Protective Coatings for Turbine Blades

Protective Coatings

for Turbine Blades

Y. Tamarin

Materials Park, Ohio 440730002www.asminternational.org

2002 ASM International. All Rights Reserved.Protective Coatings for Turbine Blades (#06738G)

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Copyright 2002by

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First printing, September 2002

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Comments, criticisms, and suggestions are invited, and should be forwarded to ASM International.Prepared under the direction of the ASM International Technical Book Committee (20012002), Charles A.Parker, ChairASM International staff who worked on this project included Steve Lampman, Manager of Book Acquisitons;Bonnie Sanders, Manager of Production; Carol Terman, Production Project Manager; and Scott Henry,Assistant Director of Reference Publications.

Library of Congress Cataloging-in-Publication Data

Tamarin, Y.Protective coatings for turbine blades / Y. Tamarin.

p. cm.Includes bibliographical references and index.

1. Aircraft gas-turbinesBlades. 2. Protective coatings. 3.Gas-turbinesMaterials. I. ASM International. II. Title.

TL709.5.B6 T36 2002629.134353dc21

2002027690

ISBN: 0871707594SAN: 2047586

ASM InternationalMaterials Park, OH 440730002


Printed in the United States of America

Cover: Test blades (shiny) in a low-pressure turbine. Source: Advances in Turbine Materials, Design, and Manufacturing,Proceedings of the Fourth International Charles Parsons Turbine Conference, The institute of Materials, 1997



Contents

Foreword.......................................................................................................................v

Preface......................................................................................................................... vi

Chapter 1: Introduction....................................................................................................1

Chapter 2: Choosing Optimum Coatings for Modern Aircraft Engine Turbine Blades.................5Conditions of Turbine Blade Operation ...............................................................5Requirements Imposed on Turbine Blade Coatings................................................8Principles of Choosing Coatings for Aircraft Engine Turbine Blades ........................8Causes of Coating Failures on Aircraft-Engine Turbine Blades.............................. 10

Chapter 3: Technological Processes for Deposition of Protective Coatings to Turbine Blades .... 25Diffusion Coatings ........................................................................................ 25Overlay Coatings .......................................................................................... 38

Chapter 4: Phase Composition and Structure of Coatings on Superalloys ............................... 55Phase Composition and Structure of Diffusion Coatings....................................... 55Phase Composition and Structure of Ni-Cr-Al, Ni-Co-Cr-Al, and Co-Ni-Cr-AlAlloys and Overlay Coatings Made of Them...................................................... 69Phase Composition and Structure of Overlay Coatings......................................... 71

Chapter 5: Phase and Structural Changes in Coatings during High-Temperature Tests .............. 79Changes of Phase Composition and Structure in Diffusion Coatings....................... 80Changes of Phase Composition and Structure in Overlay Coatings at High-Temperature Tests ......................................................................................... 87

Chapter 6: Turbine Blade Coating Protective Properties ...................................................... 97Protective-Properties Evaluation Methods .......................................................... 97Heat Resistance of Aluminides and Alloys for Overlay Coatings ..........................100Heat Resistance of Coated Superalloys.............................................................106Resistance of Overlay-Coated Alloys to Hot Corrosion .......................................109Resistance of Coated Superalloys to Hot Corrosion ............................................113

Chapter 7: The Effect of Protective Coatings on the Mechanical Properties of Superalloys .......119Thermal Expansion Coefficients and Elasticity Modulus of Coating Alloys ............120Mechanical Properties of Coating Alloys ..........................................................124Thermal Stresses in Superalloy Coatings ..........................................................128Effect of Coatings on High-Temperature Strength of Superalloys..........................134

iii



iv / Protective Coatings for Turbine Blades

Effect of Coatings on Fatigue Strength of Superalloys ........................................138Effect of Coatings on Thermal Fatigue of Superalloys ........................................142Coating Effect on Thermomechanical Fatigue of Superalloys ...............................150

Chapter 8: Electron Beam Thermal Barrier Coatings..........................................................161Ceramics for EB Evaporation .........................................................................162Main Features of TBC Deposition Technique....................................................165Thermophysical Properties of Condensed Ceramics ............................................175TBC Ceramic Layer Durability .......................................................................180

Chapter 9: Some Principles of Strength Designing for Turbine-Blade Protective Coatings ........195Calculation of Stresses and Strains in Coatings..................................................195Thermal Barrier Coatings ..............................................................................203

Appendix....................................................................................................................211

Index .........................................................................................................................217

iv



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CHAPTER 1

Introduction

THE DAMAGE TO blade surfaces caused byoxidation and hot corrosion results in consider-able deterioration of the mechanical propertiesof blades and shortens their service lives. Thatis why protection of turbine blade surfacesagainst damage has been a common problem.Solving this problem is critical for designingcost-effective and reliable aircraft engines andstationary gas turbine units. The idea to apply alayer with protective properties to the surface ofa nickel superalloy was first practiced in the1960s and found a wide application area. Sincethen, aircraft engine turbine blades with protec-tive coatings have been used.

In the 1960s, extensive research into the prop-erties of various coatings demonstrated that dif-fusion aluminide coatings had the best protec-tive properties on turbine blades made of nickelsuperalloys. Owing to their properties, such ashigh oxidation resistance, high stability whendeposited on nickel superalloys, and a favorablecombination of physical and mechanical prop-erties, the diffusion aluminide coatings havebeen the predominant type of aircraft engine tur-bine blade coatings used for several decades.The simplicity of the technique used in formingdiffusion coatings on the blade surfaces in thepowder mixtures containing free or bonded alu-minum and aluminum halides contributed totheir most common use. Translation of diffusionaluminide coatings in a commercial practice andthe work on improvement of their protectiveproperties and further development of their de-position techniques started concurrently.

This further development was the result of theneed for improved properties of protective coat-ings in the media containing sulfur compounds.Diffusion aluminide coatings with the increasedchromium content in their outer layer (chrom-izing-aluminizing method) had found applica-tion in the protection of aircraft engine blades

suffering from hot corrosion. Alloying alumin-ide coatings with silicon (silicification in com-bination with aluminizing) turned out to be onemore technique that allowed the improvement ofcoating resistance to hot corrosion and oxida-tion. Since the 1970s, both modifications of thecoatings have been used for protection of aircraftengine turbine blades.

The studies aimed at modifying diffusion al-uminide coatings are ongoing. Some of the coat-ings designed (e.g., Pt-Al) are used widely inaircraft engines and stationary gas turbine units.

Designing new-generation engines with in-creased inlet gas temperatures resulted in theshortening of the surface lives of the blades pro-tected with diffusion coatings. Low ductility ofdiffusion coatings and high stresses generated incooled blades cause the formation of numerousthermal fatigue cracks in the outer layer of dif-fusion coatings on the blade surfaces. In updat-ing diffusion coatings, designers seek to im-prove primarily the resistance of the coatings tothermal fatigue. The problem of improving thecoating property in question arises from the lim-ited technological feasibility of changing itscomposition, structure, and consequently, itsphysical and mechanical properties.

The studies aimed at increasing ductility ofdiffusion aluminide coatings resulted in thegradual decrease of the aluminum contents ofthe coatings from 34 to 36%, typical of the coat-ings used in the 1960s, down to 20 to 24%, typ-ical of the coatings used since the 1970s. Thistrend, however, resulted in a decrease in oxida-tion resistance.

High labor consumption and expenditure ofenergy on the formation of diffusion aluminidecoatings on turbine blades in powder mixturesand the need for the use of halides in the tech-nological process stimulate the seeking of alter-native techniques for forming diffusion coatings

Protective Coatings for Turbine Blades Y. Tamarin, p1-3 DOI: 10.1361/pctb2002p001

Copyright 2002 ASM International All rights reserved. www.asminternational.org

2 / Protective Coatings for Turbine Blades

on turbine blades. Coating deposition fromslurry allowed considerable improvement of theprocess conditions under which the coatingswere formed on the turbine blades. The methodhas experienced a wide application area in air-craft engine manufacturing. Further techniqueimprovements resulted in the development ofvarious gas aluminizing methods. These meth-ods allowed less employment of powdermixtures and higher efficiency of gas transfer ofaluminum to blade surfaces. By now, the tech-niques of blade aluminizing that eliminate com-pletely the need for covering the blades withpowder mixtures have been developed and areenjoying application. In this case, aluminumtransfer to the blade surface occurs by circula-tion of gaseous aluminum halides formed in aspecial reaction zone of the unit.

A further advancement in the practice of dif-fusion coating deposition was the use of electric-arc evaporation of aluminum alloys and theirtransfer to the blade surface in the form ofplasma. In this case, the need for the use of ha-lides and powder components is eliminatedcompletely, and it opens up the possibility forthe accurate control of diffusion coating thick-ness and its location on the blade surface. Thistechnique is used for forming the diffusion coat-ings modified with silicon, yttrium, and otherelements on turbine blades.

At the same time, it is noteworthy that alltechnological processes of diffusion coating de-position used in turbine blades production formthe coatings with similar protective properties.As a rule, the choice of a certain technique forcoating application depends on requirements forits quality, technological equipment available,and personnel skills.

Although diffusion coatings have been enjoy-ing wide application, they do have a limitation:They do not allow deliberate changes of theirproperties to meet the requirements imposed byspecific blade service conditions. This restrictionexists because all diffusion coatings always arebased on Ni-Al and Ni3Al aluminides with typ-ical combinations of physical-chemical and me-chanical properties. This drawback follows fromthe principle of diffusion coating formation dueto interaction of aluminum transferred to the sur-face from an aluminizing medium and the com-ponents of a nickel superalloy. It cannot be over-come by modernization of coating compositionsand techniques of their deposition. That is why,as a following step in developing protectivecoatings for turbine blades, it was quite natural

to apply some new principles of their formation,such as deposition of the alloys, which compo-sitions met complex requirements to protectivecoatings for turbine blades most adequately.

Design of electron-beam, electric-arc, plasmamethods, and equipment for their realizationtriggered the change to the new principles of de-velopment and application of protective coatingsto turbine blades. Since the 1980s, some variantsof overlay coatings have found application forthe protection of turbine blades. Unlike diffusioncoatings, overlay coatings admit the control oftheir composition, thickness, and combination ofphysical-chemical and mechanical propertieswithin wide ranges. It allows the choice of over-lay coatings, which have the properties optimalfor certain engine types and their service con-ditions.

New principles of coating deposition openedup new possibilities for purposive improvementof coating compositions and variation of theirproperties. However, a wide range of possiblevariants of coatings compositions and tech-niques for their deposition presents the problemof making an adequate choice in coatings forcertain engine types.

For this reason, development of calculationmethods for determining stresses and strains inthe coatings on turbine blades and thus their ser-vice lives under conditions of high thermalstresses further advanced the work on overlaycoatings. To calculate the values in question, itis necessary to have certain information onchemical and mechanical properties of the coat-ings within a wide range of probable composi-tions.

Basically, a new stage of progress in the areaof turbine blades protection has been achievedwith designing thermal barrier coatings. Theouter layer is formed of stabilized zirconium ox-ide-base ceramics. This coating type allows notonly protection of a coating surface against cor-rosion damage as all previously designed coat-ings did, but it was also the first to allow pro-tection of the structure of a superalloy in cooledblades against changes caused by exposure tohigh temperatures and, hence, against its soft-ening.

An apparently simple idea of designing a ce-ramic layer on the blade surface called for ex-tensive research and development for its reali-zation that involved designing techniques forproduction of special ceramics, their evaporationwith electron beam, special test methods, andcalculations of service lives for thermal barriercoatings on turbine blades.

Introduction / 3

As is well known, metal structural materialsproperties depend on their chemical composi-tions and structures. Control of metal materialscompositions and structures makes it possible toproduce a great variety of structural alloys withrequired properties. Similarly, protective coat-ings properties depend on their chemical com-positions and structures. And yet, the ap-proaches worked out for structural materialshardly remain valid for coatings because the re-lationships connecting chemical compositions,structures, and properties in thin coating layersdo not reveal themselves clearly. Besides that, acoating-superalloy system is a dynamic one.The chemical and phase compositions of a coat-ing, as well as its structure, vary continuouslyunder blade service conditions at elevated tem-peratures. The initial coating properties on theblade and those it has at any moment of itsoperation in the turbine may be quite different.

Since the first use of protective coatings onthe turbine blades of different designations, hun-dreds of papers that deal with their propertieshave been published. However, to create a da-tabase covering the papers published and storingthe information required for the proper choiceof coatings and for calculations of the servicelives of coated blades, one should overcome theproblems of coatings identification and lack ofdata on chemical and phase compositions andstructures of the coatings studied. No unifiedstandard methods for research into coatingsproperties are available. In most cases it makesimpossible comparison and classification of thetest results obtained.

Special databases should be created to solvethe problems of description and systematizationof the steadily increasing number of protectivecoatings for turbines. The databases should pro-vide systematized, accumulated information onthe coatings properties studied and on the ex-perience of their use in practice. Systematizationof the results presented by laboratories andfirms, which deal with the problems of turbineblades protection in the form of databases, may

allow further simulation of changes in coatingsproperties, as well as formalization of compo-sition-structure-property relationships. Besidesthat, the characteristics required for coated bladeservice life calculations can be derived.

This book presents knowledge accumulatedby the author over many years of his work onthe research and development of diffusion andoverlay protective coatings for aircraft engineturbine blades. The book includes detailed de-scriptions of some technological processes fordiffusion and overlay coatings deposition, to-gether with compositions of the coatings usedfor turbine blades protection, coatings structuresafter their deposition on the blade surfaces, andtheir changes during high-temperature testing.The results of a number of tests for coatings andsamples of the alloys used for coating depositionare discussed also. Some chapters deal with ther-mal barrier coatings, technological features oftheir deposition, and test results, as well as withthe principles of calculations of stresses andstrains in the coatings on blades.

The composition-property regression equa-tions derived on the basis of research into alu-minide alloys of Ni-Cr-Al, Ni-Co-Cr-Al, andCo-Ni-Cr-Al systems used for deposition ofoverlay coatings are given in the Appendix.They allow simulation of aluminide alloys prop-erties. These simulations cover the compositionsof both as-deposited protective coatings andthose after their long-term operation on turbineblades. In some cases the simulations obtainedmay be used for assessment and prediction ofcoatings properties. Many coatings dealt with inthis book are used currently for turbine bladesprotection.

The information on technological processesand coatings properties presented in the bookmay be useful to the specialists who work in thefield of turbine blades protection, to the scien-tists interested in research into materials behav-ior at hot corrosion, and to the students whosespecialty is protection against hot corrosion.

CHAPTER 2

Choosing Optimum Coatings forModern Aircraft Engine TurbineBlades

COATED TURBINE BLADES boast servicelives of 2 to 5 times those of their uncoatedcounterparts. However, coating protection ef-fects are manifested only if the coating has beenchosen correctly, both in terms of blade designand engine running conditions.

Conditions of Turbine Blade Operation

The main parameters determining the oper-ating conditions of turbine blade surfaces of air-craft gas-turbine engines are the gas temperatureat the turbine inlet; the pressure, speed, and com-position of the gas flow; the stresses due to theeffect of the centrifugal and gas dynamic forces;and the irregularity of the temperature field.

The gas temperature at a turbine entry is themost important parameter determining the spe-cific power, specific weight, and efficiency of anaircraft engine. During the last 20 years, themean mass temperature of the gas at a turbineentry increased from 1200 to 1300 K to 1700 to1800 K. Together with the turbine temperaturegrowth, the pressure ratio of the compressor hasgrown significantly as well; it has reached 20 to30. The combined growth of the temperature andthe pressure ratio results in significantly moreintensive heat flows in turbines; they make 1 to2 MW/m2 for modern engines.

One of the main properties of aircraft gas-tur-bine engines is a significant number of variableconditions when the level of temperatures andstresses differs significantly from those mea-sured in stationary conditions. Statistical analy-sis shows that for aircraft engines, about 20% of

the flaws developed at operation are flaws re-sulting from the cyclic temperature variations(Ref 1). Figure 2.1 shows the operating diagramof an airliner for a typical medium-range flight.

Every 1.5 h flight has four corresponding cy-cles of condition variation; that is, 10,000 h ofengine life correspond to 27,000 cycles ofvariation in flight conditions. Every time an en-gine reaches takeoff power, the nonuniformityof the temperature field in the turbine blades in-creases significantly. Running the engine undercruising conditions does not eliminate the effectof nonuniform temperature and stress fields onthe blades. The number of varied powers forwarplane engines, especially fighters, is signifi-cantly greater than that for passenger aircraft en-gines.

The surface layers of aircraft-engine turbineblades during operation contact the combustionproducts of the aviation fuel. This is hydrocar-bon-based (98 to 99%), with the remaining 1 to2% including sulfur, nitrogen, oxygen com-pounds, and trace metals as well as their com-pounds. A widespread impurity in fuels is va-nadium oxide (V2O5), whose content depends onthe deposit of the source oil. The fuel sulfur hasthe form of its various compounds, mercaptanes,sulfides, disulfides, and thiophenes. The totalsulfur content of an aviation fuel is in the rangeof 0.05 to 0.25%.

Aviation fuels usually get impurities from theenvironment and from the production process,resulting from corrosion and rubber decompo-sition and so forth. Typical impurities of jet fuelsare as follows: C, H, S, N, Fe, Si, Sn, Ca, Mg,Pb, Na, Ba, Ti, Ni, Al, Cu, Zn, Mn, and O.

Protective Coatings for Turbine Blades Y. Tamarin, p5-23 DOI: 10.1361/pctb2002p005

Copyright 2002 ASM International All rights reserved. www.asminternational.org


Fig. 2.1 Operating diagram of an airliner engine. Source: Ref 1

The quantity and the composition of the im-purities depend on local operation conditions, onthe fulfillment of requirements to fuel controlsand purification, and on aircraft fueling condi-tions (Ref 2 to 4).

In a combustion chamber, the process of fuelcombustion is realized at the temperature of2273 to 2473 K. Hydrocarbon combustion is aprocess of their oxidation by the oxygen in air,resulting in carbon dioxide and water, whichmay contain some carbon monoxide (CO), H2,methane (CH4), and hard carbon particles. Thecomplete combustion of 1 kg of fuel requires15 kg of air. The total factor of air excess incombustion chambers compared to the amountnecessary for complete fuel combustion is 3.5 to4.5; it may change in dependence of the flightconditions. The specific fuel consumption formodern engines provides 0.75 to 0.90 kg/kgf ofthrust.

The high-temperature gas flow at the turbineinlet contains mainly oxygen, carbon dioxide,water vapors, and some amounts of CO, H2,CH4, hard carbon particles, compounds of sul-fur, and metals present in fuels. Various chem-ical compounds may come into the gas flowfrom the environment. These are sea salts fromflight over sea regions and various compoundsthat are exhausts of industrial enterprise. As arule, they contain sulfur, salts of alkaline metals,

and other compounds, depending on the industrysituated in the flight area.

The temperature pattern of the gas flow at theturbine inlet is highly inhomogeneous; the in-homogeneity is usually caused by the design ofa combustion chamber, by the design and layoutof burners, and by other factors determining theflow-gas dynamics and the perfection of the fuelcombustion process. The measurements of thegas temperature before the nozzle guide vanesshow that a difference between the extreme gastemperatures at a nozzle block inlet may be 400to 500 C.

The nozzle vanes of a high-pressure (HP) tur-bine situated just behind the combustion cham-ber are the most heated elements of an engine.The temperature of the cooled nozzle vanes is200 to 300 C lower than that of the gas flow,but the thermal inhomogeneity of the gas at aturbine inlet results in some vanes working attemperatures differing significantly from the av-erage ones. The temperature measurement ofvanes has shown, that the thermal pattern of thevanes leading edges in the middle section differsby 200 to 250 C. The thermal inhomogeneityof the nozzle vanes results in significantly dif-fering rates of flaw development in vanes situ-ated in different parts of an engine relative to thecombustion chamber baffles and burners.

The gas temperature pattern inhomogeneity isobserved not only by the hot-duct circumfer-

Choosing Optimum Coatings for Modern Aircraft Engine Turbine Blades / 7

ence, but by height, too. The characteristic tem-perature curve by the vane height shows an areaof maximal temperature in the vane mid-part andlower temperatures at end areas. Nevertheless,other temperature patterns may also be ob-served. The greatest difference between the mid-dle and peripheral areas of the vanes is 100 to150 C. The temperature gradient value growsat the transition to the gas high-temperature con-ditions.

The vane and blade thermal state depends onthe engine power; it changes from 450 to 500 Cto 1000 to 1100 C within a few seconds at theengine transition from idle power to take-offconditions. The temperature inhomogeneity ofthe vane and blade surfaces increases signifi-cantly under the transient conditions of startup,acceleration, and run. The maximal thermalstresses of the vanes, especially at their surfaces,within the coating areas are observed at thesevery conditions. The computed data that char-acterize variations in cooled-vane temperaturesat the engine startup show that the difference inthe temperatures between the vane back andtrailing edges may reach 500 to 600 C. Thecomputation results on the thermal stresses cor-responding to the above temperature patternsshow that their values attain 400 to 500 MPa(Ref 5).

The analysis of the strains in a vane causedby such thermal stresses reveals that some areasof a vane are in a state of plastic flow. The levelof thermal stresses and strains in a vane surfacelayer may exceed significantly the mean level ofthe vane stresses due to the presence of the tem-perature gradients along the cooled-vane wallsection and due to different physical and me-chanical properties of the protective coating andsuperalloy. The accumulation of thermal varia-tions at operation results in an accumulation ofresidual strains, which, in their turn, result inthermal fatigue cracks on the vane surfaces, firston the leading and trailing edges. These cracksare the most widespread flaws of vanes.

Analysis of the vane operating conditionsshows that the combined effect of high tempera-tures, gas dynamic forces, and thermal stressesin vanes causes cracks due to thermal fatigue,burnouts, and warping. The higher temperatureof the engine gas resulted in higher temperaturegradients in the vane surface layers, and the useof cast nozzle blocks caused higher stiffness ofrestraint of individual vanes. As a result, thethermal fatigue cracks have become the mostwidespread flaw of the vanes during tests and

operation, despite the use of efficient film meth-ods for cooling.

One of the critical elements of an aircraft gas-turbine engine is its turbine blade. It is its ther-mal state that determines the maximal admissi-ble gas temperature in the turbine and thereliability and life of the turbine and the enginein whole. The blades are affected by centrifugaland gas dynamic forces causing tension, bend,and torsion of the blades. The high tempera-tures and stresses, the unstable conditions ofheating and loading, and the possibility of res-onance vibrations make a blade one of the mostcomplicated elements of an engine.

In contrast to the nozzle vanes, the gas tem-perature field before the turbine blades is moreuniform. The high rotation speed of a blade re-sults in a more homogeneous thermal state ofindividual blades, if compared with the vanes.The temperature gradient by the blade surfacedepends on the gas-temperature curve by heightand by the specific operation of the blade cool-ing system. The highest temperature area is, asa rule, in the upper third of a blade airfoil por-tion; this is explained by the need to place itwithin the area with the lowest centrifugal loads.The temperature distribution by the turbine sec-ond stage blades is even more homogeneous,though the non-cooled, second-stage blades ofsome engines have higher average temperaturesthan the ones of the first stage.

The limitations due to high-temperaturestrength (heat resistance) of the nickel-based su-peralloys and the necessity to improve the pa-rameters of modern aircraft engines pushed thedesign of sophisticated and efficient cooling sys-tems; they permitted lower average temperaturesof a blade wall by 300 to 400 C, compared tothe gas temperature. The higher temperature andcompression rates in modern engines result ingreater temperature irregularities in the blades.When the gas temperature increases by 100 C,the temperature gradient on the surface of acooled blade increases by 40 to 50 C; the com-pression rate effect is similar. The temperaturepattern of the modern cooled blades is complexenough: the areas of high temperature, 1000 to1100 C, whose size is often 10 to 30 mm2, arealternated by the average (900 to 950 C) andlow (750 to 800 C) temperature areas. The tem-perature gradient by the surface and wall sectionof the blades causes high thermal stresses, reach-ing maximal magnitudes at engine operation un-der non-cruise conditions (Ref 6 to 8).

A blade is affected by tensile stresses causedby centrifugal forces. Their levels are not usu-


ally higher than 200 to 250 MPa. A blade is alsoaffected by the bending stresses caused by gasdynamic forces. These stresses are determinedby the difference in pressures at the concave(pressure) and convex (back) surfaces of a blade.The combined static stresses caused by the ex-ternal forces affecting turbine blades are distrib-uted irregularly along a blade. The highest ten-sions are observed near the blade root, the lowestones exist at the end of a blade airfoil portion.The section with the least static strength securitymargin or with the maximal exposure to flawsdue to the long-term static load is usually be-tween the maximal stress and maximal tempera-ture sections (Ref 9 to 11).

The vibration loads are also critical for theblades; these loads result from the inhomoge-neity of the gas flow hitting the turbine wheelrim. The low-frequency vibrations are caused bythe flow thermal inhomogeneity due to the pres-ence of burners and pillars; the high-frequencyones result from the finite number of nozzlevanes. The resonance vibrations are the mostwidespread cause of the surface fatigue cracksand rupture of the turbine blades. The level ofvariable stresses in the blades reaches 80 to 100MPa (Ref 12).

Requirements Imposed on TurbineBlade Coatings

Any coating deposited on turbine blades (orvanes) must offer protection within a specifiedperiod of service life against destructive attacksof high-temperature corrosion and erosion,when the said components are exposed to a flowof fuel combustion products containing aggres-sive ingredients and solid particulate matter.

With this in mind, the coating must meet thefollowing requirements:

It must withstand hot corrosion, oxidation,and erosion when placed into a flow of gaswhose parameters are similar to those of tur-bine gases.

It must safely withstand the static and alter-nate stresses applied to the blade surface; tothis end the coating must have the requisitecombination of strength and ductility.

It must show good stability and not be de-stroyed by interaction with the substrate.

It must not degrade the blade material me-chanical properties.

Even after a long service life, coated bladesand vanes must have better mechanical proper-ties than their uncoated counterparts due to pro-tecting the blade surfaces from damage and soft-ening.

In addition to the requirements to metal coat-ings listed previously, the ceramic coatings(thermal barrier coatings, or TBCs) must ensurelower average wall temperatures in cooledblades and protect superalloys against softening.TBCs must level off the temperature over theblade surface and reduce thermal stresses duringengine transient running.

The techniques involved in depositing protec-tive coatings must guarantee that the coatingshave the required composition, thickness, andstructure, and, as such, the requisite set of physi-cal-chemical and physical-mechanical charac-teristics. Of utmost importance is the need tomaintain these characteristics depositing protec-tive coatings on blades under the commercialconditions.

Three technological principles are applied toform coatings on the surfaces of aircraft engineturbine blades: Chemical-thermal treatment (aluminizing) of

blade surfaces in the media, containing alu-minum or aluminum with additives of otherelements (such as chromium, silicon, yttrium,etc.). A diffusion aluminide coating is builtup as aluminum or its compounds interactwith the superalloy surfaces. At present, dif-fusion aluminide coatings are used to protect80 to 90% of all aircraft turbine blades.

Depositing the overlay coatings by evaporat-ing special alloys under vacuum and con-densing vapor or plasma with coating com-ponents on blade surfaces

Forming coatings from powders by arcplasma spraying

Principles of Choosing Coatings forAircraft Engine Turbine Blades

The requirement to attain highly protectivecoatings (i.e., adequate resistance to hot corro-sion and oxidation) conflicts with the demand topreserve high mechanical properties in thecoating-superalloy system.

Until recently, engine tests have remained themain method of examining the properties of andchoosing optimized coatings for blades. How-


Fig. 2.2 Scheme of the choice of a coating for aircraft engineturbine blades

ever, the high cost of systematic studies of coat-ing protectiveness using this approach has fu-eled the desire to seek shorter equivalent trials.Under these conditions, laboratory research intophysical-chemical and physical-mechanicalcoating properties and their relation to blade ser-vice life facilitates significantly the selection ofoptimal chemical compositions of the coatings,reduces the number of their field tests, andpromises high cost efficiency of the study.

Shown in Fig. 2.2 are the key parameters toconsider in selection of coatings for turbineblades. The starting points to be taken into ac-count when choosing a coating of optimum lifeexpectancy include: Experience in applying protective coatings on

blades running under similar conditions Analysis of causes of coating damage on

blades after long-time tests or service opera-tion

Presumable alterations in running conditionsof newly designed blades (changed parame-ters of gas flow, values of thermal stresses,running conditions, etc.)

Results of research into physical-mechanicalproperties of coatings and their effects on themain properties of the superalloy from whichthe blades are made (resistance to thermal fa-tigue, endurance, high-temperature strength)Of prime significance for the protectiveness

of a coating is its resistance to hot corrosion andoxidation in the temperature range of 600 to

1200 C, that is typical of the running conditionsof modern-engine turbine blades. The mecha-nism of coating damage in this range can varydepending on the temperature and compositionof gas in turbine, blade surface temperature, anddesign. Coatings are chosen from the laboratoryresults of coating tests for heat resistance andhot corrosion.

Actual engine running conditions can resultin blade surfaces being exposed to temperatureseither below or higher than 900 C. Moreover,the demands on coating life expectancy at tem-peratures above 900 C may conflict with lifeexpectancy requirements imposed on bladesworking in corrosive media at temperatures be-low 900 C. When choosing coatings, due atten-tion should be paid to possible application ofengines in coastal regions or in areas with airheavily laden with industrial impurities. Shouldthe blade damage be mostly caused by hot cor-rosion, then coatings with high resistance to thistype of corrosion should be preferred. Alterna-tively, multilayer coatings can be used, giventhat individual layers of such coatings offer highresistance to oxidation and hot corrosion.

Coating thickness is undoubtedly a very im-portant characteristic of its protectiveness; coat-ing resistance to hot corrosion and oxidation isdirectly proportional to the thickness of coatingused. It should be borne in mind, however, thatthe thickness of coatings applied to turbineblades are restricted, by both the probable de-crease in mechanical properties of the bladeswith thick coatings and by technological prob-lems involved in depositing thick coatings. De-sign problems must also be taken into account(e.g., the reduction of the open flow area in theturbine).

As overlay coatings add to the mass of bladesthey are applied to, hence, the centrifugal loadson the blades and turbine disk grow. This mustbe counted when thick coatings are used. Vari-able thickness coatings are most advantageous;a thicker layer is deposited on the area that ismost likely to be attacked by corrosion. Otherareas are thinner-coated.

No correlation between coating thickness andthermal fatigue resistance has been noted inoverlay coatings up to a thickness of 120 lm. Apositive correlation seems more likely in thickercoatings prone to formation of various defects,which, in turn, are conducive to the emergenceof thermal fatigue cracks. When using coatingsthicker than 120 lm, the blade fatigue limit mustbe taken into consideration.


Table 2.1 Alloy composition for overlay coat-ings deposited on turbine blades

Chemical composition, wt%Alloy Ni Co Fe Cr Al YCo20Cr12AlY Base 1822 1113 0.20.6Co23Cr12AlY Base 2224 1113 0.20.6Co26Cr9AlY Base 2528 810 0.30.6Co25Cr4AlY Base 2426 3.55 0.30.6Co30Cr6AlY 02 Base 2832 57 0.30.6Co28Cr10FeY Base 812 2630 0.1Co22Ni23Cr12AlY 1825 Base 2224 1113 0.30.5Co8Ni23Cr12AlY 610 Base 2224 1113 0.20.6Ni20Co20Cr12AlY(a) Base 1822 1822 1113 0.20.6Ni8Co20Cr12AlY(a) Base 610 1822 1113 0.20.6Ni20Cr12AlY(a) Base 1822 1113.5 0.20.6Ni20Cr5AlY Base 1822 46 0.20.6Ni36Cr5AlY Base 3538 46 0.20.5Fe25Cr5AlY Base 2326 46 0.20.5Fe22Ni24Cr5AlY 2025 Base 2325 46 0.20.5

(a) Composition used in aviation. Source: Ref 13, 14

Coating composition and thickness can be de-termined once the required service life has beenspecified. Once the coating composition hasbeen determined, one must then consider howlong the coating of given composition will staysteadily on the blade surface after long exposure.Coating protectiveness is limited by diffusioninto the alloy of those elements that control re-sistance to oxidation and hot corrosion. The dif-fusion progress can cause the decrease in alu-minum, chromium, and cobalt contents of thecoating. It can call for the use of special methodsfor retarding diffusion processes by introducingbarrier layers based on phases that are stablewith respect to both the coating and the alloy.On the other hand, as coating elements diffuseinto the alloy, a zone of poorer heat resistanceis formed under the coating. It is a factor that,in thin-walled blades, leads to reduction of themargin of safety. No coating lacking the requi-site stability can be used for turbine-blade pro-tection.

The life of a coated turbine blade depends onits mechanical characteristics and the stresses onthe surface. Thermal fatigue cracks have becomethe commonest cause of service life reductionand discarding of blades. Calculation of thestresses and strains in turbine blade constructionis mandatory in coating selection for modern tur-bine engines.

Engine testing is the final step in coating se-lection and estimating useful life on turbineblades. Defects in a coating on blade surfacesmay be caused by hot corrosion and oxidation,low resistance of blade construction to thermal

fatigue, poor high-temperature strength and en-durance of the substrate, or by a combination ofthese factors. In order to assess life expectancyof coated blades in long-term tests, their tem-perature schedule must be as close to the actualengine running conditions as possible. The mat-ter is that in the course of engine-equivalentquick tests, where the turbine blades are runningat higher than actual service temperatures, theprocesses occurring on the blade surfaces, aswell as between the coating and superalloy, maysignificantly differ from those occurring at lowertemperatures. Elevation of temperatures duringequivalent quick tests may be misleading due toreduced condensation of aggressive componentsfrom the gas flow on the blade surfaces and sug-gest longer-than-actual service lives. Alterna-tively, such tests may degrade coating resistanceowing to sharp acceleration of diffusion pro-cesses and rapid destruction of protective ox-ides.

Analyzing the results of engine tests pinpointsthe coating-failure mechanism and enables, ifnecessary, appropriate modifications to the com-position, thickness, and/or design of multilayercoatings.

Chemical compositions of some alloys usedfor coating deposition on turbine blades are pre-sented in Table 2.1 (Ref 13, 14). It is noteworthythat the coating composition depends on the cur-rently used technological process of its deposi-tion. It can differ from that of the initial alloy.In this book, the coatings are marked with thegrades of the alloys used for their deposition.Otherwise, when the coatings are marked ac-cording to their chemical compositions, the re-spective notes are supplied.

Chemical compositions of the materials usedfor applying diffusion coating to aircraft turbineblades are presented in Table 2.2 (Ref 15 to 21).Chemical compositions of the superalloys usedfor research into coating properties are presentedin Table 2.3 (Ref 22, 23).

Causes of Coating Failures on Aircraft-Engine Turbine Blades

The investigation of coating damage, detectedon the blades of engines after a long time inservice, is an important stage in the choice of anoptimum coating. Careful investigation of thenature of damage and the sources of origin willenable the designer to choose a coating that has


Table 2.2 Chemical compositions of the materials used for forming diffusion coatings

Method of coating Grade Chemical composition of alloy (mixture), % massPack calorizing(a) Al Granules of alloy of Fe(3550%)Al; 2% NH4ClPack calorizing(a) Al Granules of alloy of Fe(7075%)Al; 2% NH4ClPack calorizing(a) Al Powder 98%Al, 2% NH4ClPack calorizing under vacuum(b) AlCr Powder (78%Al; 3540%Cr; Al2O3); 0.30.4% NH4Cl-balanceSlurry aluminizing(c) Al Powder of (100% Al) in colloxiline solution of amilacetate

AlSi Powder of (Al550%Si) in colloxiline solution of amilacetateElectric arc vacuum deposition(d) Al Alloy: 100% Al

AlSiY Alloy: Al base; 4.015.0% Si; 0.62.0% Y

(a) Ref 15, 16. (b) Ref 17, 18. (c) Ref 16, 19. (d) Ref 20, 21

Table 2.3 Superalloys used for aircraft turbine blades

Chemical composition(b), wt%Alloy (a) Cr Ti Mo W Re Ta Al Co Hf Nb B Zr CNi10CrWMoCo 10.0 10.0 5.0 4.5 5.0 0.10JS6K 11.0 3.0 4.0 5.0 5.5 4.3 0.15JS6U 8.5 2.5 1.8 10.0 5.5 9.5 1.0 0.15VJL12U(c) 9.5 4.5 3.0 1.4 5.5 9.5 0.8 0.15JS6F(c) 5.5 1.0 0.9 12.0 5.3 9.4 0.9 1.6 0.11JS30 7.0 1.9 0.7 11.8 5.2 8.5 0.8 0.9 0.015 0.05 0.15JS26(c) 5.0 1.0 1.1 11.7 5.8 9.0 1.6 0.015 0.05 0.15JS32 5.0 1.0 8.3 4.0 4.0 6.0 9.0 1.5 0.015 0.05 0.15MAR-M-200 9.0 1.9 12.5 4.7 10.0 1.8 0.015 0.05 0.14MAR-M-002 9.0 1.7 11.0 2.5 5.5 10.0 1.5 0.015 0.05 0.15CMSX-4 6.5 1.0 0.6 6.0 3.0 6.5 5.6 9.0 0.1

(a) JS alloys are designed in All-Russian Institute of Aviation Materials (VIAM). (b) Average values. (c) These alloys contain 1.0% V. Source: Ref 22, 23

Fig. 2.3 Depth of high-temperature corrosion inNi20Cr12AlY overlay coating vs. test temperature at

testing for heat resistance (); 200 h hot corrosion (o), and enginetest, 100 h (D)

the highest life under certain running conditionsand point out ways to improve existent coatingsand develop alternative systems.

Two temperature regimes, one at 600 to 850C and one above 1050 C (Fig. 2.3), have beensingled out as the ones where coatings are sub-jected to intensive damage. This result wasachieved while analyzing aluminide-based coat-ings for their protectiveness. At 600 to 850 C,hot corrosion was noted to develop in coatingsowing to the presence of aggressive compoundsconstituted mostly of sulfur and vanadium. Thetemperature boundaries and intensity of corro-sion will depend on a number of factors: com-position of protective coating, quantity of ag-gressive compounds on the surface, and durationof tests. During engine tests, the intensity of hotcorrosion will be affected (besides temperature)by the gas pressure in the turbine and by designfeatures of the blades and vanes which can fa-cilitate or impede condensation of aggressivecompounds on different areas of the surface.

In the interval from 900 to 1000 C, onlymodest oxidation has been noted due to the highprotective properties of oxides Al2O3 Cr2O3forming on the surface of this type of coating.In laboratory heat resistance tests in the tem-perature range of 900 to 950 C, the diffusionaluminide-based coatings, 50 lm thick, with


aluminum content of 34 to 36%, have mani-fested their ability to protect superalloys fromoxidation during 10,000 h.

At temperatures above 1000 C, the rate ofcoating oxidation rises. At the same time, thediffusion processes between the coating and al-loy speed up, resulting in the aluminum andchromium content of the coating dropping ap-preciably and the rate of oxidation growing. Adiffusion coating 50 lm thick (34 to 36% Al)has exhausted its protective ability in 300 to 500h when tested in the laboratory at 1100 C forheat resistance.

The trends that have been noted during labo-ratory tests are also observed on engine tests.Hot corrosion develops when engines are runcontinuously. The rate of coating damage is gen-erally lower when testing coated blades in therange of 700 to 850 C (in comparison with lab-oratory tests). This phenomenon is attributed tolower amounts of aggressive compounds fromthe gas flow in contact with the blade surfaces.

Oxidation processes at 900 to 950 C go onturbine blades more intensively than during heatresistance tests. Generally, coatings are oxidizedduring high-temperature cycles of engine run-ning. At such temperature, the life expectancyof a diffusion coating 50 lm thick (34 to 36%Al) spans between 6000 and 8000 h when theengine is run at maximum load. However, lifeexpectancy dramatically drops when the tem-perature on blade surface is elevated up to 1050to 1100 C. When the blades were rig-tested at1050 C, their diffusion-type protective coat-ing failed in 80 to 100 h during engine runningat high temperatures.

Hot Corrosion of Turbine Blades. Sinceaviation fuels contain insignificant quantities ofsulfur and vanadium, they do not destroy coat-ings through hot corrosion below service livesof 5,000 to 10,000 h. The environment is themain source of aggressive components enteringthe hot channel of aircraft engines. This applies,first of all, to sea regions, where a diverse varietyof compounds contained in the seawater get intothe engine hot channel.

Another source of aggressive components isthe waste gas of industrial plants. The locationof test rigs or the region of engine operation maybe the cause of this mechanism of coating fail-ure.

Investigations carried out have revealed twomodes of hot corrosion on turbine blades of air-craft engines. Those are (a) pit and (b) uniformcorrosion modes (Ref 16). Pit corrosion emerges

and progresses first of all on those areas of theblade surface that work continuously at 620 to760 C (Fig. 2.4). Such surfaces are located onthe blade airfoil near the fir-tree root and shroudplatform. As early as in the first dozen hours oftesting or operation, numerous pits may appearon the said surfaces. These pits, filled with cor-rosion products and looking like small nodules,gradually, as time goes on, spread towards theblade central region.

Pit corrosion proceeds on most-lengthy work-ing cycles of engine operation. As the analysisof turbine work parameters has shown, thehigher the gas flow velocity, the lesser the cor-rosion damage. Variation of pressure and air sur-plus coefficient produces no significant effect onblade corrosion. Of much greater significanceare the blade design features and the local aero-dynamics. For the same temperatures on the suc-tion and pressure surfaces, the pressure surfacewill suffer more from pit corrosion than the suc-tion surface where corrosion attack is much lesssevere. The poorer the fuel quality, the more se-vere is the corrosion.

Research has been conducted into the micro-structure of diffusion coatings in corrosion-in-flicted areas on the blades removed from differ-ent engines, either rig-tested or service-run, indiverse conditions. The research has proved thatthe microstructure was the same in all casesstudied. The oxides penetrating into coatingouter zones are encircled with light edging, con-sisting of Ni3Al compound (Fig. 2.4b). Thelonger the time of testing or running, the greaterthe depth of penetration of oxides into the coat-ing. This, however, continues only until innerzones with high contents of chromium-basephases are reached. At this point, the oxide pen-etration is lowered and the inner zone of thecoating retards corrosion penetration into basicmetal. The analysis of the chemical compositionof pit corrosion products has shown that, apartfrom the coating elements, they contained from5 to 10% V and from 0.5 to 1.0% S.

The higher the blade working temperatures,the less the pit corrosion. Thus, no pit corrosionwas observed on turbine blades from enginesthat had been running at 900 C for as long as8,000 to 10,000 h.

Elevation of temperatures, higher compres-sion rates, and greater gas flow velocities in en-gine turbines have altered the outer appearanceof corrosion-stricken blades (Fig. 2.5). Thisshould be attributed, first of all, to the elevationof gas temperatures up to 1330 to 1400 K and


Fig. 2.4 (a) Pit corrosion location vs. blade surface temperature (service time around 4000 h); (b) Blade surface in the zone of pitcorrosion. 20; (c) Microstructure of diffusion coating in pit corrosion zone. 300. (Light phase around oxides is Ni3Al.)

to the resulting changes in the composition, flu-idity, and thermodynamic properties of aggres-sive compounds. The condensation of aggres-sive compounds is helped by a high temperaturegradient between gas flow and surface of cooledblades. Higher gas flow pressure and velocityenhance erosion effects upon the corrosion prod-ucts by briskly sweeping the latter off the bladesurfaces.

Instead of formation of individual corrosionpits, uniform corrosion has been noted. The lat-ter affects large areas of the blade surface, par-ticularly the pressure surface and the leadingedge exposed to temperatures of 780 to 850 C.Consisting of aluminum, chromium, and nickeloxides, the corrosion products penetrate deeplyinto the coating and the superalloy (Fig. 2.6).

A unique feature of hot corrosion is that a thin


Fig. 2.5 Uniform corrosion zone location vs. blade surface temperature (test time around 3000 h).

Fig. 2.6 Microstructure of JS6U-alloy blade surface underlayer of oxides. 5000

layer of metal with sulfur precipitates emergesunderneath the corrosion products in the high-temperature zone. Such microstructure was ob-

served in all blades stricken with hot corrosion.As for vanadium, its content in corrosion prod-ucts is either insignificant or nil.

The uniform corrosion is typical of continu-ous engine running. Uniform hot corrosion isnoted on engine blades running at high tempera-tures (gas temperature from 1650 to 1700 K).This should be attributed to effective coolingsystems providing, on certain surfaces (e.g., inthe lower zone of the blade airfoil), favorabletemperature conditions for aggressive com-pounds to condense.

The choice of a coating to protect engineblades from hot corrosion within the requiredlife period is generally based on the results oflaboratory tests conducted in line with unifiedmethods. Coating service life is understood astime to coating failure caused by hot corrosion.The correctness of coating choice and predictionof its service lifespan is totally dependent onhow sound the selected laboratory methods are.Accelerated tests may be fraught with such se-rious errors that the results of the tests cannot beused to predict lives. Moreover, one cannot ruleout errors involved in comparative analysis ofthe properties of different coatings.

Life expectancies of coatings can be predictedfrom the results of tests described in Phase andStructural Changes in Coatings during High-Temperature Tests, Chapter 5. The approxi-mate life expectancy of a new coating in termsof its corrosion-resistance can be inferred from


Fig. 2.7 Service life of Ni-Co-Cr-Al-Y coatings 100 lm thickcalculated from data on laboratory-tested Ni-Co-Cr-

Al-Y alloys in GZT ash, 850 C, 200 h

the rate failure of an existing engine-run coatingand the results of comparative hot-corrosion lab-oratory tests of both the existing and new coat-ings.

The suggested approach to inferring blade-lifeexpectancy is simplified and substantiated bytest experience in using the laboratory methodsand then applying the data generated in selectingthe most appropriate coatings for aircraft en-gines. The best way of improvement here (i.e.,in estimating the coating life expectancies) is notto sophisticate further the relevant calculationsbut, rather, to select the right test methods andto accumulate data on the properties of differentcoatings. The problem of accumulation of dataon hot corrosion can be solved, at reasonablecost, by using the alloys, imitating possible com-positions of coatings, and applying statisticalmethods of processing the test results.

Figure 2.7 illustrates the results of life-expec-tancy calculations for 100 lm thick coatingstested at 850 C under laboratory conditions forthe Ni-Co-Cr-Al-Y alloys. The same data caneasily be obtained for other systems too. Addi-tional investigations will be required to inferfrom these data life expectancies for coatedblades.

Ceramic coatings provide new opportunitiesto lengthen blade-life expectancy in a hot-cor-rosion environment. The effects of such coatingscan be dual: Adequate protection of metal coating from

aggressive components with proportionallengthening of life expectancy of a double-

layer coating (ceramic-metal coating). Inpractice, the pertinent solution may be to in-crease coating thickness.

The ceramic coating alters the thermal stateof blade surface conditions (elevates the sur-face temperature). This reduces the amount ofaggressive compounds condensing on the sur-face. It also enhances the fluidity of such ag-gressive compounds, thus facilitating their re-moval by gas flow.

This factor has been confirmed by tests. Whentested in the laboratory, the ceramic coatings hadno significant effect, though tests on power unitsshowed the increase in life expectancy of blades.Further investigations and development of spe-cial ceramic compositions will be necessary tobring ceramic coatings on a par with the require-ments imposed on coatings for hot-corrosionprotection.

Oxidation of Turbine Blade Operating atTemperatures above 1000 C. As the tem-perature of a blade surface increases from 750to 850 C to 900 to 950 C, the life expectancyof coatings rises. In this temperature interval,diffusion coatings guarantee adequate protectionof aircraft-engine turbine blades within 6000 to8000 h. Increasing blade surface temperature upto 1050 to 1100 C results in appreciable de-creases in diffusion coating life expectancy. Atsuch temperatures the coatings fail mostly dueto consumption of aluminum for the formationof the protective oxide. Besides, the aluminumcontent was reduced owing to interdiffusion ofaluminum and nickel between the coating andsuperalloy. As Fig. 2.8 shows, a turbine blade,after being tested for 100 h, displayed full oxi-dation of the coating (blade wall thereafter) in30 to 40 h when exposed to a temperature of1100 C.

Figure 2.9 illustrates oxidation rates in a dif-fusion coating deposited on the blade from su-peralloy JS6U. The coating has been built up byslurry aluminizing and it had an aluminum re-serve of 40 g/m2. After being tested for100 h, the coating outer zone was found to befully destroyed at a narrow strip on the leadingedge 2 to 4 mm broad and 10 to 20 mm highwhen exposed to the maximum temperature of1050 C. The consumption rate of aluminum forprotective oxide formation in the leading edgeregion was 0.5 g/m2 h. With the temperatureat the blade inlet edges elevated to 1100 C, thediffusion coating with aluminum reserve of 60


Fig. 2.8 External appearance of a blade with diffusion coat-ing after 100 h of testing

Fig. 2.9 (1) Oxidation depth of diffusion and (2)Ni20Cr12AlY overlay coatings on the leading edge

of a blade during testing

to 70 g/m2 was found to be fully oxidized in 30to 40 h of tests.

Figure 2.9 illustrates the depths to which ox-idation has penetrated in the overlay coatingNi20Cr12AlY, 120 lm thick, with aluminum re-serve 200 g/m2. Measurements were taken onthe leading edges of turbine blades. When run-ning under maximum loads, the temperature onthe leading edge was 1050 C. Test durationis indicated as a summed value for all running.In comparison with diffusion coatings, the over-lay coatings had greater aluminum reserves,which made the life expectancy 4 times aslong as that of diffusion coatings. The designedaverage consumption of aluminum and chro-mium for protective oxide formation inNi20Cr12AlY coating varied within 0.5 to 0.6g/m2 h.

As for the phase transformations and struc-tural changes in the coating Ni20Cr12AlY oc-curring on the leading edge of the blade, theyare similar to those typical of laboratory tests forheat resistance. After 140 h, a thin layer of c-solid solution is formed on the coating surface,controlling the kinetics of oxidation processes.As test time gradually increases to 300 h, the c-solid solution covers a greater zone and itsboundary spreads further into the coating. Thediffusion of aluminum, and to a much lesser ex-tent chromium, from the coating central zonestoward the surface makes the coating composi-tion change in the direction NiAl cNi3Al cc. In this case 10 to 15% of aluminumand chromium diffuse under the coating into thezone of superalloy interaction.

As the Ni20Cr12AlY coating is becoming ox-idized, the surface of a blade is deformed in arippled or folded fashion (Fig. 2.10). Thisphenomenon can be attributed to gas flow pres-sure on the coating that is soft at these tempera-tures. Its deformation is affected by compressionstresses occurring in the coating at temperaturesabove 1000 C.

Should the working temperatures on turbineblades with the overlay coating Ni20Cr12AlYbe raised to 1100 to 1150 C, the coating pro-tection capability reduces dramatically, pro-vided that the described mechanism of coatingdamage prevails. Figure 2.11 shows typicalstructures of Ni20Cr12AlY coating on turbineblades due to overheating from 1000 to 1150 Cin certain zones.

The oxidation mechanism remains the sameon all surface areas. Its nature can be explainedby consumption of certain elements (aluminumstanding the first) for oxide formation. The rate


Fig. 2.10 External appearance of Ni20Cr12AlY coating on a turbine blade after 500 h of testing. (a) Leading edge. (b) Pressuresurface. 10

of aluminum and chromium consumption varieswith the temperature level. While in zones at1000 C, the coating has oxidized to an insig-nificant depth (Fig. 2.11a); in the zone at 1150C, phase transformations associated with c-solid solution formation have been completed.Aluminum content there has been reduced to3%, and the processes of internal oxidation havestarted (Fig. 2.11d). Aluminum and chromiumlevels consumed to form the interaction zone donot exceed 10 to 15%, even at maximum tem-peratures.

The heat resistance of diffusion and overlaycoatings is discussed against the background oflaboratory tests, enabling protective propertiesof different coatings to be compared, given thatcorrect methods and proper criteria have beenchosen. However, the test data, pertinent to spe-cific changes of mass (g/m2) cannot be used incalculations of coating life expectancies onblades. For the same mechanism of oxidation,the extent of blade damage from oxidation willbe many times that of laboratory samples.

In laboratory tests and in tests on the blades,the protective function is performed not by thecoating, but rather by the oxides that are formedon its surface. In fact, the coating acts as a re-serve that supplies aluminum and chromium forthe formation of oxides if they are damaged. Inthis case it will be the aluminum reserve (mAl)that emerges as a criterion for coating life ex-pectancy (Ref 16). This has been substantiatedby data, which showed that all aluminum con-tained in a diffusion coating on superalloys iseventually consumed for oxide formation. Thealuminum reserve criterion correlates wellwith the results of both laboratory and field tests.

Aluminum reserve, mAl in a diffusion coat-ing is found from the following function:

n

m K X q hAl i i ii1

in g/m2 where Xi is aluminum mass fraction ina definite zone of coating; qi is the density of


Fig. 2.11 Microstructure of Ni20Cr12AlY coating in the blade zones with gradual increasing of working temperatures from (a) 1000C to (d) 1150 C. 500. In (a), (b), and (c), a dark phase in the coating is c-solid solution; a light one is Ni3Al. In (d), a

dark phase in the coating is oxides; a light one is c-solid solution.

that zone of coating; hi is the thickness of thatzone of coating; and K is the alloying factor.

Aluminum reserve in the coating equals theincrease of specific mass of samples at alitizing,

and it can be found out experimentally. The ef-fect of alloying the diffusion coatings can becounted by introducing the coefficient K ,which for aluminized coatings equals unity (K


1), while for the modified coatings alloyedwith chromium, silicon, and yttrium, it takes thevalue KAl1.11.6.

The aluminum reserve criterion can also beused in comparing life expectancies of overlaycoatings, having the same or nearly the samesystem of alloying. When determining alumi-num reserve in Ni-(Co)-Cr-Al system, Cr con-centration must be taken into account too. Otherelements must be counted by introducing suit-able alloying factor K:

m (X 0.5 X ) q h KAl Al Crin g/m2. The coefficient 0.5 for chromium massfraction XCr in this case was taken from the re-gression equation, interlinking the effect of theseelements (Al and Cr) on the oxidation kineticsin Ni-Co-Cr-Al system alloys. The tests wereconducted at 1200 C.

When aluminum consumption for oxide for-mation on turbine blades is known, then thecoating life expectancy will be directly propor-tional to the aluminum reserve. The coating lifeparameter at temperatures above 1000 C willdepend on the effects of interaction between thecoating and superalloy. In both laboratory andengine tests, a part of the aluminum reserve islost in diffusion of aluminum and chromiumfrom the coating into the superalloy. In labora-tory tests the losses maybe considerable com-pared with aluminum consumption for oxideformation, but in testing on turbine blades theoxide consumption will always exceed interac-tion losses.

Since interaction zone dimensions (betweendiffusion/overlay coatings on the one hand andnickel-base superalloys on the other) do not ex-ceed 30 to 40 lm when tested/run for up to 1000h at temperatures below 1000 C, these may bedisregarded.

Thermal Fatigue of Turbine Blades. As aresult of higher gas temperatures in aircraft en-gines and the development of turbine bladeswith highly efficient cooling, the thermalstresses in blades have risen. Thermal fatiguecracks in blades have become the commonestdefect. It has been suggested that all thermal fa-tigue cracks occurring in turbine blades of dif-ferent aircraft engines can be conventionallyclassified into two types, proceeding from eithercrack origins within the substrate or from ther-mal stress cracking in coatings (Ref 24): Thermal fatigue cracks of the first type Thermal fatigue cracks of the second type

(fragmentation)

Differences in the appearance and microstruc-ture between the two types are as follows:

The first type of thermal fatigue cracks aregenerally few. In the process of operation,they penetrate deeply into the blade wall, re-ducing the thermal stresses arising betweencertain zones of the surface and eventuallyleading to blade failure.

The second type of thermal fatigue cracks(fragmentation) can appear on blade surfacesduring the first hours of blade service life.Usually these are numerous fine cracks, oc-casionally covering the full blade surface.The depth of their penetration is usually lim-ited to within the coating thickness; thereaf-ter, the cracks either discontinue their growthor continue at a very slow rate. The uniquefeature of such cracks is that their distributiondensity is very high; distances between adja-cent cracks are comparable with the coatingthickness.

The research into the fragmentation phenom-enon shows that its occurrence is due to the fol-lowing:

High gradients of temperatures across theblade wall section, the blades being providedwith highly efficient cooling systems

Increased transient running on the engine Sharp changes of heating on transient running

(thermal shocks), during which the rate oftemperature rise and drop is measured in hun-dreds of degrees per second

Difference between the thermal expansioncoefficients (TECs) of coating and superalloy

Insufficient ductility of coatings in the areawhere maximum strains occur

Fragmentation emerges as a result of the coat-ing-superalloy system exposure to high thermalstresses and strains. Illustrated in Fig. 2.12 is aturbine blade made from the superalloy JS6F,with a diffusion coating and after testing for 100h. The entire pressure surface of the blade is cov-ered with a net of the second type of thermalfatigue cracks, penetrating not beyond the coat-ing thickness.

The optimum way of removing fragmentationis to reduce thermal stresses in the blade coatingto as low a level as possible. Bringing closer thecoefficients of thermal expansion of the coatingand superalloy can do this. While calculating thecyclic service life of a coated blade, the designercan predict the emergence of fragmentation


Fig. 2.12 (a) External appearance of a blade with fragmentation exposed by capillary flaw detection after 500 h of testing. (b)Microstructure of a coating in a fragmentation zone. 500

cracks and thereby choose coatings that offer thebest life expectancy.

The principles of a cyclic service life calcu-lation for a coating deposited on a blade surfaceare dealt with in Chapter 8, Strength Designingof Turbine Blade Protective Coatings. The cal-culation is based on stresses and strains arisingin the coating due to centrifugal forces, bendingmoments, nonuniformity of blade thermal con-ditions, and thermal stresses caused by differentTECs of the coating and the alloy protected.

In performing calculations, it must be bornein mind that the composition of the coating andthe whole complex of its physical and mechan-ical characteristics are subject to continuouschanges with time. The rate of such changes ishigher when the temperature of the blade surfacerises. The coating cyclic life expectancy must becalculated for the original and final coating com-positions. Should the coating properties undergoessential changes with time, interim calculationsmay be necessary.

While protecting the blade from oxidation, thecoating simultaneously produces extra stressesand strains on the blade surface. These stressesand strains damage the coating and reduce itscyclic service life. This reasoning holds onlywith respect to an ideal blade, because in actualconditions, when uncoated blades are exposed

to vigorous oxidation, the resultant loss of bladeservice life far exceeds the losses caused by thecoatings.

With high strains in blade coating (Deic 0.5%), the rate at which thermal fatigue cracksemerge in the coating (fragmentation) may exceedthe rate of blade surface damage from oxidation.In this case, the deposition of coatings would bringabout a shortened cyclic life of the blade.

Figure 2.13(a) and (b) illustrate the results ofcyclic tests conducted on engines whose turbineblades are made from superalloys JS6F andJS6U with diffusion and overlay coatingsNi20Cr12AlY, Ni20Co20Cr12AlY, andCo25Ni20Cr12AlY. The number of blades withfragmentation grows with test time. After 10 to50 h, all blades are covered with a network ofcracks. No cracks have been observed on theblades with Ni20Cr12AlY coating.

Spalling of Ceramics on Turbine Bladeswith Thermal Barrier Coatings. The develop-ment of thermal barrier coatings (TBC) offersnew horizons in further lengthening of turbineblade life expectancy. Not only does the TBCprotect the blade surface from hot corrosion andoxidation, it also offers thermal protection andthus shields the superalloy from softening dur-ing high-temperature exposure. The TBC is ba-sically a structure consisting of a bond coat and


Fig. 2.13 Number of (a) JS6F and (b) JS6U blades with frag-mentation vs. time of testing. 1, Ni20Cr12AlY

coating; 2, aluminized coating; 3, Ni20Co20Cr12AlY coating; 4,Co23Ni20Cr12AlY coating

an outer ceramic layer. The latter is based onzirconium oxide and serves as the TBC mainelement.

The properties of the TBC ceramic layer areentirely dependent on the mode of its formation.If the electron beam (EB) technique is appliedand if all process problems have been satisfac-torily settled, the ceramic layer will have a spe-cific columnar structure. The typical features ofsuch a structure are low tensile strength and lowelastic modulus in the direction normal to thecrystallite-growth axis. As thermal stresses andstrains arise, the ceramic layer freely breaks intofragments. The individual fragments are sizedby such factors as stress level and the ceramiclayer thickness. Any strains in the blade causefragmentation. This property protects the ceram-ics from high stresses and spalling.

Life expectancy of a TBC working at tem-peratures above 1000 C is totally dependent onthe adhesion of the ceramic layer. The EB tech-nique enables ceramic layers with original ad-hesion strength higher than 70 MPa to be ob-tained.

Depositing a TBC on the blade alters its char-acteristics. This must be noted when designingblades with TBC. The thermal barrier effect DTis controlled by the ceramic layer thickness, d; itsthermal conductivity, k; and the thermal fluxthrough the blade wall, Q such that D T Q d/ k.

There are certain restrictions on the extent towhich ceramic thickness can be increased. Inview of its low tensile strength, the ceramic layeris unable to carry mechanical loads produced bycentrifugal forces. Hence, extra loading is ap-plied to the turbine blades (and disk). This dic-tates that it is better to make the ceramic layerof variable thickness throughout blade surface.Maximum thickness is needed in the areas wheremost intensive thermal fluxes will be experi-enced or where the thermal field is nonuniform.In other zones, the ceramic layer may have min-imum thickness or be absent. The optimumthickness of a ceramic layer on the blade varieswithin 120 to 250 lm. At lower thickness, theremay be rapid oxidation of the ceramic-bond coatboundary and poor adhesion strength. At greaterthickness, considerable stresses resulting in spal-lation are likely to occur.

The composition of a bond coat and its thick-ness are chosen as when using metal heat resis-tant coatings. Since TBCs are employed onblades with high thermal stresses, preferenceshould be given to bond coats offering higherresistance to thermal fatigue cracking. Once theceramic layer has spalled, the bond coat has toprotect the blade from oxidation.

The main cause of TBC failure is spallationof the ceramic layer from the blade while beingexposed to higher temperatures for a long time(Fig. 2.14). The spallation is assisted by tear-offstresses occurring on the convex blade surfaceswhen thermal stresses occur inside the ceramiclayer. These internal compression stresses reachtheir maximum as the blade is cooled down toroom temperature.

The onset of TBC failure is betrayed by a net-work of microcracking (fragmentation) appear-ing on the surface of the blade. Fragmentationof a ceramic layer does not indicate its failure.The blades in such a state may be exploited forquite a long time. Fragmentation appears onsome surface areas where tensile stresses can begenerated in the ceramic layer under the certainengine operation conditions.

TBC life expectancy can be estimated by ad-hesion strength that has to be maintainedthroughout the service life of the coating. Ac-cording to this criterion, the adhesion strength


Fig. 2.15 Elongation of blades with (1) Ni20Cr12AlY coat-ing and with (2) TBC vs. service time

Fig. 2.14 Blades with a spalled ceramic layer after 800 cy-cles of testing at 200 1100 C

must exceed the shearing stress needed to sepa-rate the ceramic coating from the surface. Fail-ure of the ceramic layer in the TBC is evidenceof a reduced adhesive strength on the ceramic-

bond coat boundary. Spallation of ceramics afterhigh-temperature tests takes place even in theabsence of stresses in the sample. The effect isassociated with the formation of various Al2O3-type oxides on the ceramics-metal interface.Such oxides reduce the adhesive strength andalter the stress condition on the boundary be-cause their physical and mechanical propertiesdiffer from those of ceramic and metal layers inTBCs.

All the above processes, leading to ceramiclayer spallation, can be allowed for by the inte-gral characteristic of adhesion strength, ra,found experimentally in relation to the tempera-ture and duration of tests. The criteria applicableto the determination of adhesive strength for aceramic layer on blades with TBC are discussedin Chapter 8, Strength Designing of TurbineBlade Protective Coatings.

To retain adhesive strength with time is themost formidable challenge to further lengthen-ing of TBC life expectancy. Two prerequisites,however, must be fulfilled for this: to reduce ox-ygen diffusion mobility in the lattice of zirco-nium oxide and to prevent fragmentation of ce-ramics in high-temperature applications.

Benefits that TBCs provide include the op-portunity to elevate the gas temperature at theturbine inlet without changing the temperatureof the turbine blades. Most often, however,TBCs are used to extend blade service life. Es-timation of the TBC benefit is a difficult task.The issue is that while comparing the bladeswith and without TBC, some side effects occurthat affect the blade temperature (e.g., reductionof perforated hole diameter). Analysis of plasticstrains in identical blades with and withoutTBCs, tested in the engine under similar con-ditions revealed lesser elongation in blades withTBC in comparison to their counterparts withoutTBC (Fig. 2.15); that is, they demonstrate higherhigh-temperature strength.

Use of TBCs allows higher temperatures inengine turbines to be employed, thoughZrO2 Y2O3 oxide imposes significant restric-tions on its use at the temperatures above 1150C. These restrictions are due to high oxygen-diffusion mobility in zirconium oxide. The de-velopment of new ceramics for high-tempera-ture application is the most promising route tofurther TBC improvement.

REFERENCES

1. N.D. Kuznetsov, Problems of ThermocyclicStrength of Gas Turbine Engines Elements,Probl. Prochn., Vol 6, 1978, p 38


2. Y.B. Chertkov, K.F. Rybakov, and V.H.Zrelov, Impurities and Purification Meth-ods of Oil-based Fuels, Moscow, Khimiya,1970

3. Ya.B. Chertkov and V.G. Spirkin, Use ofJet Fuels in Aviation, Moscow, Transport,1974

4. I.H. Shishkov and V.B. Belov, Aviation Lu-bricants, Fuels, and Special Fluids, Mos-cow, Transport, 1979

5. I.A. Birger, B.F. Schorr, and I.V. Demi-anushenko, Machine Elements, Moscow,Mashinostroyeniye, 1975, 455 p

6. N.D. Kuznetsov, Designing the Strength ofa Long Life Gas Turbine Engine, Probl.Prochn., 5, 1976, p 39

7. N.D. Kuznetsov, Strength of Gas TurbineEngine Turbine Elements under ComplexLoads and Related Problems, Probl.Prochn., 3, 1982, p 1014

8. M.Ia. Ivanov and V.P. Pochuev, Problemsof Designing High-Temperature Turbinesof Modern Aircraft Engines, Conversion inMachine Building of Russia, 5, 2000, p 3446

9. V.H. Abiants, Jet Engines, Moscow, Mash-inostroyeniye, 1985

10. Ye.N. Bogomolov, Service Processes inCooled Turbines of Gas Turbine Engineswith Perforated Blades, Moscow, Mashi-nostroyeniye, 1987

11. S.Z. Kopelev, M.N. Galkin, A.A. Harin, andI.V. Chevtchenko, Thermal and HydraulicCharacteristics of Cooled Gas TurbineBlades, Moscow, Mashinostroyeniye, 1993,176 p

12. G.P. Dolgopolenko, M.D. Romanov, andV.V. Gatin, Plane and Helicopter Gas Tur-bine Engines, Moscow, Mashinostroyeniye,1983

13. D.H. Boon, Overlay Coatings for ImprovedOxidation/Corrosion Protection and Duc-tility for High-Temperature Applications,Airco Temescal, May, 1977, 12 p

14. B.A. Movhan and I.S. Malachenko, HeatResistant Coatings Deposited in Vacuum,Kiev, Naukova Dumka, 1983, 232 p

15. A.G. Andreeva, V.V. Terekova, and G.D.Fomenko, Heat-Resistance Coatings ofNickel-Base Alloys, High-TemperatureCoatings, Leningrad, Nauka, 1967, p 96110

16. Y.A. Tamarin, Heat Resistant DiffusionCoatings for Turbine Blades, Moscow,Machinostroenie, 1978, 133 p

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21. E.B. Kathanov and Y.A. Tamarin, Protec-tive Coatings: Effective Path of Increase ofReliability of the Blades, Aviation Materialson the Eve of 21st Century, Moscow,VIAM,1994, p 296304

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CHAPTER 3

Technological Processes forDeposition of Protective Coatings toTurbine Blades

A GREAT VARIETY of techniques for de-position of protective coatings to aircraft turbineblades have been designed. The use of eachtechnological process is warranted by a numberof reasons including requirements to the prop-erties of protective coatings, production equip-ment available, personnel experience, and com-mon practice. The preference of new techniquesto conventional ones is justified only in caseswhere fundamental changes in the properties ofprotective coatings are required. For example,new techniques are necessary in the case of de-signing a new turbine. As a rule, superseding inthe field of technology involves considerable fi-nancial expenditure connected with the purchaseof equipment and materials and training of per-sonnel.

Diffusion Coatings

Application of Diffusion Coatings inPowder Mixtures with HalideActivators

The diffusion saturation of superalloys withaluminum and with aluminum combined withother elements (Al-Cr, Al-Si) from the powdermixtures with halide activators is a widely usedmethod of aluminide coating formation on tur-bine blades of aircraft gas turbine engines. Thewide application of this method is justified bythe fact that different mixtures used for alumi-nizing ensure a wide range of aluminide coatingswith aluminum contents varying from 18 to 69%Al (here and subsequently all compositions are

given in percents by weight). Coatings alloyedwith chromium, silicon, and other elements canbe formed. The method of aluminizing in thepowder mixtures with activators offers such ad-vantages as simple production process andequipment, a wide temperature range for alu-minizing (from 500 to 1200 C), highly eventhickness and homogeneity of the producedcoatings, and low production cost.

Powder mixtures for aluminide coating for-mation are composed of aluminum or its alloyswith iron, chromium, and other elements (as anactive component), aluminum oxide (as an inertfiller), and a halide activator (usually NH4Cl).When the container with the saturating mixtureand aluminized parts is heated, the following re-actions take place:

NH

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Protective Coatings for Turbine Blades