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126) Letokhov. V. S.:Moushev. V. G.; Shekalin. S. V. Sov. Phvs. JETP 1981, 54, 257. Heidberg, J.; Stein, H.; Riehl, E. Phys. Rev. Lett. 1982, 4 9 , 666. Heldberg, J.; Stein, H.; Riehl, E. Surf. Sci. 1983, 126, 183. Mashni, M.; Hess, P. Chem. Phys. Lett. 1981, 77,541. Mashni, M.; Hess, P. Appl. Phys. 8 1982, 28,224. Chuang. T. J. J . Chem. Phys. 1982, 76,3829. Chuang, T. J. J . Vac. Sci. Techno/. 1982, 20,603. Chuang, T. J.; Seki, H. Phys. Rev. Lett. 1982, 49,382. Seki, H.. Chuang, T. J. Solid State Commun. 1982, 4 4 , 473. Cowin, J. P.; Auerbach, D. J.; Becker, C.; Wharton, L. Surf. Sci. 1978, 78,545. Wedler, G.; Ruhmann, H. Surf. Sci. 1982, 121,464. Slutsky, M. S.;George, T. F. Chem. Phys. Lett. 1978, 57,474. Slutsky, M. S.;George, T. F. J . Chem. Phys. 1978, 70,1231. Lin, J.; George, T. F. J . Chem. Phys. 1980, 72,2554. Jedrzejek, C.: Freed, K. F.; Efrima, S.;Metiu, H. Surf. Sci. 1981, 109. I91
(41) Lica, D.; Ewing, G. E. Chem. Phys. 1981, 58,385. (42) Gortel, 2. W.; Kreuzer, H. J.; Piercy, P.; Teshima, R. Phys. Rev. 8 1983, 27. 5066. (43) Yates, J. T.; Zlnck, J. J.; Sheard, S.;Weinberg, W. H. J. Chem. Phys. 1979, 70,2266. (44) Umstead, M. E.; Talley, L. D.; Tevault, D. W.; Lin, M. C. Opt. Eng. 1980. 19.94. (45) Panchenkov, G. M.; Lebedev, V. P. "Chemical Kinetics and Catalysis", MIR Publlshers: Moscow. 1976: Chaoter IO. (46) Lin, J.; George, T. F. C h e h . Phys. Lett. 1979, 66,5. (47) Chandrasekhar, S. Rev. Mod. Phys. 1943, 15, 1. (48) Collins, F. C.; Kimball, G. E. J . Colloid Sci. 1949, 4 , 425. (49) Stowell, M. J. Philos. Mag. 1972, 26. 349. (50) Freeman, D. L.; Doll, J. D. J . Chem. Phys. 1983, 78,6002. (51) Freeman, D. L.; Doll, J. D. J . Chem. Phys. 1983, 79,2343. (52) Doll, J. D.; Freeman, D. L. Surf. Sci. 1983, 134,769. (53) Lln, J.: George, T. F. J . Chem. Phys. 1983, 78,5197. (54) Goldanskii, V. 1.: Namoit, V. A.; Khokholov. R. V. Sov. Phys. JETP 1978. 43, 1226. (55) Karlov, N. V.; Lukyanchuk, B. S . Sov. J . Quantum Electron. 1981, 1 1 , 909. (56) Metiu, H.; Gadruk, J. W. J . Chem. Phys. 1981, 74,2641. (57) For recent studies see: Osgood,R. M., Jr.; Brueck, S. R. J., Ed "Laser Diagnostics and Photochemical Processing for Semlconductor Devices"; Elsevier: New York, 1983. [This is Volume 17 of the Materials Research Society Symposium Proceedings.] (58) Ehrlich, D. J.; Tsao, J. Y. J . Vac. Sci. Tech. 8 1983, 1 , 969. (59) Ehrlich, D. J.; Osgood, R. M.. Jr.; Deutsch, T. F. I€€€ J . Quantum Eiectron. 1980, QE-16.1233.
(60) Ehrlich, D. J.; Osgood, R. M., Jr. Appl. Phys. Lett. 1980, 3 6 , 698. (61) Wood, T. H.; White, J. C.; Thacker, E. A. in ref 57,p 35. (62) Ehrllch, D. J.; Tsao, J. Y. I n ref 57, p 3. (63) McWllliams, B. M.; Herman, I. P., Mltlltsky, F.; Hyde. R . A,; Wood, L. L. Appi. Phys. Lett. 1983, 43,946. (64) Herman, I.P.; Hyde, R. A.; McWilllams, B. M.; Weisberg, A. H.; Wood, L. L. I n ref 57,p 9. (65) Tsao, J. Y.; Ehrlich, D. J. I n ref 57,p 235. (66) Ehrllch, D. J.; Tsao, J. Y. Appl. Phys. Lett. 1982, 4 1 , 297. (67) Lin, J.; George, T. F. J . Appl. Phys. 1983, 5 4 , 382. (68) Lundqvist, S. I n "Surface Science", International Atomic Energy Agency: Viennia, 1975. (69) Murphy. W. C.; George, T. F. Surf. Sci. 1982, 114, 189. (70) Murphy, W. C.; Lee, K.-T.; George, T. F. Surf. Sci. 1983, 127,L156. (71) Murphy, W. C.; Berl, A. C.; George, T. F.; Lin, J. I n ref 43, p 273. (72) Murphy, W. C.; George, T. F. J . Phys. Chem. 1982, 86,4481. (73) George, T. F.; Lam, K.-S.; Hutchinson, M.; Murphy, W. C. I n "Advances in Laser Spectroscopy", Vol. 2,Garetz, B. A,; Lombardi, J. R., Ed.; Wiley: New York, 1983. (74) Murphy, W. C.; George, T. F. J. Chem. Phys. 1984, in press. (75) van Driel, H. M.; Sipe, J. E.; Young, J. F. Phys. Rev. Lett. 1982, 4 9 , 1955,and references therein. (76) Sipe, J. E.; Young, J. F.; Preston, J. S . ; van Driel, H. M. Phys. Rev. 8 1983, 27, 1141. (77) Brueck, S. R . J.; Ehrlich, D. J. Phys. Rev. Lett. 1982, 48, 1678. (78) Lord Rayleigh Philos. Mag. 1907, 14,60. (79) Sheng, P.; Stepheman, R. S.;Sanda, P. N. Phys. Rev. 8 1982, 26, 2907. (80) Hutchinson, M.; Lee, K.-T.; Murphy, W. C.; Beri, A. C.; George, T. F. I n "Laser-Controlled Chemical Processing of Surfaces", Johnson, A. W.; Ehrlich, D. J., Ed.; Elsevier: New York, in press. [This is Volume 29 of the Materials Research Society Symposium Proceedings.] (81) Murphy, W. C.; Huang, X.-Y.: George, T. F. Chem. Phys. Lett. 1984,
104,303. (82) Huang. X.-Y., Unpublished results at The University of Rochester, Rochester, NY, 1984. (83) Auston, D. H.;Brown, W. L.; Celler, G. C. Bell Laboratories Record July/Aug 1979, 187. (84) Lyman, J. Electronics 1977, 50(15), 81. (85) Mason, E. Electronics 1979, 52(5),90. (86) Kaplan, R. A.; Cohen, M. G.; Klu, K. C. Electronics 1980, 53(4),137. (87) Brlnton, J. B. €lectfonics 1981, 54(8),39. (88) Waller, L. Necfronics 1980, 53(21),46.
Received for review May 1, 1984 Accepted June 7, 1984
Ceramic Thermal Barrier Coatings Curt
H. Liebert' and Robert A.
Mlller
Nafionai Aeronautics and Space Administration, Lewis Research Center, Cleveland, Ohio 44 135
Ceramic thermal barrier coating research is reviewed in periods from the early 1950's through the early 1970's, from the seventies to the eighties, and for the past few years. Recent major developments include use of ceramic coatings on various components of gas turbine engines such as first-stage turbine vane platforms of advanced production engines, uncooled integrally bladed turbine wheel rims, and cooled combustion transition sections. These developments have been realized through improvements in coating materials and processing conditions and through heat transfer and aerodynamic analysis. Future improvements will be facilitated when coating failure mechanisms are more fully understood. Current understanding of failure mechanisms is reviewed.
Introduction Zirconia-based thermal barrier coatings have attracted much attention because of their ability to provide thermal insulation (Liebert and Stepka, 1977) for gas turbine components exposed to very hot combustion gases. Much of the ongoing research and development of these coatings
is directed toward improved durability of metal components in high-performance military and commercial aircraft gas turbine engines and in diesel engines. Thermal barrier coatings (TBC) have been used in aircraft gas turbine engines since the late 1960's to extend the life of combustor liners. Presently TBC's are used to a limited extent in the turbine section of some production engines to increase component life and they are just starting to be used to reduce cooling air requirements. Current research efforts are primarily directed toward developing coatings which can be used with extended reliability in the combustor and turbine sections of advanced gas turbine engines. Therefore, considerable attention is being directed to understanding and modeling the durability behavior of TBC's in the high temperature and pressure environment of these engines. Benefits to be realized from the use of TBC's include increased engine efficiency and power through higher combustion temperatures or decreased cooling air requirements. Component durability may be improved through decreased base metal
This article not subject to U.S. Copyright. Published 1984
by the American Chemical Society
Ind. EM. Chem. Rod. RW. Dw., Vol. 23. No. 3, I984 S4S
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I
1 's\ r,
Figure 2. Arc plasma spraying of turbine blade.
1 Figure 1. Ceramic coated turbine blade.
temperatures or moderation of temperatranaientn, and component costa may be reduced through elimination of elaborate cooling schemes. T h e TBC Concept A duplex-layer thermal bnrrier coating on a supatalloy turbine blade is shown in Figwe 1. The outer layer of the coating consists of a zirconia-based ceramic. The choice of ceramic thickness depends on gas turbine component design requirements. The thickness typically is in the range of 0.01 to 0.06em. Adherence of the zirconia ceramic to the metallic turbine blade is a critical requirement for a satisfactory TBC. Adhesion of zirconia directly to the superalloy substrate is relatively poor but it can be considerably improved ifa metallic bond coat is used between the ceramic and superalloy substrate (Liebert and Stepka, 1977). The bond coata are usually taken from the MCrAlY alloys, e.g. NiCrAlY, and are typically 0.01 cm thick. Both ceramic and bond coat layers can easily be applied individually by the plasma spray process as shown in Figure 2. Metal or ceramic powder is fed into an extremely high-temperature, gas-stabilized arc plasma. The powder particles melt as they are rapidly propelled to the substrate upon which they splat to build up a coating to any desired thickness. Before the bond coating is applied, the superalloy substrate is typically roughened by grit blasting with aluminum oxide to increase the adherence of the bond mat to the superalloy substrate. Details of typical plasma spray processing are given by Liebert and Stepka (1977). Early TBC Research and Development: 1950 to 1972 Most of the early TBC research was conductad at the NASA-Lewis F&searcb Center and was applied to the thermal and corrosion protection of uncooled or air-cooled gas turbine blades and to water or fuelaoled rocket engine wds. The early work relating to gas turbine engines involved engine operation at gas temperatures,pressures, and heat fluxes which were relatively low compared to today's engines Ekutoo and Clure, 1953;Schaefer et aL, 1953;Price
et aL, 1973). Schaefer et al. (1953)described experimenta in which SF-99silicone or tributyl borate was mixed with Jp-4 fuel to deposit coatings of silicon dioxide or boric oxide on turbine blades during engine operation. These coatings had a negligible effect on the measured blade temperatures and deteriorated after several hours of engine operation. Other attempts to apply TBC's to turbine hardware are described by Bartoo and Clure (1953).where enameling or glazing processes are evaluated. These coatings also spalled and cracked after several hours of testing. Use of TBC's in rocket engines which generated much higher heat fluxesresulted in considerable thermal protection. However, durability was poor. Curren et al. (1972)described investigations to determine the durability of coatings when subjected to higher heat fluxes. Coating systems investigated were thin layers of molybdenum, nichrome, tungsten, alumina, zirconia, and chromia. The coatings were all either plasma-sprayed or slurry coated and cured in place. Curren et al. (1972)pointed out the need for improved coating application control techniques. TBC Research in the Seventies It has only been in the past 10 years that the feasibility of using TBC's in modern gas turbine engines has been demonstrated. A major breakthrough was the successful completion of testa in a moderately high heat flux 5-75gas turbine reaearch engine by Liebert et al. (1976)and Liebert and Stepka (1977),where the significant heat insulating capability of thin TBC's was experimentallydemonstrated. The experiments showed that a 0.028 cm thick coating of Zr0,-12 wt 7O Y,O, over a 0.010 em thick coating of NiCrAlY provided a 190 K reduction in leading edge metal temperature. This TBC represented a breakthrough in thermal barrier coating technology. Its success was based on three significant departures from prior practices: (1) yttria was used to stabilize the zirconia rather than calcia or magnesia, (2) a NiCrAlY alloy was used as the bond coat rather than nicbrome, and (3) the coating system was applied in two thin layera rather than three (Stecura, 1976; Stecura and Liebert, 1977). Subsequent testa in more advanced and higher heat flux gas turbine engines showed that although this two-layer zirconia TBC system was very promidi, durability at the high gas temperatures and pressures would have to be improved (Sevcik and Stoner, 1978). Since then more durable coatings have been developed through adjustments in bond coat composition (Stecura, July, 1979; Gedwill, 1980),ceramic composition (Stecura, Jan 1979)and plasma spray parameters (Stecura, 1981). Of special note is the observation that the most durable ceramic coatings are those formed from zirconia containing 6 to 8 wt 70yttria.
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Much of this durability research is discussed by Miller et al. (1980), who also describe other investigations such as the response of the zirconia TBC to the presence of inorganic salt deposits (Hodge et al., 1981; Bratton et al., 1982). These studies pertained to ground and marine gas turbine engines where it would be desirable to burn lower quality fuel than is used in aircraft gas turbine engines. The latter studies indicated that it may not be possible to use zirconia TBC’s under conditions of severe contamination, but that in certain less severe conditions they could be very useful. It was also shown by Zaplatynsky (1982) that laser-glazed coatings performed significantly better in dirty fuels than as-sprayed coatings. The development of the TBC’s has also been advanced through heat transfer and aerodynamic research. Heat transfer work dealing with the required high reflectivity of stabilized zirconia is described by Liebert (1978). The high reflectivity property is especially useful in combustor and flameholder applications. A quasi-three-dimensional heat transfer model for accurate calculation of heat flux and temperature gradient through TBC’s and prediction of base metal temperatures is described by Liebert and Gaugler (1980). Experimental aerodynamic investigations (Stabe and Liebert, 1975) have shown that the rough, as-sprayed ceramic coating on aircraft gas turbine vanes reduced aerodynamic efficiency. However, polishing the coating after spraying results in a high aerodynamic efficiency.
Current and Future TBC Research and Development One of the more demanding applications for TBC’s involves their use on convection-cooled turbine vane and blade airfoils, platforms, and rims of advanced aircraft gas turbines. Recent programs have demonstrated that TI3C‘s are sufficiently advanced for current use in advanced Pratt and Whitney JT9D engines to coat first-stage vane platforms (Sheffler et al., 1982). Use of the TBC’s in these engines results in improved base metal durability and simpler internal cooling systems which have the advantage of slightly reduced cooling air requirements. The use of TBC’s on uncooled turbine wheel rims results in reducing transient heat flux into the disk, thus decreasing base metal temperature. Decreased base metal temperatures reduce disk thermal stress and improve base metal durability. Other studies (Siemers and Hillig, 1981) have shown that plasma sprayed TBC’s are sufficiently durable for use on second-stage air-cooled blades of advanced gas turbines and that an integrated design methology should be pursued to take optimum advantage of TBC’s. In addition, it was also shown that durability can be improved through use of low-pressure, plasma-spray bond coats. Control of bond coat and ceramic thickness of TBC’s to conform with design specifications is always necessary, and work described by Fetheroff et al. (1981) demonstrated the feasibility of a computer-aided, fully automatic plasma spray system which assures that precision coating thicknesses are attained. Future improvements may include use of alternative processes for ceramic deposition, such as electron-beam physical vapor deposition (Sheffler et al., 1982) or use of alternative base-metal materials. Fiber metals are one alternative. They are fabricated from randomly oriented alloy fibers which are diffusion bonded together to make up a porous metal matrix (Kascak et al., 1981; Bak, 1983). Preliminary aerodynamic and heat transfer tests (Bak, 1983) have demonstrated that a fiber metal core shaped into an airfoil form and sprayed with a bond coat can support a ceramic which is twice as thick as the zirconia
currently used on solid metal airfoil shells. Efficient convection cooling is provided by passing air through the porous metal matrix. Characterization of Material Properties The industrial designer requires mathematical modeling techniques for TBC design on gas turbine hardware and these models require accurate values for material properties. Specific values of the mechanical and thermal properties of stabilized zirconia and bond coat alloys are reported for a broad range of temperatures by Wilkes and Lagedrost (1973), Shiembob (1977),Kaufman et al. (1978), Liebert (1979), Hodge et al. (1981),Anderson et al. (1982), and Berndt (1983). Thermal properties, such as thermal conductivity and thermal diffusivity, are known within an uncertainty of 20 to 30%. This uncertainty will be decreased when improvements in measurement techniques become available. Characteristics of plasma-sprayed zirconia which make it useful in a thermal barrier coating system are its low thermal conductivity, which is one to two orders of magnitude lower than aircraft metal alloys, and high reflectance, which at practical thicknesses of about 0.033 cm is about 1.5 to 3 times that of typical metallic coatings. The low thermal conductivity provides an effective thermal barrier between the base metal and hot gases. In addition, the coating reflects away much of the gas radiation impinging on the ceramic surface (Liebert, 1978). Thus, less heat is transferred to the interior of the component and therefore less cooling air is needed. Also, since the gas is hotter because of the reduced heat transfer to the coolant, the thermal efficiency of the engine cycle will be increased. Gas-side thermal barrier coating surfaces are hotter than uncoated surfaces. This leads to another advantageous use of TBC’s whereby concentrations of unburned hydrocarbons can more easily burn off the surface and emissions of carbon dioxide and carbon can be reduced. Thin coatings of stabilized zirconia on MCrAlY bond coatings have a partial transmission of heat radiation from the bond coat through the ceramic (Liebert, 1978) and this transmission is reduced as ceramic coating thickness increases. Also, ceramic emittance varies with thickness. These transmittance and emittance variations are significant enough on clean coatings to permit radiation pyrometer detection of changes in ceramic thickness incurred during engine operation. Variations in temperature or color due to spalling can also be noted. The work described by Liebert (1978) also points out that, within an uncertainty of about 1to 3%, the spectral reflectance does not change when temperature, type of zirconia ceramic stabilizer, surface roughness, surface color (very light yellow to light yellow orange), and when view angle was varied from 0 to 80” from a direction normal to the surface. Industrial Tests of Yttria-Stabilized Zirconia TBC’s The interest expressed by various manufacturers, both within and outside the aircraft industry, and the need to evaluate TBC’s under a wide variety of conditions has encouraged the NASA Lewis Research Center to make formal and informal technical assistance arrangements with a number of commercial organizations. Under these arrangements NASA furnished plasma spray coatings on uncooled and convection cooled parts. In turn, industry tested and coated hardware in their engines or in rigs closely simulating their field conditions and provided NASA with data and evaluations based on their particular needs. Finally, recommendations were made by industry for improvement of coating procedures. These arrangements provided industry with hands-on experience with
Ind. EW. chsm.Prod. Rea. Dev.. Vol. 23. No. 3.1984 S47
TBC's and the succ888 of these endeavors encourages industry-wide acceleration of the implementation of TBC technology. The resulta of the work produced by the technical assistancea"entaarediscussed hy Liebert (1979)and Liebert and Levine (1982). To protect proprietary interests, hardware and teat conditions are not always described in great detail. The references deecribe experimental thermal protection and durability testa on: (1) fmtstage turbine vanes for an advanced gas turbine engine, (2) combustor transition sections, (3) coated turbme vanes and blades operating in full-scale military aircraft engines, (4) coated turbine blades operating in a commercial aircraft engine, (5) coated combustor dome, scroll, and turbine nozzle shroud tested in a full-scale military, ground-based gas turbine engine, (6) rig tests of a combustor liner for a gas turbine auxiliary power unit, (7) a partially coated turbine plenum for a vehicular gas turbine engine, and (8) coated valves and head operating in a full-scale diesel engine. Several observations were consistently noted in all of these diverse TBC applications. The coating lowered bane metal temperatures, protected metal parta from warping, increased metal part life, and reduced cooling air requirements. In applications where the gas temperature was higher than the melting point of the base metals, burning and melting were eliminated. For example, the uncoated leading edges of fmt-stage turbine vanes for an advanced gas turbine engine fatigued and melted after initial short test times in full-scale engine cycle testa. When coated with the TBC's, no cracking or melting of the air m l e d metal walls was observed. ALSO, the thermal protection provided by TBC's on flame impingement surfam of combustor transition sections eliminated metal cracking during accelerated mission testing with no deleterious effectto the coating. This result led to the use of TBC's on combustor transition sections of production hardware. The results of the testa on coated diesel engine parta showed that TBC's reduced heat loases by 5% while exhibiting excellent adherence. Based on these early results, the manufacturer believes that ceramic-insulated adiabatic diesel engine parts are potentially useful. More examples of the successful use of TBC's are presented by Liebert (1979)and Liebert and Levine (1982).
Research into TBC Failure Mechanisms The mechanisms of coating failure are not completely understood, and there are areas of disagreement between different investigators. Improved understanding is essential because it seto guide further advancements and leads toward the development of life-modeling capabfitiea. Therefore, basic understanding of failure mechanism has h m e an important part of coatings r m b at the Lewia Research Center. Thermal barrier coatings are generally observed to fail hy spalling within the ceramic layer at a location near the irregular ceramic/bond coat interface. Spalling is characteristic of 'compressive failure", i.e., failure which occurs when the ceramic has been placed in biaxial compression parallel to the interface and tension perpendicular to thin interface. The fact that coatings show a greater tendency to fail in compression than in tension contrasta with the conventional view of monolithic ceramics. Monolithic ceramics may break under tension. Coatings will crack perpendicular to the interface, but such cracks may even be beneficial because they may serve to arrest cracking parallel to the interface. The stresses leading to compressive failure can arise either during rapid heating or after cooling. Stresses are encountered on rapid heating
u-p
7r0.
- "4
-
SEM photomicrograph of the cross &ion of a t h e d barrier mating system after e x p u r e tn 12 I-h rig eyelea at F/A 0.058. Specimen has failed (delaminaLFd) but not yet spalld. Figam8.
because of the temperature gradients which develop throughout the thickness of the coating. Stresses are encountered after cooling because of thermal expansion mismatch between the ceramic and the metallic layers. These cooling stresses may increase as the 'stress-free temperature" (Sevcik and Stoner, 1978)of the ceramic/ metallic system increases by any of several possible hightemperature stress relaxation processes. Cooling mode compressive stresses tend to maximize near the interface where failure is observed. Heating mode compressive stresses maximize a t the outer surface. One author has argued that heating mode stresses could still lead to the observed failure near the ceramic/bond coat interface (McDonald and Hendricks. 1980). Recent experiments showed that TBC failures in fun" and Mach 0.3 burner rig testa initiate as a result of the stresses encountered on cooling (Miller and Lowell, 1982). Failure could even be initiated on cooling after a single, severe isothermal exposurebut only if the specimen was exposed in an oxidizing environment. The initiation of failure by heating-mode stressea can be ruled out because as-sprayed and pre-exposed specimens do not fail after up to loo0030.second heating cyclea This 30-8 heating cycle is thought to be long enough to allow the development of maximum heating stresses but short enough to ensure that the bond coat does not attain maximum temperature. Therefore no oxidation or other type of high-temperature degradation can occur during these short heating cycles. T w o conditions were chosen for the pre-exposcues. In one, a specimen was exposed in an inert atmosphere furnace at a temperature which was high enough to caw failure on cooling in a single cycle if the environment were oxidizing. In the other, a specimen was exposed in the burner rig for a single cycle. Ita duration was longer than that required to fail the coating when cycled hourly. The absence of coating failure after many subsequent heating cycles further supports the assumption that damage must to initiate during the cooling mode for specimens ex& Mach 0.3 burner rigs. Also, the absence of failure on heating of the preex& specimens underscores the importance of both oxidation and cycle frequency to the initiation of coating failure in the cooling mode. A cross-sectional micrograph of a specimen which had been cycled hourly in the combustion gases of a Mach 0.3 burner rig is shown in Figure 3. The specimen has suffered a delamination type of failure but it has not yet spalled. This figure is also typical of specimens which failed after a single, more severe isothermal exposure. An oxide layer, primarily alumina, has grown (throughout the course of exposure) on the NiCrAlY bond coat a t the ir-
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Ind. Eng. Chem. Prod. Res. Dev., Vol. 23,No. 3, 1984
regular interface with the ceramic layer and there is also internal oxidation at the splat boundaries. A delamination crack has formed in the ceramic layer parallel to and near the interface with the bond coat and a large gap has opened in this region. The gap has opened to relieve residual stress in the ceramic. The size of this gap must be directly related to the magnitude of this residual stress. Thus if this residual stress could be minimized during processing, the tendency for the coating to delaminate should decrease. The delaminated region subsequently spalls on heating because it behaves as a rapidly heating, constrained plate. That is, the delaminated region heats up rapidly because the heat is not efficiently conducted out from its unattached underside. The cooler surrounding regions constrain the delaminated ceramic layer as it expands. The resultant stresses cause buckling leading to crack extension and, eventually, spalling. If delaminated specimens are cycled in a furnace the heat flux is too low to cause spalling of the ceramic. Rather, tensile cracks perpendicular to the interface develop in the delaminated region. These tensile cracks must result from the increased circumference of the material above the gap. Observation of these cracks at the surface provides the critical indication of failure in the furnace tests. In summation, stresses leading to delamination occur on cooling and arise from thermal expansion mismatch between the ceramic and the metallic layers. Spalling subsequently occurs during heating in burner rig tests. Cracking is observed at the surface in furnace tests. Specimen test lives are strongly influenced by oxidation, cycle frequency, and residual stress. The magnitude of these cooling-mode stresses may be affected by any of several processes which effect the ceramic-metal system, e.g., mechanical “creeping” of the ceramic (Firestone et al., 1982), bond coat flow, bond coat-substrate diffusion, phase transformations (Miller et al., 1981) or oxidation. Therefore the exact failure mechanism may be complicated-involving the interaction of several processes. Of these, oxidation of the bond coat at the irregular interface with the ceramic is the time-dependent process which most highly correlates with coating behavior in burner rig and furnace tests (Miller and Lowell, 1982). Even though heating mode stresses do not play a significant role in the initiation of failure at lower heat fluxes, it is possible that they may play an important role at higher heat fluxes. Experiments are underway to determine whether failure on heating becomes significant at higher heat fluxes. Preliminary results indicate that optimum coatings do not spall at heating rates greater than those encountered in engines. The response of the ceramic layer to stress must be characterized. This is because failure occurs within the ceramic layer even though stresses arise as the environment acts on the ceramic/metal system. In one study, coating test lives for zirconia-yttria ceramics were shown to be highly correlated with the presence of a quenched-in, “martensitically nontransformable” tetragonal phase (Miller et al., 1983). Compositions rich in this phase also contained minor but possibly significant amounts of the monoclinic phase. Possible relationships between composition and performance are discussed by Miller et al. (1983). These phases, which were observed in the assprayed ceramics, are quite stable below about 1470 K (Miller et al., 1981), and so thermally activated phase transformations do not appear to be a first-order factor in coating failure. The structure of the ceramic layer, as controlled by the processing, is a crucial factor controlling the tolerance of
the ceramic to cyclic stress. The density of the ceramic layer is one important factor (Tucker et al., 1976; Stecura, 1981). Attempts to further control structure through processing are described by Anderson and Sheffler (1983). Monolithic and single-crystal zirconia-based ceramics have been studied extensively (Heuer et al., 1981; Cawley, 1984). These studies involve such techniques as X-ray diffraction analysis, transmission electron microscopy, microhardness, and fracture toughness measurements. They have provided a background of basic research which can aid understanding of plasma-sprayed, yttria-stabilized zirconia. Tensile adhesion tests (Levine, 1978; Shankar et al., 1982), acoustic emission analysis (Berndt; Shankar et al., 1983; Berndt; to be published) and residual stress analysis may be expected to contribute to understanding the fracture processes leading to coating failure. Fracture toughness measurements using methods developed specifically for plasma sprayed ceramics (Berndt, 1981) are also required. Concluding Remarks TBC technology has advanced to the point were they are now being employed, to a limited extent, in the turbine section of production gas turbine engines. Work is continuing to further improve, understand, and model coating behavior. Efforts to transfer the technology to other aerospace and nonaerospace applications are also continuing. Literature Cited Anderson, C. A.; Lau, S. K.; Bratton. R. J.; Lee, S. Y.; Rleke, K. L.; Allen, J.; Munson, K. E. Feb 1982, NASA CR-165619, DOE/NASA/0110-1. Anderson, N. P.; Sheffler, K. D. Sept 1983, NASA CR-168251, PWA-577729. Bak. D. J. Des. News 1983, 39(1), 92. Bartoo. E. R.; Clure, J. L. July 16. 1953, NACA RM E53E18. Berndt, C. C. I n “Proc. of the Inter. Conference on the Ultrastructure Processing of Ceramics, Glasses and Composites”, Hench, L. L.; Ulrlch, D. R., Ed.; to be published by Wlley. Berndt, C. C. I n “Proceedings of Surface Engineering”; NATO Advanced Study Instltute; Singhal, S. C.; Kossowski, R. Ed.; to be published by Plenum Press. Berndt, C. C.; McPherson, R. Mater. Sci. Res. 1981, 14, 619-28. Bratton, R. J.; Leu, S. K.; Lee, S. Y. July 1982, NASA CR-165545. Cawley, J. D. March 1984, NASA TP-2286. Curren, A. N.; Grisaffe, S. J.; Wycoff. K. C. Jan 1972, NASA TM X-2461. Fetheroff, C. W.; Derkacs, T.; Matay, I.M. Mar 1981, NASA CR-165418, TRW ER-8019-F. Firestone, R. F.; Loaen. W. R.; Adams, J. W.; Bill, R. C., Jr. Ceram. Ens. Sei. Proc. 1982,-3, 158-71. Gedwill, M. A. Sept 1980, NASA TM-81567, DOEINASA/2593-18. Heuer, A. L.; Hobbs, L. W., Ed. Science and Technology .. of Zirconia, A&. Ceram. 1981, 3. Hodge, P. E.; Miller, R. A.; Gedwill, M. A.; Zaplatynsky, I. March 1981, NASA TM-81716. .... - . . . -. Kascak, A. F.; Liebert, C. H.; Handschuh, R. G.; Ludwig, L. P. Dec 1981, NASA TP-1942, AVRADCOM TR-814-7. Kaufman, A.; Llebert, C. H.; Nachtlgall, A. J. Dec 1978, NASA TP-1322. Levine, S. R. May 1978, NASA TM-73792. Llebert, C. H. Thin SolM Fllms 1978, 53, 235-40. Liebert, C. H. Thin SolM Nlms 1979, 64, 329-33. Llebert, C. H.; Gaugier, R. E. Thin SolM Nlms 1980, 73, 471-5. Llebert, C. H.; Jacobs, R. E.; Stecura, S.; Morse, C. R. Sept 1976, NASA TM X-3410. Liebert, C. H.; Levine, S. R. Sept 1982, NASA TP-2057. Llebert, C. H.; Stepka, F. S. J. Aircr. 1977, 14, 487-93. McDonald, G.; Hendricks, R. C. Thin SolM Films 1980, 73, 491-6. Miller, R. A.; Garilck, R. G.; Smialek, J. L. Am. Ceram. SOC.Bo//. 1983, 62, 1355-8. Miller, R. A.; Levine, S. R.; Stecura, S. Jan 1980, AIAA-80-0302. Miller, R. A.; Lowell, C. E. Thin SolM Films 1982, 95, 265-73. Miller, R. A.; Smialek, J. L.; Garlick, R. G. A&. Ceram. 1981, 3 , 241-53. Price, H. G. Jr.; Schacht, R. L.; Quentmeyer, R. J. Nov 1973, NASA TN P 7392. Schafer, L. J., Jr.; Stepka, F. S.; Brown, W. B. Mar 30, 1953, NACA RM E53A19. Sevclk, W. R.; Stoner, B. L. Jan 1978, NASA CR-135360, PWA-5590. Shankar, N. R.; Berndt, C. C.; Herman, H. Ceram. Eng. Sci. PIoc. 1982, 3(NO. 9-10), 772-92. Shankar. N. R.: Berndt. C. C.: Herman. H.: Ranaswamv. S.: Am. Ceram. Soc. Bull. 1983, 62(No. 5), 614-9. Sheffler, K. D.; Grazlani. R. A.; Slnko, G. C. April 1982. NASA CR-167964, PWA-55 15-176. Shiembob, L. T. April 1977, NASA CR-135183, PWA-5521. Siemers, P. A.; Hllllg, W. B. Aug 1981, NASA Cr-165351. Stabe, R. G.: Liebert, C. H. Jan 1975, NASA TM X-3191.
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349
Ind. Eng. Chem. Prod, Res. Dev. 1984, 23, 349-360 Stecura, S. Sept 1976, NASA TM X-3425. Stecwa, S. Jan 1979, NASA Th4-78976. Stecwa, S.July 1979, NASA TM-79206. Stecwa, S.March 1981, NASA TM-81724. Stecwa, S.; Liebert, C. H. U.S. Patent 4055705, Oct 1977. Tucker, R. C., Jr.; Taylor, T. A.; Weatherly, M. H. In “Proceedings of the 3rd Conference on (3es Turbine Materlals in a Marine Environment”, Bath,
England, Sept 20-23, 1976; Session VII, Paper 11. Wllkes, K. E.; Lagedrost, J. F. March 1973, NASA CR-121144. Zaplatynsky, I . Thin SolM Fllms 1982, 9 5 , 275-84.
Receiued for review March 1, 1984 Accepted June 11, 1984
Catalyst Deactlvatlon during Direct Coal Liquefaction: A Revlew Deepak S. Thakur’ InterNorth, Inc., Corporate Research Dlvlsbn, Omaha, Nebraska 681 17
Michael 0. Thomas Sandla National Laboratorles, Albuquerque, New Mexico 87185
A review of the research and development efforts in catalytic direct liquefaction has shown that, Justas in petroleum residuals refining, catalyst lifetime is more important than initial activity. Deacthration studies indicate that contamination by 20-30 wt % carbonaceous material and 2-8 wt % inorganics results in deactivation best described by the shell progressive deactivatkm mechanism. Metals contamination is confined to the outer several hundred microns of the catalyst, and pore volume distributions show a marked change toward higher average pore size in support of this mechanism. Close examination of the mechanistic behavior of catalysts in direct liquefaction indicates that their primary function is hydrogenation of solvents.
I. Introduction The present uncertainty in the supply of liquid fuels has created interest in developing alternate sources. Coal, oil shale, and tar sands are the main candidates. Abundant supplies of coal in most parts of the world provide a basis for development and optimization of this alternate source for liquid fuels. In Germany, the shortage of petroleum reserves during World War I1 caused the initiation of extensive development efforts for conversion of coal into liquids. German technology, which involved the hydrogenation of coal into liquid products under high pressure and temperature, served as a starting point for current research efforts. The ultimate goal in these efforts is to reduce the capital investment by (1)improving thermal efficiency and hydrogen utilization, (2) carrying out the process under less severe conditions, and (3) broadening the range of acceptable feed-coal properties. A number of coal liquefaction processes have been developed in recent years; they can be grouped into three broad categories: (1)pyrolysis, (2) direct liquefaction, and (3) indirect liquefaction. Pyrolysis involves the heat treatment of coal at temperatures above 400 “C and low pressures to convert the coal into gases, liquids, and char. In direct liquefaction, coal, a coal-derived hydrogen donor solvent and molecular hydrogen (gas) react under temperature and pressure with or without catalysts to yield liquid hydrocarbons. Indirect liquefaction is carried out by allowing coal to react with oxygen and steam to produce carbon monoxide and hydrogen. Subsequent hydrogenation of carbon monoxide in the presence of a catalyst produces liquid products. The main purpose of this paper is to review the results of studies on the deactivation of catalysts used in direct
coal liquefaction. Much of the work on catalyst deactivation in coal liquefaction processes has been performed at Amoco, PETC, and Sandia National Laboratories in conjunction with ongoing processes, especially H-Coal. These studies will be reviewed in detail. Because of the difficulties associated with model compound studies, only information collected in coal processing is incorporated in this discussion. The use of a catalyst in direct liquefaction has several advantages: (1)It minimizes hydrogen consumption. (2) It permits milder temperature conditions than those required in noncatalytic coal liquefaction (thermal liquefaction). (3) It produces lesser amounts of undesirable products like gases and heavy resids. (4) It increases distillate (middle range oil) production. Catalytic coal hydrogenation also has certain disadvantages. The most important are the complexity of the reaction system, the catalyst cost, and catalyst deactivation. Before proceeding to discuss the results on catalyst deactivation, it is necessary to address the salient features of the process and also the principal functions that catalysts play in coal liquefaction. 11. Thermal Liquefaction Process As previously indicated, goals for coal liquefaction include the production of distillate products in high yields, and minimum consumption of hydrogen. Many current mechanistic descriptions of coal conversion have two common themes. (1)Liquefaction (dissolution) of coal, which is a primary step in upgrading coal into liquid fuels, is a thermally induced free-radical reaction (Wiser et al., 1971; Larsen, 1978). In general, organic free-radical reactions are classically nonselective and hence they may or may not require the presence of a catalyst. (2) Many of the liquefaction products can be converted to lighter fractions. The conversion of coal into such materials may follow many parallel (Shah et al., 1978; Neavel, 1976) or series paths (Thomas and Traeger, 1979; Weller et al., 1951). This step is accompanied by substantial reduction in average molecular weight and heteroatom (S, N, and 0) content. The processes developed to date utilize hydrogenolysis of coal in the presence of a hydrogen-donor solvent. Two descriptive representations of coal liquefaction are given in Figure 1. There is considerable debate as to the most appropriate representation for liquefaction; for our initial
0196-4321/84/1223-0349$01.5~~0 0 1984 American Chemical Society
