EP0139396A1 - Combustion turbine blade with varying coating - Google Patents

Combustion turbine blade with varying coating Download PDF

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Publication number
EP0139396A1
EP0139396A1 EP84305738A EP84305738A EP0139396A1 EP 0139396 A1 EP0139396 A1 EP 0139396A1 EP 84305738 A EP84305738 A EP 84305738A EP 84305738 A EP84305738 A EP 84305738A EP 0139396 A1 EP0139396 A1 EP 0139396A1
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EP
European Patent Office
Prior art keywords
coating
blade
cooler
temperature
coated
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Granted
Application number
EP84305738A
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German (de)
French (fr)
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EP0139396B1 (en
Inventor
Edwin Alfred Crombie Iii
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CBS Corp
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Westinghouse Electric Corp
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Publication date
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Publication of EP0139396A1 publication Critical patent/EP0139396A1/en
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Publication of EP0139396B1 publication Critical patent/EP0139396B1/en
Expired legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/28—Selecting particular materials; Particular measures relating thereto; Measures against erosion or corrosion
    • F01D5/288—Protective coatings for blades
    • C—CHEMISTRY; METALLURGY
    • C23—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; CHEMICAL SURFACE TREATMENT; DIFFUSION TREATMENT OF METALLIC MATERIAL; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL; INHIBITING CORROSION OF METALLIC MATERIAL OR INCRUSTATION IN GENERAL
    • C23C—COATING METALLIC MATERIAL; COATING MATERIAL WITH METALLIC MATERIAL; SURFACE TREATMENT OF METALLIC MATERIAL BY DIFFUSION INTO THE SURFACE, BY CHEMICAL CONVERSION OR SUBSTITUTION; COATING BY VACUUM EVAPORATION, BY SPUTTERING, BY ION IMPLANTATION OR BY CHEMICAL VAPOUR DEPOSITION, IN GENERAL
    • C23C4/00—Coating by spraying the coating material in the molten state, e.g. by flame, plasma or electric discharge
    • C23C4/02—Pretreatment of the material to be coated, e.g. for coating on selected surface areas
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/005—Selecting particular materials

Definitions

  • the present invention relates to a turbine blade, for a land-based or marine combustion turbine, and in particular to turbine blades provided with coatings for protecting such blades.
  • Land-based or marine-type combustion turbines present difficult problems of blade materials. Near the tip of the blades, the temperatures are often 1700°F or more. Down near the base of the blade (near the shaft), temperatures are much cooler, for example, approximately 1000°F.
  • turbines are commonly operated with fuels containing corrosive impurities such as sulfur and vanadium. Further, corrosion-causing compounds such as sea salt or fertilizer are often ingested in with the air drawn in by the turbine compressor.
  • Such problems are significantly worse with land-based and marine combustion turbines as compared to aircraft (aircraft turbines are operated with cleaner fuel and significantly less contaminated air).
  • a turbine blade for a combustion turbine, said blade having a hot end portion at least a portion of which is designed to operate at a temperature in excess of 1500°F, a cooler end portion at least a portion of which is designed to operate at a temperature of less than 1250°F, and an intermediate portion at least a portion of which is designed to operate at between 1250 and 1500°F, said blade comprising: a hot end portion coated with a low creep-type coating which is resistant to high temperature oxidation, a cooler end coated with a ductile-type coating which is resistant to sulfide corrosion, and an intermediate portion, which is coated with a mixture of said hot end coating and said cooler end coating.
  • the coating on this portion of the blade must be creep resistant.
  • the cooler temperature of the blade especially those portions less than about 1250°F
  • an intermediate zone which is a mixture of the two coatings, must be used in order to prevent problems such as abrupt chemical discontinuities in the coating or stress concentrations.
  • the coatings are applied by plasma spraying and the intermediate portion is a graded coating giving a smooth transition from the hot end coating to the cooler end coating.
  • the range of temperatures of many gas turbine blades (as used herein, the term “blades” is used to mean turbine components having airfoil portions whether rotating or stationary, e.g., including the stationary parts which are sometimes called “vanes") generally exceeds the range of effectiveness of any single type of coating. This is in part due to the chemical/thermal stability of a coating in the various deleterious corrosive environments and partly due to the physical/mechanical properties of the coating itself.
  • This invention enables the use of a multiple composite coating system that enables the designer to maximize coating capabilities without the usual compromises (especially with regard to reduced physical/mechanical properties above or below the ductile/brittle transition temperature which are inherent to any given coating composition).
  • Figure 1 shows a blade with a portion designated 10 as the hot end part, and a cooler end portion 12.
  • a gas turbine blade may have an operating temperature profile ranging from about 1000°F at the base of the gas path surface to nearly 1800°F at the outermost tip region. Because the deleterious species and compounds are stable only through certain temperature ranges, application of a singular coating system has inherent limitations. A coating system which is most effective in preventing low temperature class II type corrosion in the range of 1000°F to 1450 o F, for example, could be applied through the lower portion of the airfoil and, a high temperature corrosion resistant composition applied to the upper portion (away from the center axis) of the airfoil where blade temperatures are highest.
  • the inherent ductility of most coating systems currently employed for environmental protection is generally equal to or greater than that of the base alloy to which it is applied. Premature failure of the blade due to brittle coating behavior and crack initiation is therefore not likely. Consequently, the coating that exhibits the best environmental protection may be utilized.
  • ductile-type coating means coatings which have a ductility of greater than or equal to that of the base metal at a given operating temperature. The correlation of coating and base metal ductily can be demonstrated in Figure 4.
  • Figure 2 shows three zones of coatings, with a hot-end coating 14 at the top and a cooler-end coating 16 at the bottom, with a transition zone 18 in the intermediate portion, the transition zone 18 being coated with a mixture of hot-end coating and cooler-end coating.
  • This transition zone 18 eliminates a sharp transition between the hot end coating and the cooler end coating. As a variation in the coating in an abrupt manner would result in poor thermal/mechanical properties and the possibility of uncoated areas resulting from less than perfect alignment, the transition needs to be gradual. Generally, this transition zone 18 will be at least 2 inch in height.
  • the coating is applied by plasma spray. If pack cementation techniques were used, additional handling would be required and masking would present difficulties with little or no control over interdiffusion between masked areas. It would be very difficult, therefore, to control the transition from one coating chemistry to the adjacent coating chemistry.
  • plasma spray could, for example, be done with an argon flood or low pressure plasma spray.
  • the transition zone could be formed by applying the coating compositions one at a time (e.g., by applying the hot-end coating with its thickness tapering from full thickness at the top end of the transition zone down to essentially zero thickness at the lower end of the transition zone and then applying the cooler-end coating with a maximum thickness at the lower transition zone and tapering down to near zero at the upper end of the transition zone, preferably followed by appropriate heat treatment), the coating is preferably applied by a system such as shown in Figure 3 where the transition zone 18 is accomplished by spraying a powder premixed by the hopper system.
  • the hot end coating composition designated "A”
  • the cooler end coating designated "B”
  • the feeding mechanism of the powder hoppers containing A and B compositions can be programmed to deliver the proper powder or powder mixture to the mixing vessel 26 which in turn supplies the gun 24.
  • the composition is initially 100% A, then an A-rich mixture becoming richer and richer in B, then a B-rich mixture and finally a 100% B coating.
  • all three zones 14, 18, and 16 will have a height of at least inch.
  • a coating system similar to Figure 3 can be used to coat more than three zones.
  • a third hopper with a "C" type coating composition could be added to apply an erosion resistant coating (or extended corrosion or lower temperature ductility coating, etc.) in this area (preferably using an additional transition zone).
  • prior-art single coatings can fail mechanically due to insufficient creep strength, but that this problem is generally in the high temperature regions, above the ductile/brittle coating transition temperature. Failures also can be caused by poor ductility below the brittle/ductile transition temperature of such a single coating.
  • a low temperature corrosion resistant coating with good low temperature ductility can be used on the lower portion of the blade airfoil.
  • a high temperature corrosion resistant coating with good high temperature creep resistance is applied to the upper portion of the airfoil. Problems at the interface of the two regions are avoided by using the blended composition in the intermediate zone of the airfoil.
  • the hot end (designed to operate above about 1500°F) can, for example, use MCrAlY coatings (with M being Ni and/or CO).
  • the cooler end coatings be similar to the MCrAlY (with M being Fe or FeNi or combinations thereof).
  • Figure 4 shows typical ductility variations with temperature for coatings and nickel-based superalloys.
  • the ductility of coating A is equal to or greater than the base metal alloys at temperature above about 1350°F and the ductility of coating B is equal to or greater than the ductility of the base metal alloys above about 1050°F.
  • the corrosion resistance of coating A is greater than that of coating B above about 1400°F while below about 1300OF coating B has a corrosion resistance at least as good as that of coating A.
  • the coating system of this invention provides improved protection against low coating ductility problems (above e.g. 1000°F) and against corrosion problems.
  • the transition zone which is coated with a mixture of the coatings is to be generally greater than inch in height.
  • the location of the transition zone can vary with various coatings, but at least a portion of this transition zone will be in a portion of the blade which is designed to operate at a temperature of between 1250 and 1500°F.
  • at least a portion of the transition zone is to be at a part of the blade which is designed to operate at between 1300 and 1450°F and most preferably at 1350°F.

Abstract

A turbine blade for a combustion turbine, in which the blade is provided with a hot-end coating, a cooler-end coating, and an intermediate transition zone with a mixture of the two end coatings. The blade has a hot end portion operable at a temperature in excess of 1500°F, a cooler end portion operable at a temperature of less than 1250°F, and an intermediate zone portion operable at between 1250 and 1500°F. The blade with hot end portion is coated with a low creep-type coating which is resistant to high temperature corrosion and oxidation, the cooler end portion is coated with a ductile-type coating which is resistant to sulfide corrosion, and the intermediate portion coated with a mixture of the hot end coating and the cooler end coating.

Description

  • The present invention relates to a turbine blade, for a land-based or marine combustion turbine, and in particular to turbine blades provided with coatings for protecting such blades.
  • Land-based or marine-type combustion turbines present difficult problems of blade materials. Near the tip of the blades, the temperatures are often 1700°F or more. Down near the base of the blade (near the shaft), temperatures are much cooler, for example, approximately 1000°F. In addition, such turbines are commonly operated with fuels containing corrosive impurities such as sulfur and vanadium. Further, corrosion-causing compounds such as sea salt or fertilizer are often ingested in with the air drawn in by the turbine compressor. Such problems are significantly worse with land-based and marine combustion turbines as compared to aircraft (aircraft turbines are operated with cleaner fuel and significantly less contaminated air).
  • According to the present invention, a turbine blade, for a combustion turbine, said blade having a hot end portion at least a portion of which is designed to operate at a temperature in excess of 1500°F, a cooler end portion at least a portion of which is designed to operate at a temperature of less than 1250°F, and an intermediate portion at least a portion of which is designed to operate at between 1250 and 1500°F, said blade comprising: a hot end portion coated with a low creep-type coating which is resistant to high temperature oxidation, a cooler end coated with a ductile-type coating which is resistant to sulfide corrosion, and an intermediate portion, which is coated with a mixture of said hot end coating and said cooler end coating.
  • Advantageously, it is to be noted that not only must the hotter portion be protected against high temperature oxidation type corrosion, but that the coating on this portion of the blade must be creep resistant. Conversely, the cooler temperature of the blade (especially those portions less than about 1250°F) must be protected against sulfide-type corrosion and must have high coating ductility to prevent crack propagation. Further, it has been found that an intermediate zone, which is a mixture of the two coatings, must be used in order to prevent problems such as abrupt chemical discontinuities in the coating or stress concentrations. Preferably, the coatings are applied by plasma spraying and the intermediate portion is a graded coating giving a smooth transition from the hot end coating to the cooler end coating.
  • The invention will now be described, by way of example, with reference to the following drawings in which:
    • Figure 1 is an elevation of a blade showing typical tip intermediate and base temperatures;
    • Figure 2 is a blade elevation showing three coating zones;
    • Figure 3 shows a system for applying the coatings of this invention, and
    • Figure 4 is a graph of typical ductilities for coatings and superalloy base materials at various temperatures.
  • The range of temperatures of many gas turbine blades (as used herein, the term "blades" is used to mean turbine components having airfoil portions whether rotating or stationary, e.g., including the stationary parts which are sometimes called "vanes") generally exceeds the range of effectiveness of any single type of coating. This is in part due to the chemical/thermal stability of a coating in the various deleterious corrosive environments and partly due to the physical/mechanical properties of the coating itself. This invention enables the use of a multiple composite coating system that enables the designer to maximize coating capabilities without the usual compromises (especially with regard to reduced physical/mechanical properties above or below the ductile/brittle transition temperature which are inherent to any given coating composition).
  • Referring to the drawings, Figure 1 shows a blade with a portion designated 10 as the hot end part, and a cooler end portion 12.
  • A gas turbine blade may have an operating temperature profile ranging from about 1000°F at the base of the gas path surface to nearly 1800°F at the outermost tip region. Because the deleterious species and compounds are stable only through certain temperature ranges, application of a singular coating system has inherent limitations. A coating system which is most effective in preventing low temperature class II type corrosion in the range of 1000°F to 1450oF, for example, could be applied through the lower portion of the airfoil and, a high temperature corrosion resistant composition applied to the upper portion (away from the center axis) of the airfoil where blade temperatures are highest.
  • At the hot end of the blade the inherent ductility of most coating systems currently employed for environmental protection is generally equal to or greater than that of the base alloy to which it is applied. Premature failure of the blade due to brittle coating behavior and crack initiation is therefore not likely. Consequently, the coating that exhibits the best environmental protection may be utilized.
  • At the cooler end 12 of the blade (generally here the end towards the 1000°F temperature), it has been discovered that unusually high ductility for these temperatures is required in addition to resistance to low temperature sulfide-type corrosion. As used herein, the term "ductile-type coating" means coatings which have a ductility of greater than or equal to that of the base metal at a given operating temperature. The correlation of coating and base metal ductily can be demonstrated in Figure 4.
  • Figure 2 shows three zones of coatings, with a hot-end coating 14 at the top and a cooler-end coating 16 at the bottom, with a transition zone 18 in the intermediate portion, the transition zone 18 being coated with a mixture of hot-end coating and cooler-end coating. This transition zone 18 eliminates a sharp transition between the hot end coating and the cooler end coating. As a variation in the coating in an abrupt manner would result in poor thermal/mechanical properties and the possibility of uncoated areas resulting from less than perfect alignment, the transition needs to be gradual. Generally, this transition zone 18 will be at least 2 inch in height.
  • Preferably, the coating is applied by plasma spray. If pack cementation techniques were used, additional handling would be required and masking would present difficulties with little or no control over interdiffusion between masked areas. It would be very difficult, therefore, to control the transition from one coating chemistry to the adjacent coating chemistry.
  • Although any type of plasma spray could be used, a non-oxidizing plasma spray system is thought to be the most practical. As most such coatings require an inert atmosphere or vacuum, such plasma spraying could, for example, be done with an argon flood or low pressure plasma spray.
  • Although the transition zone could be formed by applying the coating compositions one at a time (e.g., by applying the hot-end coating with its thickness tapering from full thickness at the top end of the transition zone down to essentially zero thickness at the lower end of the transition zone and then applying the cooler-end coating with a maximum thickness at the lower transition zone and tapering down to near zero at the upper end of the transition zone, preferably followed by appropriate heat treatment), the coating is preferably applied by a system such as shown in Figure 3 where the transition zone 18 is accomplished by spraying a powder premixed by the hopper system. Thus, the hot end coating composition (designated "A") and the cooler end coating (designated "B") are loaded into separate hoppers 20, 22. As the plasma gun 24 traverses the blade airfoil (under programmed computive control to maintain coating thickness profile), the feeding mechanism of the powder hoppers containing A and B compositions can be programmed to deliver the proper powder or powder mixture to the mixing vessel 26 which in turn supplies the gun 24. As the plasma gun 24 moves down the airfoil, the composition is initially 100% A, then an A-rich mixture becoming richer and richer in B, then a B-rich mixture and finally a 100% B coating. Generally all three zones (14, 18, and 16) will have a height of at least inch.
  • The specification of U.S. Patents Nos. 3,545,944 and 3,020,182 describe similar systems being used for different purposes.
  • It can be seen that a coating system similar to Figure 3 can be used to coat more than three zones. For example, if erosion (or corrosion or coating ductility) were a problem on some particular portion of the blade, a third hopper with a "C" type coating composition could be added to apply an erosion resistant coating (or extended corrosion or lower temperature ductility coating, etc.) in this area (preferably using an additional transition zone).
  • It should be noted that prior-art single coatings can fail mechanically due to insufficient creep strength, but that this problem is generally in the high temperature regions, above the ductile/brittle coating transition temperature. Failures also can be caused by poor ductility below the brittle/ductile transition temperature of such a single coating. By using different coatings in the high temperature region and the cooler temperature region, a low temperature corrosion resistant coating with good low temperature ductility can be used on the lower portion of the blade airfoil. A high temperature corrosion resistant coating with good high temperature creep resistance is applied to the upper portion of the airfoil. Problems at the interface of the two regions are avoided by using the blended composition in the intermediate zone of the airfoil.
  • It is felt that current coating systems are compromises in an attempt to perform adequately over a wide range of conditions, and are not optimized for providing either the high temperature corrosion resistance with high creep strength required in the hot end or the low temperature corrosion high ductility required in the cooler end.
  • Generally, it is anticipated that the hot end (designed to operate above about 1500°F) can, for example, use MCrAlY coatings (with M being Ni and/or CO). Similarly, it is anticipated that the cooler end coatings be similar to the MCrAlY (with M being Fe or FeNi or combinations thereof).
  • Figure 4 shows typical ductility variations with temperature for coatings and nickel-based superalloys. The ductility of coating A is equal to or greater than the base metal alloys at temperature above about 1350°F and the ductility of coating B is equal to or greater than the ductility of the base metal alloys above about 1050°F. The corrosion resistance of coating A is greater than that of coating B above about 1400°F while below about 1300OF coating B has a corrosion resistance at least as good as that of coating A. Thus, the coating system of this invention provides improved protection against low coating ductility problems (above e.g. 1000°F) and against corrosion problems.
  • Again, the transition zone which is coated with a mixture of the coatings is to be generally greater than inch in height. The location of the transition zone can vary with various coatings, but at least a portion of this transition zone will be in a portion of the blade which is designed to operate at a temperature of between 1250 and 1500°F. Preferably, at least a portion of the transition zone is to be at a part of the blade which is designed to operate at between 1300 and 1450°F and most preferably at 1350°F.

Claims (5)

1. A turbine blade, for a combustion turbine, said blade having a hot end portion at least a portion of which is designed to operate at a temperature in excess of 1500°F, a cooler end portion at least a portion of which is designed to operate at a temperature of less than 1250°F, and an intermediate portion at least a portion of which is designed to operate at between 1250 and 1500°F, said blade comprising: a hot end portion coated with a low creep-type coating which is resistant to high temperature oxidation, a cooler end coated with a ductile-type coating which is resistant to sulfide corrosion, and an intermediate portion, which is coated with a mixture of said hot end coating and said cooler end coating.
2. A blade as claimed in claim 1, wherein a mixture of said hot end coating is applied over at least inch of blade height.
3. A blade as claimed in claim 1 or 2, wherein said cooler end coating is chosen from the group consisting of MCrAlY, where M is Fe or FeNi.
4. A blade as claimed in any one of claims 2 to 3, wherein said coatings are applied by plasma spray.
5. A turbine blade, for a combustion turbine, constructed and adapted for use, substantially as hereinbefore described and illustrated with reference to the accompanying drawings.
EP84305738A 1983-08-29 1984-08-22 Combustion turbine blade with varying coating Expired EP0139396B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US52760683A 1983-08-29 1983-08-29
US527606 1983-08-29

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EP0139396A1 true EP0139396A1 (en) 1985-05-02
EP0139396B1 EP0139396B1 (en) 1988-07-13

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EP (1) EP0139396B1 (en)
JP (1) JPS6062603A (en)
KR (1) KR850001950A (en)
CA (1) CA1217433A (en)
DE (1) DE3472698D1 (en)
IE (1) IE55513B1 (en)
MX (1) MX159535A (en)

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WO1986004615A1 (en) * 1985-02-01 1986-08-14 Ingard Kvernes Aluminium-based article having a protective ceramic coating, and a method of producing it
FR2659088A1 (en) * 1990-03-02 1991-09-06 Gen Electric PROCESS FOR THE FORMATION OF DISKS CONSISTING OF TWO ALLOYS
WO1993005194A1 (en) * 1991-09-05 1993-03-18 Technalum Research, Inc. Method for the production of compositionally graded coatings
DE19615549A1 (en) * 1996-04-19 1997-10-23 Asea Brown Boveri Device for the thermal protection of a rotor of a high pressure compressor
EP0844368A2 (en) * 1996-11-26 1998-05-27 United Technologies Corporation Partial coating for gas turbine engine airfoils to increase fatigue strength
WO1998041735A1 (en) * 1997-03-18 1998-09-24 Abb Stal Ab A device for a guide blade arranged in a rotary machine
GR1003298B (en) * 1999-01-08 2000-01-18 Interceramic S.E. �.�. Method of selective priming of lamina with metal ceramic materials and construction of special features parts using them in a single production stage
WO2000025005A1 (en) * 1998-10-22 2000-05-04 Siemens Aktiengesellschaft Product with a heat insulating layer and method for the production of a heat insulating layer
EP1070769A1 (en) * 1999-07-22 2001-01-24 ALSTOM POWER (Schweiz) AG Process for coating a locally diversely stressed component
EP1101832A1 (en) * 1999-11-19 2001-05-23 Basf Aktiengesellschaft Method for the combinatorial production of a library of materials
WO2002018674A2 (en) * 2000-08-31 2002-03-07 Siemens Westinghouse Power Corporation Thermal barrier coating system for turbine components
US6534975B2 (en) 2000-01-15 2003-03-18 Alstom (Switzerland) Ltd Nondestructive method for determining the thickness of a metallic protective layer on a metallic base material via a different type of layer between the metallic protective layer and the metallic base material
EP1352992A2 (en) * 2002-04-12 2003-10-15 Ford Global Technologies, LLC A method for selective control of corrosion using kinetic spraying
EP1499754A1 (en) * 2002-04-30 2005-01-26 Ebara Corporation Abrasion resistant surface treatment method of a rotary member, runner, and fluid machine having runner
EP1895102A1 (en) * 2006-08-23 2008-03-05 Siemens Aktiengesellschaft Coated turbine blade
EP1927677A1 (en) * 2006-11-14 2008-06-04 United Technologies Corporation Thermal barrier coating for combustor panels
US20100143110A1 (en) * 2006-11-03 2010-06-10 Mtu Areo Engines Gmbh Vane for a compressor or a turbine of an aircraft engine, aircraft engine comprising such a vane and a method for coating a vane of an aircraft engine
EP2354454A1 (en) * 2010-02-02 2011-08-10 Siemens Aktiengesellschaft Turbine blade with variable oxidation resistance coating
US20120308842A1 (en) * 2011-05-31 2012-12-06 Schmidt Wayde R Composite article having layer with co-continuous material regions
EP2325441A3 (en) * 2009-11-13 2013-01-23 United Technologies Corporation Gas turbine engine component with discontinuous coated areas and corresponding coating method
EP2607787A1 (en) * 2011-12-22 2013-06-26 General Electric Company System and method for improved combustor temperature uniformity
EP2711441A1 (en) * 2012-09-21 2014-03-26 Reinhausen Plasma GmbH Device and method for creating a coating system
WO2014121998A1 (en) * 2013-02-05 2014-08-14 Siemens Aktiengesellschaft Fuel lances having thermally insulating coating
US9896585B2 (en) * 2014-10-08 2018-02-20 General Electric Company Coating, coating system, and coating method
EP3438414A1 (en) * 2017-08-04 2019-02-06 MTU Aero Engines GmbH Blade for a flow machine with different diffusion protection layers and method for production
US11261742B2 (en) 2013-11-19 2022-03-01 Raytheon Technologies Corporation Article having variable composition coating

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GB696715A (en) * 1951-02-07 1953-09-09 Metro Cutanit Ltd Improvements in blades for gas turbines and method of manufacture thereof
FR2367833A1 (en) * 1976-10-15 1978-05-12 Bbc Brown Boveri & Cie Corrosion protective coating for gas turbine parts - consists of electrodeposited nickel or cobalt contg. inclusions
FR2406000A1 (en) * 1977-10-17 1979-05-11 United Technologies Corp ARTICLES IN NICKEL, COBALT AND / OR IRON SUPERALLY COATED, RESISTANT TO OXIDATION AND WEAR
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Publication number Priority date Publication date Assignee Title
US2431660A (en) * 1944-12-01 1947-11-25 Bbc Brown Boveri & Cie Turbine blade
GB616432A (en) * 1946-08-30 1949-01-21 Power Jets Res & Dev Ltd Improvements relating to turbine rotors and the like bladed structures
GB696715A (en) * 1951-02-07 1953-09-09 Metro Cutanit Ltd Improvements in blades for gas turbines and method of manufacture thereof
FR2367833A1 (en) * 1976-10-15 1978-05-12 Bbc Brown Boveri & Cie Corrosion protective coating for gas turbine parts - consists of electrodeposited nickel or cobalt contg. inclusions
FR2406000A1 (en) * 1977-10-17 1979-05-11 United Technologies Corp ARTICLES IN NICKEL, COBALT AND / OR IRON SUPERALLY COATED, RESISTANT TO OXIDATION AND WEAR
GB2046369A (en) * 1979-04-04 1980-11-12 Rolls Royce Gas turbine blade

Cited By (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
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JPS6062603A (en) 1985-04-10
EP0139396B1 (en) 1988-07-13
KR850001950A (en) 1985-04-10
CA1217433A (en) 1987-02-03
MX159535A (en) 1989-06-28
JPH02521B2 (en) 1990-01-08
IE55513B1 (en) 1990-10-10
DE3472698D1 (en) 1988-08-18
IE842094L (en) 1985-02-28

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