Thermal and Environmental Barrier Coatings for SiC/SiC CMCs in Aircraft Engine Applications

Int. J. Appl. Ceram. Technol., 1 [4] 291-301 (2004)

Ceramic Product Development and Commercialization

IIIIIntrntrntrntrntroductionoductionoductionoductionoduction

Ceramic matrix composite (CMC) materials haverecently received significant attention from both govern-ment and industry as a key material technology for meet-ing future propulsion needs, as CMCs provide a signifi-cant potential improvement in fuel consumption andthrust-to-weight ratio compared to metallic materials.1 Aceramic composite material comprised of silicon carbide(SiC) fiber with a silicon carbide matrix is one such ma-terial. The SiC/SiC CMC material developed jointly un-der U.S. Government (NASA, DoD, and DoE) and GEprograms utilizes the preferred Melt Infiltration (MI) pro-cess to achieve the desired thermomechanical properties.The MI SiC/SiC material has the capability of operatingat an average temperature exceeding 1204ºC (2200°F).

The feasibility of implementing SiC/SiC CMC ma-

terial technology into turbine engines has been studied atGE Aircraft Engines over the last 10 years. As shown inFig. 1, the candidate components include all major hard-ware throughout the turbine hot section. If the designrequirements are met successfully for this wide range ofcomponents, engine performance benefits should be re-alized in a variety of applications including military andcommercial.

The engine performance benefits associated withimplementing CMCs are attributed to the material’s rela-tively low density (approximately one-third of typical Ni-base superalloys) and higher temperature capability. Ad-ditional weight benefit may be realized when applyingCMCs to rotating airfoils because of the impact on theentire rotor structure. In the future, the temperature ca-pability of the current SiC/SiC CMC system should im-prove by an additional 93ºC (200°F), or a potential ben-efit over superalloys exceeding 204ºC (400°F). The highertemperature capability enables lower cooling air require-ments, leading to a potential reduction in NO

x emissions

for combustors and an increase in thrust-to-weight ratio

Thermal and EThermal and EThermal and EThermal and EThermal and Envirnvirnvirnvirnvironmental Bonmental Bonmental Bonmental Bonmental Barrier Coatingsarrier Coatingsarrier Coatingsarrier Coatingsarrier Coatingsfor Sfor Sfor Sfor Sfor SiC/SiC/SiC/SiC/SiC/SiC CMCs in AiriC CMCs in AiriC CMCs in AiriC CMCs in AiriC CMCs in Aircraft Ecraft Ecraft Ecraft Ecraft Engine Angine Angine Angine Angine Applications*pplications*pplications*pplications*pplications*

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GE Aircraft Engines, Cincinnati, OH

Accommodating relatively high engine component surface temperatures is a significant challenge that must be overcomein order to fully realize engine performance benefits associated with implementing CMC materials having relatively low densityand higher temperature capability. The design surface temperature for selected components is expected to exceed the 1414ºC(2577°F) melting point of silicon, a key CMC constituent, and thus protective coatings are required. This paper discusseschallenges in developing the coating system, specifically addressing issues related to the selection of materials, designing thecoating architecture, and the test methodology.

* Presented at the 27th Annual International Conference on Advanced Ceramics andComposites, Cocoa Beach, FL, January 27th, 2003, and at the Environmental BarrierCoatings for Microturbine and Industrial Gas Turbine Ceramics Workshop, Nashville, TN,November 19th, 2003.

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292 Vol. 1, No. 4, 2004International Journal of Applied Ceramic TechnologySpitsberg and Steibel

for turbines that incorporate CMC airfoils, as shown inFig. 2. However, accommodating the relatively high com-ponent surface temperatures is a significant challenge thatmust be overcome in order to fully realize these engineperformance benefits. The design surface temperaturesare expected to exceed the 1414ºC (2577°F) silicon melt-ing point for selected components, thus, protective coat-ings are required. Even with the surface temperature be-low the melting point of key CMC constituents, envi-ronmental and thermal protection of the component sur-face most likely will be needed.

CCCCCurrurrurrurrurrent Sent Sent Sent Sent Status of Etatus of Etatus of Etatus of Etatus of Envirnvirnvirnvirnvironmental and onmental and onmental and onmental and onmental and ThermalThermalThermalThermalThermalBBBBBarrier Coatings for CMCsarrier Coatings for CMCsarrier Coatings for CMCsarrier Coatings for CMCsarrier Coatings for CMCs

When exposed to an oxidative environment, SiCforms a protective slow-growing silicon oxide scale (SiO

2).

However, it has been experimentally demonstrated2-4 thatSiC can undergo a significant weight loss in a water-va-por containing environment, as shown in Fig. 3. Thisoccurs by the surface recession of the silica scale due tothe reaction with the water vapor as follows:

SiO2 + 2H

2O = Si(OH)

4 (gas) (1)

Fig. 1. CMC applications in the hot section of jet engine.

Fig. 2. CMC payoffs in NOx emissions (left) and Specific Fuel Consumption (SFC) and Thrust-to-Weight ratio (right).

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Thermal and Environmental Barrier Coatings for SiC/SiC CMCs in Aircraft Engine Applicationswww.ceramics.org/ACT 293

The surface recession mechanism is very sensitive tothe gas velocity and can be dramatically accelerated as the

velocity increases.5,6 For turbine engine conditions, thevelocity factor is usually included in the gas Mach num-ber and/or heat transfer coefficient. Surface recession wasindeed observed during the first field engine tests of SiC/SiC CMC combustor liners in land-based power genera-tion engines.7,8

Significant CMC surface recession can be expectedin a typical combustion environment containing about10% water vapor at a pressure of 10-30 atm, especiallysince these conditions must be withstood for lifetimesexceeding 1000 h. The recession may reduce the net-sec-tion thickness of the component wall or contribute todimensional changes in the case of airfoils.

An environmental barrier coating (EBC) system ca-pable of protecting the CMC component surface fromrecession over the life of the component was developedunder the EPM program.2,9 As shown in Fig. 4, the sys-tem consists of three layers: a silicon layer adjacent to theCMC, a transition layer consisting of a mixture of mul-lite (2Al

2O

3 - 3SiO

2) and barium strontium alumino-sili-

Fig. 3. Mass change and surface recession for sintered SiC at1200ºC in 90% steam.

Fig. 4. As-processed (by air plasma spray) and heat-treated 3-layer EBC system with BSAS environmental barrier layer; X-ray mapsshow higher Sr and Al content, as well as lower Si content in the secondary phase (brightly imaging phase within BASA layer).

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cate (BSAS), and a protective layer of stoichiometric BSAS(25 mol% BaO + SrO, 25 mol% Al

2O

3, 50 mol% SiO

2).

As deposited using a conventional plasma-spray techniquefollowed by a prolonged heat treatment,2 the top layerpredominantly exhibits a Celsian crystal structure, with asmall percentage of a secondary phase with a strontium-rich composition (Fig. 4). This three-layer system is oftenreferred to as a baseline or Gen 1 EBC coating system.

EBC development efforts have continued since thetermination of the NASA EPM program in 1999. Theseactivities have included microstructure and processingimprovements, as well as scale-up and component leveldemonstration.10,11 In addition, development efforts atGeneral Electric have also been focused on developingcoating systems with higher temperature capability forlong-life applications in power-generation engines, as wellas a thermal-environmental barrier coating (T/EBC) sys-tem with significantly higher surface temperature capa-bility than the baseline BSAS system. The T/EBC coat-ing system will potentially consist of a relatively low-con-ductivity ceramic top layer material selected from a mate-rial family other than silicates12-14 due to the relativelyhigh temperature capability requirement. Currently, thefeasibility of establishing a T/EBC system that survivescyclic testing at a surface temperature approaching 1700ºC(3100ºF) for 50 h has been demonstrated. Further effortsto mature this coating system for the full life requirementare necessary, and the related challenges will be discussedin subsequent sections.

Coating DCoating DCoating DCoating DCoating Design Resign Resign Resign Resign Requirequirequirequirequirementsementsementsementsements

Prior to initiating development of a new coating sys-tem, the coating performance goals dictated by the com-ponent design must be defined and subsequently trans-lated into specific requirements for the coating materialsystem. Specific requirements for the coating performancein combustors and airfoils are summarized in Table I. Thecombination of high surface temperature, moisture con-tent, and high gas velocities (Mach) generate harsh reces-sion conditions in jet engines, and these conditions aretypically more severe than power-generation engines. Inaddition, jet engine field experience on metallic compo-nents suggests that corrosion is also a concern. The cor-rosive environment may result from sea salt or sand-typedeposits of calcium-magnesium-alumino-silicates (CMAS)that tend to melt at the high application temperatures.The coatings must also withstand erosion potentially in-duced by particles passing through the hot section of theengine. Larger particles (>500 µm) can potentially induceimpact damage to the coatings, a phenomenon commonlyreferred to as foreign object damage (FOD).

The various engine operation conditions that thecoatings must withstand are schematically summarizedin Fig. 5. In addition to the specific requirements high-lighted previously, Fig. 5a depicts the severe temperaturegradients incurred through the coating thickness. Thethermal shock conditions resulting from the cyclic natureof the engine operating cycle are shown schematically in

TTTTTable I. able I. able I. able I. able I. TTTTTypical Coating Dypical Coating Dypical Coating Dypical Coating Dypical Coating Design Resign Resign Resign Resign Requirequirequirequirequirements for Combustors and Airements for Combustors and Airements for Combustors and Airements for Combustors and Airements for Combustors and Airfoilsfoilsfoilsfoilsfoils

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Fig. 5b. Thus, maintaining the coating mechanical in-tegrity is a significant challenge because a CTE differen-tial between coating layers is inevitable.

MMMMMaterial Saterial Saterial Saterial Saterial System Dystem Dystem Dystem Dystem Devevevevevelopment Challengeselopment Challengeselopment Challengeselopment Challengeselopment Challenges

Successful development of a new coating system forSiC-based CMC turbine components depends on theability to address four major challenges: 1) Establishingthe test methodology, or more specifically, demonstratingthe performance goals on a laboratory-scale prior to in-troducing a coating system into the engine; 2) Identify-ing a material composition with adequate recession resis-tance; 3) Optimizing the coating deposition process toattain acceptable coating thickness variation; 4) Design-ing a coating system with the desired thermomechanicalcapability. These challenges are each discussed further inthis section.

TTTTTesting Mesting Mesting Mesting Mesting Methodologyethodologyethodologyethodologyethodology

An effective laboratory testing approach must ad-equately assess the potential for the coating to survive inthe engine. Besides investigating the capability of newcoating materials, the testing methodology is complicatedby the need to address the requirements for new compo-nents or different engines. These advanced engine require-ments may exceed the capability of existing test equip-

ment. Consequently, existing test methods must be modi-fied to support implementation of the coating system re-quired for CMC turbine engine components. Table IIsummarizes currently employed laboratory tests and pro-vides a comparison of the respective test capabilities rela-tive to the engine conditions. Unfortunately, none of theexisting tests simulates all of the required conditions.Therefore, a combination of the lab tests must be per-formed to assess the coating performance.

Selected test methods have been enhanced to en-able adequate characterization of the CMC coatings. Astate-of-the-art Laser Gradient Rig test has been devel-oped at NASA15 and adapted to enable the testing of EBCsand T/EBCs. Most recently, this test was upgraded toincorporate a steam environment. Also, a Flame Jet Gra-dient Test has been developed at GE Aircraft Engines.This test utilizes a mixture of propane and oxygen gasesthat impinge on the 25-mm disk-shaped specimen. Thespecimen is cooled from the back side with compressedair, and the temperature on the front and back sides ismeasured using a pyrometer. This system is capable ofattaining surface temperatures up to 1760ºC (3200ºF)while imparting through-thickness temperature gradientsapproaching 537ºC (1000ºF), as well as various thermalshock conditions representative of the engine cycle. Thegas velocities are relatively high compared to the LaserGradient Rig, with the moisture content being about 7-10% at atmospheric pressure. Even though both of these

Fig. 5: a) Potential failure mechanisms for T/EBC coatings and b) schematics for typical jet engine temperature cycle.

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TTTTTable II. Comparison of Conditions of Laboratorable II. Comparison of Conditions of Laboratorable II. Comparison of Conditions of Laboratorable II. Comparison of Conditions of Laboratorable II. Comparison of Conditions of Laboratory y y y y TTTTTests with the Eests with the Eests with the Eests with the Eests with the Engine Conditions.ngine Conditions.ngine Conditions.ngine Conditions.ngine Conditions.

Legend: FCT – furnace cycle test; HSCF – high steam cyclic furnace; Laser – Laser Gradient Test at NASAGlenn Research Center; Jet – Flame Jet Gradient Test at GE Aircraft Engines; Question mark - indicatesthat this condition is not precisely controlled.

Fig. 6. Burner Rig Test at GE Aircraft Engines.

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conditions are not precisely controlled, the test can besufficient for relative coating screening and ranking interms of moisture and, in particular, gas velocity sensitiv-ity. Comparative testing of the same T/EBC systems inboth the Laser and Flame Rigs typically yields similar re-sults in terms of number of hours “survived” in the test.The Flame test is somewhat more severe due to the radialtemperature gradients induced by the localized flame jetas compared to the relatively uniform heating by the lasersource in the Laser Rig.

The combined hot corrosion and oxidation test(CHCO), commonly referred to as a Burner Rig test andtypically used for evaluation of metallic environmentalcoatings, has been adapted to accommodate testing ofEBCs and T/EBCs. As illustrated in Fig. 6, the enginefuel is burned in the combustor at a specific fuel-to-airratio and the exhaust gases are directed through the nozzleonto the coating specimen. The specimen orientation rela-tive to the flame may be adjusted to alter the gas im-pingement angle. The gas temperature and Mach num-ber are controlled in this test. The typical moisture con-tent attained in this test from the jet fuel combustion isabout 10%. In addition, a sea salt solution may be in-jected into the flame jet to enable testing under the hotcorrosion conditions. Since the test is at atmospheric pres-sure, the salt concentration should be adjusted to accountfor the pressure difference between the test and applica-tion conditions. This general testing approach, describedelsewhere,16,17 is routinely used by GE Aircraft Enginesfor the CHCO testing of metallic environmental coat-ings for airfoils. Similarly, ceramic particles of selectedsize and quantities can be injected during the test cycle,in this case into the flame, to enable evaluating the coat-ing response to erosion and impact conditions. This ap-proach is analogous to the testing methodology employedon ceramic TBCs for airfoils.18 Due to these advantages,the Burner Rig test is actively used as a part of the EBCand T/EBC evaluation by GE Aircraft Engines. The1315ºC (2400ºF) temperature limitation should be noted,however.

RRRRRecession Iecession Iecession Iecession Iecession Issue for Combustorsssue for Combustorsssue for Combustorsssue for Combustorsssue for Combustors

CMC component surface recession can present anissue for both cooled and uncooled components. Thus,the EBC should be designed to minimize surface reces-sion. Based on the SiC recession model developed byNASA,8 the Gen 1 BSAS-type coating recession rate isabout two orders of magnitude less than the SiC rate

(based on a generally agreed upon assumption for the es-timated silica activity in BSAS of approximately 0.01).According to this model, which was developed for theSiC forming SiO

2 scale (with the unit silica activity) for

laminar gas flow conditions, the surface recession K, inµm/h, due to silica volatilization by the reaction with watervapor (Eq. 1) is determined by the equation:

K ~ 10–5654/[T+273] × P1.5 × V0.5 × aSiO2

(2)

where P is total pressure at 10% H2O, atm; V is the gas

velocity, m/s; a is the activity of silica; and T is the tem-perature, ºC.

The model was validated for gas velocities up to about0.1 Mach, and the recession rate was shown to be inde-pendent of the gas impingement angle. An approxima-tion of the field engine test results on BSAS-coated com-bustion liners in land-based engines11 indicates that coat-ing recession was reasonably close to the predicted be-havior. An assessment of the BSAS recession rate for theaircraft turbine combustor liner applications is shown inFig. 7. Total recession exceeding 200 µm is predicted atthe higher temperature and pressure conditions. Thisapproach likely underestimates the actual recession ratessince these types of conditions are outside of the gas ve-locities for which the model was developed and validated.Indeed, BSAS surface recession was observed in the at-mospheric Burner Rig test at temperatures above 1371ºC(2500ºF) and Mach 0.3 after only 50 h, even though watervapor content was only on the order of 10%, and themodel would not predict a visible recession for these con-ditions. The initial stages of surface recession include poreformation and pitting, preceding the thickness loss atlonger exposures, as opposed to the uniform reduction ofthe coating thickness. Surface phase changes somewhatsimilar to those observed in the engine tested liners afterthousands of hours of exposure11 were also observed in

Fig. 7. Estimated recession rates at combustor conditions inaccordance with the NASA model.8

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the Burner Rig test. Additional work is needed to de-velop a mechanistic understanding of the BSAS recessionmechanism and develop models adequately describing thismechanism at a given heat transfer condition. At the sametime, it is clear that the coating surface recession for theBSAS-type coatings under the jet engine combustor ap-plication conditions can present a significant risk due toimposed variations in the coating thickness.

Coating Coating Coating Coating Coating Thickness Thickness Thickness Thickness Thickness VVVVVariationariationariationariationariation

Efforts focused on understanding the implication ofcoating thickness variation on the performance of a cooledcombustor liner should consider the effects on the sur-face temperature and the CMC componentthermomechanical stresses. If the EBC-coated CMC com-bustor wall is considered as one system under the tem-perature gradient, a significant portion of this gradient is“taken” by the coating, thus providing the potential tomaintain the CMC component stress within the designlimits of the CMC strength capability. As shown by sche-matic finite-element calculation results provided in theupper portion of Fig. 8, the minimum coating thicknessrequirement is driven by the CMC stress limit. Hence,one may be tempted to satisfy this minimal thickness re-quirement by increasing the nominal or mean coatingthickness target for the as-deposited coating. However, as

shown in the lower portion of Fig. 8, the coating surfacetemperature capability may be exceeded if the coatingthickness is too high. Analysis shows that the acceptablethickness range, as indicated by the bar in the bottom ofFig. 8, may be as narrow as 100 µm. A 100-µm thicknesstolerance imposes a strict requirement on the coatingdeposition process as well as tolerance for the surface re-cession. The upper thickness limit will be increased if thecoating surface temperature capability is improved, andoptions include modifying the coating chemistry or add-ing a thermal barrier layer.

AAAAApprpprpprpprpproach for a Hoach for a Hoach for a Hoach for a Hoach for a High-Tigh-Tigh-Tigh-Tigh-Temperemperemperemperemperaturaturaturaturature e e e e T/EBC ST/EBC ST/EBC ST/EBC ST/EBC Systemystemystemystemystem

As was discussed in the previous sections, the designsurface temperature requirements for the CMC compo-nents are expected to exceed the BSAS thermal stabilitylimit. Thus, in addition to increasing the thickness range,high-temperature-capability materials with low thermalconductivity are required. At the same time, environmen-tal protection of the CMC substrate from the recessionand hot corrosion will still be needed. One approach to“building” a high-temperature T/EBC system would con-sist of adding a high-temperature ceramic layer, for ex-ample, zirconia-based, on top of the three-layer BSAS-type EBC system. A transition layer with an intermediateCTE will be needed to accommodate the CTE mismatchbetween the BSAS layer (about 5 ppm/ºC) and the toplayer (about 10 ppm/ºC in the case of stabilized zirco-nia). This approach is schematically shown in Fig. 9. Thisapproach will result in a five-layer system which presentschallenges in terms of processing, meeting thickness re-quirements, and cost. Therefore, new types of materialsthat could replace the three-layer EBC system, providethe environmental protection, and serve as a bond coatfor the top ceramic layer are currently sought.

For the multi-layer T/EBC system, the major mate-rial challenges arise in the areas of microstructure andphase stability as well as chemical compatibility of theindividual layers, sintering of the ceramic top layer, andmechanical integrity of the system with the CTE gradi-ent while operating under the temperature gradient cy-clic thermal shock conditions.

While defining a five-layer T/EBC coating systemarchitecture, the following factors are typically taken intoaccount: temperature capability of the individual coatinglayers as well as interfaces between the layers, thermalconductivity of each layer, minimal thickness requirementfor the environmental barrier, and the maximum total

Fig. 8. Schematics of finite element calculations for the CMCstress and coating surface temperature as a function of thecoating thickness in commercial combustor liner (blue bar in thebottom on the figure represents acceptable coating thicknessrange).

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coating thickness requirement from the design. Cyclicexposures to the maximum temperature with the tem-perature gradient result in a complex stress state. An ex-ample of the stress distribution in a five-layer coating sys-tem with zirconia-based topcoat is shown in Fig. 10. Thecalculations are made for the cases of plasma-sprayed andEB-PVD deposited top layers, assuming a stress-free stateat the top layer deposition temperature for each case.19

Sintering and aging phenomena in the layers as well asthe coating cracking will result in changes in the mechani-cal properties of the layers and the stress state. Furthermodeling efforts are required to adequately describe thesechanges and develop coating lifing models. Such modelsare critical for assessment of the coating performance asthe coating architecture, processing methods, and the layermaterials selection changes evolve as required by the de-sign in various applications of the T/EBC CMC systems.

To date, the feasibility of a five-layer T/EBC coatingsystem with EB-PVD top ceramic layer has been demon-strated for 50 1-h cycles at a surface temperature of 1675ºC(3047ºF) and a back-side temperature of 1106ºC (2023ºF)in the NASA Laser Gradient Rig. Coating spallation wasnot observed during this test, as shown in Fig. 11. Also,the performance has been validated in the Flame Jet Gra-dient test with similar temperature conditions but underthermal shock: 600 5-min cycles with 20 sec heat-up andcool-down cycle. This coating system has been success-

fully applied on CMC airfoil components for subsequentengine testing.

Upcoming engine tests planned at GE Aircraft En-gines will be absolutely critical for correlating conditionsof the lab testing to the real engine conditions. With theuncertainty in the recession mechanisms and predictionsof the recession rates for the BSAS coatings, engine testresults would provide the basis for refining the perfor-mance predictions. Engine testing on an EBC-coated

Fig. 9. Schematics of approach to designing high-temperature T/EBC system.

Fig. 10. Example of estimated stress state in a five-layer coatingsystem with air plasma-sprayed and EB-PVD zirconia-based toplayer.

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CMC combustor liner is being currently planned at GEAircraft Engines. The coating for this test was validatedin the lab tests, described in the previous sections of thepaper, and has undergone relatively short high-tempera-ture, high-pressure Combustor Rig test. A section of thecoated outer liner in the Combustor Rig, after initial hoursof rig testing, is shown in Fig. 12. The upcoming enginetest will provide a better assessment of the Gen 1 coatingcapability for the jet engine applications as well as of someother aspects of the coating technology.

SSSSSummarummarummarummarummaryyyyy

Significant progress has been made in developingenvironmental and thermal barrier coatings for SiC/SiCCMCs over the last 10 years. The development effortshave resulted in three generations of the coatings, withtheir temperature capability increasing up to 1700ºC(3100ºF). EBCs for the lower-temperature applications(Gen 1) have been optimized, scaled-up, and demonstratedat the component level. Applications of the Gen 1 coat-ings are limited to temperatures below 1371ºC (2500ºF)due to the recession and material temperature stabilityissues. Efforts toward developing higher-temperature-ca-pability coating systems (Gen 2 EBCs and Gen 3 T/EBCs)are continuing. The feasibility of utilizing these coatingsat temperatures up to 1704ºC (3100ºF) has been dem-

onstrated in the laboratory rig tests, and follow-on en-gine tests are planned at GE aircraft Engines. Success inmaturation of these coatings will depend on improvingthe test methodology as well as achieving the higher tem-perature test capability. Establishing an understanding ofthe coating failure mechanisms and incorporating theminto the coating life models is also critical. In addition,identifying new materials with the desired characteristicswould enable a reduction in the number of layers for theT/EBC system and make these coatings more practical interms of both manufacturing and prediction of performance.

AAAAAcknocknocknocknocknowledgmentswledgmentswledgmentswledgmentswledgments

The authors would like to acknowledge Mark Noe,Brian Hazel, Christine Govern, all of GE Aircraft En-gines, and Y.-C. Lau of GE GRC for their contributionsto the T/EBC development effort at GE Aircraft Engines;Karren More of Oak Ridge National Laboratory for col-laborative work on characterization of the coatings;Dongming Zhu, Bob Draper, Dave Brewer of NASAGlenn Research Center, and Inna Talmy and JimZaykovski of Naval Warfare Research Center for theircollaborative efforts; Lt. Jeremey Schroeder and CharlesStevens of Air Force Research Lab as well as Steve Fishmanand Gill London of Office of Naval Research for supportof T/EBC efforts at GE Aircraft Engines.

Fig. 11. Five-layer coating after 50 1-h cycles in the LaserGradient Rig: coating sample view (top) and coating cross-section (bottom).

Fig. 12. Section of commercial combustor outer liner in theHigh-Pressure High-Temperature Combustor Rig after initialhours of testing.

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RRRRReferefereferefereferencesencesencesencesences

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2. H. Eaton et al, “Article having silicon-containing substrate and barrier layerand production thereof ”, US 6,387,456 (2001).

3. K. More, P. Tortorelli, M. Ferber, J. Keiser, “Observations of Accelerated SiliconCarbide Recession by Oxidation at High Water-Vapor Pressures”, J. Amer.Ceram. Soc, 83 [1], 211-13 (2000).

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under Combustion Conditions:1, Experimental Results and Empirical Model”,J. Amer. Ceram. Soc, 82 [7] 1817-25 (1999).

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11. K. More, P. Tortorelli, L. Walker, J. Kimmel, N. Miriyala, J. Price, H. Eaton,E. Sun, G. Linsey, “Evaluation of Environmental Barrier Coatings on CeramicMatrix Composite Combustor Liners after Engine and Laboratory Exposures”,presented at the International Gas Turbine&Aeroengine Congress & Exhibition,Amsterdam – June 3-6, 2002, ASME 2002-GT.

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Y2O

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18. R. Bruce, “Development of 1232ºC (2250ºF), “Erosion and Impact Tests forThermal Barrier Coatings”, Tribology Transactions, 41 [4], 399-410 (1998).

19. A. Karlsson, unpublished work.

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Documents

Thermal and Environmental Barrier Coatings for SiC/SiC CMCs in Aircraft Engine Applications