testing of Ceramic Coatings Aboard MV Bill Gee; This Conference.

20. Holtman, R.L., Layne, J.L., and Schechter, B., An Investigation of Enhanced Capability Thermal Barrier Coating Systems for Diesel Engine Components; DOE/NASA/0326-1, August, 1984.

21. Yonushonis, T., Thick Thermal Barrier Coatings Development for Diesel Engines; This Conference.

22. Biehler, D., Thick Thermal Barrier Coatings for Diesel Engines; This Conferenfce.

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