Article having corrosion resistant coating

According to the invention, an article that is exposed to high temperatures, e.g., over 1000.degree. C. during operation is disclosed. In one embodiment, a turbine blade for a gas turbine engine includes a directionally solidified metallic substrate, e.g., a superalloy, which defines an airfoil, a root and a platform located between the blade and root. The platform has an underside adjacent the root, and a corrosion resistant ceramic overlay coating such as a stabilized zirconia is located on the underside of the platform and the neck. The applied coating prevents corrosion and stress corrosion cracking of blades in these regions.

BACKGROUND OF THE INVENTION
 The present invention relates generally to coatings for corrosion
 protection, and more particularly to an article having such a coating.
 Gas turbine engines are well developed mechanisms for converting chemical
 potential energy, in the form of fuel, to thermal energy and then to
 mechanical energy for use in propelling aircraft, generating electric
 power, pumping fluids etc. One of the primary approaches used to improve
 the efficiency of gas turbine engines is the use of higher operating
 temperatures. In the hottest portion of modern gas turbine engines (i.e.,
 the primary gas flow path within the engine turbine section), turbine
 airfoil components, cast from nickel or cobalt based alloys, are exposed
 to gas temperatures above their melting points. These components survive
 only because cooling air is passed through a cavity within the component.
 The cooling air circulates through this cavity reducing component
 temperature and exits the component through holes in the component, where
 it then mixes with the hot gasses contained within the primary flow path.
 However, providing cooling air reduces engine efficiency.
 Accordingly, there has been extensive development of coatings for gas
 turbine hardware. Historically, these coatings have been applied to
 improve oxidation or corrosion resistance of surfaces exposed to the
 turbine gas path. More recently, thermal barrier coating have been applied
 to internally cooled components exposed to the highest gas path
 temperatures so that the amount of cooling air required can be
 substantially reduced. Since coatings add weight to a part and debits
 fatigue life, application of the coating is intentionally limited to those
 portions of the component for which the coating is necessary to achieve
 the required durability. In the case of rotating parts such as turbine
 blades, the added weight of a coating adds significantly to blade pull,
 which in turn requires stronger and/or heavier disks, which in turn
 require stronger and/or heavier shafts, and so on. Thus there is added
 motivation to restrict use of coatings strictly to those portions of the
 blade, e.g., typically the primary gas path surfaces, where coatings are
 absolutely required.
 With increasing gas path temperatures, turbine components or portions of
 components that are not directly exposed to the primary turbine gas path
 may also exposed to relatively high temperatures during service, and
 therefore may also require protective coatings. For example, portions of a
 turbine blade that are not exposed to the gas path (such as the underside
 of the platform, the blade neck, and attachment serration) can be exposed
 to temperatures in excess of 1200 F. during service. These blade locations
 are defined at 18 and 19 in FIG. 1. It is expected that the temperatures
 these portions of the blade are exposed to will continue to increase as
 turbine operating temperatures increase.
 It is an object of the invention to provide a corrosion-resistant coating
 to prevent corrosion of components in regions not directly exposed to the
 hot gas stream.
 It is another object of the invention to provide a corrosion-resistant
 coating to prevent stress corrosion cracking on portions of turbine blades
 which are not directly exposed to a hot gas stream.
 It is yet another object of the invention to provide such a coating to
 protect against stress corrosion cracking of turbine blades in regions
 under the blade platform.
 SUMMARY OF THE INVENTION
 According to one aspect of the invention, improved durability of gas
 turbine blades is achieved through application of corrosion resistant
 coatings. A turbine blade for a gas turbine engine, typically consisting
 of a directionally solidified nickel-based superalloy, consists of an
 airfoil, a root and a platform located between the blade airfoil and root.
 The blade has a blade neck adjacent the blade root, and the platform has
 an underside adjacent the blade neck.
 In one aspect of this invention, a corrosion resistant overlay coating such
 as a stabilized zirconia is applied to the underside of the platform and
 portions of the blade neck, preferably by plasma spray. The presence of
 this coating improves component life by preventing blade corrosion by the
 salt accumulating on regions of the blade shielded from direct exposure to
 the gas path, e.g., underplatform surfaces. An additional benefit of the
 applied coating is the prevention of blade stress corrosion cracking. The
 corrosion resistant overlay coating prevents corrosion and/or stress
 corrosion cracking by acting as a barrier between the salt and
 nickel-based alloy component.
 In a more general application of the invention, the corrosion resistant
 overlay coating system may include an aluminide or platinum aluminide bond
 coat either between nickel alloy substrate and the MCrAlY layer or over
 the MCrAlY layer. The bond coat may be present to provide certain
 characteristics to the coated component. These characteristics may include
 more efficient blade repair/manufacture or improved durability.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS
 As illustrated in FIG. 1, a turbine blade composed of a superalloy material
 and incorporating the present invention is illustrated generally by the
 reference numeral 10. The turbine blade includes an airfoil 12, a serrated
 blade root 14 (used to attach the blade to the rotatable turbine disk) and
 a platform 16 located between the airfoil and serrated root. The region
 between the underside of the blade platform 18 and the root is referred to
 as the neck 19. Typically, turbine blades (and other gas turbine engine
 components) are composed of a directionally solidified nickel-based alloy,
 e.g., consisting of a single crystal or with multiple columnar grains
 oriented parallel to the direction of growth. Typical compositions of such
 alloys are shown in Table 1. Exemplary U.S. Patents describing columnar
 and single crystal and directionally solidified alloys include U.S. Pat.
 Nos. 4,209,348; 4,643,782; 4,719,080 and 5,068,084, each of which is
 expressly incorporated by reference herein. Cooling holes, which may be
 positioned on one or more portions of a turbine blade, may be provided for
 flowing cooling air over the specific portions of the airfoil during
 operation, as is known generally in the art.
 TABLE 1
 COMPOSITION OF COLUMNAR AND SINGLE CRYSTAL ALLOYS
 Alloy Type Ni Co Cr Al Mo Ta W Re Hf Ti
 Nb
 PWA 1422 DS Bal. 10 9 5 -- -- 12 -- 1.6 2
 1
 DS R80H DS Bal. 9.5 14 3 4 -- 4 -- 0.75
 4.8 --
 CM247LC DS Bal. 9.2 8.1 5.6 0.5 3.2 9.5 -- 1.4 0.7
 --
 PWA 1480 SC Bal. 5 10 5 -- 12 4 -- -- 1.5
 --
 PWA 1484 SC Bal. 10 5 5.65 1.9 8.7 5.9 3 0.1 --
 --
 Rene' N5 SC Bal. 7.5 7 6.2 1.5 6.5 5 3 0.15 --
 --
 CMSX-4 SC Bal. 9 6.5 5.6 0.6 6.5 6 3 0.1 1
 --
 It was discovered that the alkali and alkaline earth sulfate salts
 responsible for elevated temperature corrosion of turbine components
 (varying mixtures of sodium, potassium, calcium and magnesium sulfates)
 can accumulate on regions of the blade outside of the turbine gas path.
 These salts can be ingested with the inlet air in marine environments
 and/or form as a result of combustion processes. Corrosion attack of the
 blade by these salts is typically very limited at temperatures below the
 salt melting temperature (about 100 F.). With increased turbine operating
 temperature, however, blade regions shielded from the gas path can exceed
 the melting temperature of the sulfate salt resulting in accelerated
 corrosion of the blade neck and underside of the platform. It was also
 discovered that at sufficiently high stress levels, the presence of these
 salts may result in stress corrosion cracking of directionally solidified
 nickel-based turbine alloys having a single crystal or columnar grain
 structure. Stress corrosion cracking of these materials represents a newly
 discovered phenomenon.
 In one aspect of the current invention, a corrosion-resistant overlay
 coating (21 in FIG. 1a) is applied to portions of the substrate 20, such
 as the underside of the platform 18 and the neck 19 of a turbine blade to
 prevent corrosion and/or stress corrosion cracking of the blade in these
 locations. While the present invention is illustrated in FIG. 1 as a
 turbine blade, the present invention is not limited to any particular
 component. Other components exposed to relatively high stress and
 corrosive conditions would also be expected to benefit from this
 invention.
 Referring now to FIG. 1a, a corrosion inhibiting, ceramic coating is
 applied to a portion or portions of the article that are susceptible to
 corrosion and/or stress corrosion cracking. Using the turbine blade
 example, the coating is applied to the underside of the platform 18 and
 the neck 19 to prevent corrosion and/or stress corrosion cracking of the
 blade in these locations. While the present invention is illustrated in
 FIG. 1 as a turbine blade, the present invention is not intended to be
 limited to any particular component. Other components exposed to
 relatively high stress and corrosive conditions would also be expected to
 benefit from this invention. The overlay coating applied to the selected
 area(s), e.g., the under-platform surface 18 and neck 19 may be a
 conventional thermal barrier coating type material, such as a stabilized
 zirconia, e.g., 7YSZ, although the coating may also include other
 elements. We believe that other ceramic coatings may be employed with
 equal effect. See, e.g., commonly owned pending application Ser. Nos.
 09/164,700, filed on Oct. 1, 1998, and continuing prosecution application
 Ser. No. 08/764,419, filed on May 22, 1998, both entitled "Thermal Barrier
 Coating Systems and Materials" and expressly incorporated by reference
 herein.
 The coating is applied to the surface(s) to a thickness of at least about
 0.25 mils and up to about 5 mils. For rotating applications, such as
 turbine blades, the coating thickness should be adequate to ensure
 complete coverage of the area to be coated, e.g., no uncoated areas in the
 portions to be covered, and provide corrosion life necessary for providing
 protection for a typical blade service interval. The maximum coating
 thickness should be limited due to the fatigue debit associated with the
 additional weight of coatings, which typically add weight but not
 structural strength to the article. Accordingly, the thickness for
 rotating components is preferably less than about 3 mils, and more
 preferably about 2 mils.
 The coating may be applied by various processes, such as by vapor
 deposition or thermal spray. We prefer to use a plasma spray, which has
 been used previously to apply ceramic materials to other portions of gas
 turbine engine components such as the airfoil portions.
 FIG. 2 is an alternate embodiment of the present invention. Tile component
 includes an alumna forming coating between the ceramic layer and the
 substrate. For example, the component may include a substrate 20 as above,
 an alumina forming layer 22 such as an MCrAl type overlay coating or an
 aluminide layer. Both MCrAl coatings, which may include other elements
 such as Y, Hf, Si, Re and others, and aluminide coatings are known
 generally, and the particular MCrAl and aluminide compositions and methods
 of application need not be described here in detail. See, e.g., commonly
 owned U.S. Pat. No. Re 32,121 for a discussion of MCrAlY coatings and U.S.
 Pat. No. 5,514,482 for a description of aluminide coatings, both of which
 are expressly incorporated by reference herein. As is the case with the
 ceramic coating, the alumina forming layer adds weight but not structural
 strength to the article being coated, and should be no thicker than
 necessary. The ceramic, such as 7YSZ may then be applied by any suitable
 process, such as thermal spray or physical vapor deposition.
 The present invention provides significant improvement over the prior art.
 Applying a ceramic coating on selected portions of an article which are
 subjected to high temperatures and stresses, such as the under-platform
 surfaces of a turbine blade, provides superior corrosion and stress
 corrosion protection during operation. Moreover, there already exists
 significant experience in applying ceramic materials to other portions of
 these components, as well as to other components.
 While the present invention has been described above in some detail,
 numerous variations and substitutions may be made without departing from
 the spirit of the invention or the scope of the following claims.
 Accordingly, it is to be understood that the invention has been described
 by way of illustration and not by limitation.