Patent Document

BACKGROUND OF THE INVENTION 
     The invention relates generally to components of the hot section of gas turbine engines, and more particularly, to a process for depositing a coating onto a selective area of a turbine component. 
     In gas turbine engines, for example, aircraft engines, air is drawn into the front of the engine, compressed by a shaft-mounted rotary compressor, and mixed with fuel. The mixture is burned, and the hot exhaust gases are passed through a turbine coupled to a shaft. The flow of gas turns the turbine, which drives the compressor. The hot exhaust gases flow from the back of the engine, providing thrust that propels the aircraft forward. 
     During operation of gas turbine engines, at least some components within the engine, maybe in contact with high temperature gases. Such components may include, for example, blades, vanes, and nozzles used to direct the flow of the hot gases. 
     To facilitate shielding the metallic parts from the combustion gases, environmental coatings may be applied to the components. Such environmental coatings may be produced by holding the part to be coated at a temperature in an atmosphere that is rich in a certain element or elements, often aluminum. The elements diffuse onto the surface of the part and form a diffusion coating in a process known as diffusion aluminide. In one form, the environmental coating is fabricated from a diffusion cobalt aluminide, nickel aluminide or platinum aluminide. The diffusion aluminide coating surface forms an aluminum oxide scale when exposed to oxygen-containing atmospheres at elevated temperatures, thus facilitating increased resistance to additional high temperature oxidation. 
     At least some other known component coating processes demand labor-intensive processes. For example, when the component is a low pressure turbine (LPT) nozzle, known coating processes require a labor intensive masking process wherein a commercially available aluminum gettering masking tape is applied to the desired area of the turbine component. More specifically, the tape is affixed in place using a sheet metal strip. However, continued exposure to the high temperatures utilized by the coating process, may cause the sheet metal strip to warp, such that the strip fails to provide adequate support for the masking tape. As a result the masking tape may undesirably dislodge from the component during the aluminide coating process, and an undesired area of the turbine nozzle may be aluminided. 
     BRIEF DESCRIPTION OF THE INVENTION 
     In one aspect, a method for processing a substrate article is provided. The method includes masking a first portion of the substrate article with a maskant that includes a formed graphite piece that overlays and contacts the first portion of the substrate such that a second portion of the substrate is not overlaid nor contacted by the maskant; and processing the substrate article such that a coating of material is deposited on the second portion of the substrate, and wherein the maskant facilitates preventing the coating from being deposited on the first portion of the substrate article. 
     In another aspect, a method for coating a gas turbine engine turbine engine nozzle with an environmental coating is provided. The method includes masking a first portion of the turbine engine nozzle with a maskant including a formed graphite piece that overlies and contacts the first portion of the nozzle, such that a second portion of the nozzle remains exposed, and depositing a coating on the second portion of the nozzle without removing the maskant such that the maskant facilitates preventing the coating from being deposited on the first portion of the nozzle. 
     In yet another aspect, a coating mask for use in coating a substrate is provided. The mask includes a first interlocking segment end, an opposite second interlocking segment end and a body extending therebetween, each interlocking segment end is configured to interlock with a respective interlocking end of an adjacent segment such that a plurality of interlocking segments overlay a substrate article first portion. The body includes a formed graphite mask segment configured to isolate a portion of the substrate article from a coating atmosphere, and a contour surface shaped to conform to a first portion of the substrate article. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a cross-sectional side view of an exemplary gas turbine engine; 
         FIG. 2  is a perspective view of an exemplary low pressure turbine nozzle that may be used with a gas turbine engine, such as the gas turbine engine shown in  FIG. 1 ; 
         FIG. 3  is a perspective view of an exemplary inner mask segment that may be used with the nozzle shown in  FIG. 2 ; 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     Nickel-base superalloy components of gas turbines are sometimes coated with aluminum and then heated to diffuse the aluminum into the surface of the article. The aluminum-rich surface is oxidized to produce an adherent aluminum oxide scale on the surface of the article. The aluminum oxide scale is an effective barrier against further oxidation and corrosion of the component in service. 
     The aluminum coating is typically applied by a vapor phase deposition process. In one embodiment, aluminum containing a cobalt-aluminum donor alloy and a halide activator, such as aluminum fluoride gas, is contacted to the component surface under conditions such that the compound decomposes to leave a layer of aluminum deposited on the surface. The aluminum diffuses into the surface during the deposition and any post-deposition heat treatment, producing the aluminum-enriched surface region. 
     It is sometimes the case in such deposition processes that a first portion of the surface of the article is to be left uncoated, and a second portion of the surface of the article is to be coated with aluminum. In order to prevent deposition of aluminum from the aluminum-containing gas, the first (uncoated) portion of the surface of the article is physically covered with a mask. The mask prevents contact of the aluminum-containing gas to the first portion. These maskants are intended to prevent the coating vapors from reaching the surface of the article, and to prevent depletion of the alloy components from the surface of the first portion of the surface. 
       FIG. 1  is a cross-sectional side view of an exemplary gas turbine engine  10 . In one embodiment, engine  10  is an F110/129 engine available from General Electric Aircraft Engines, Cincinnati, Ohio. Engine  10  has a generally longitudinally extending axis or centerline  14  extending in a forward direction  16  and an aft direction  18 . Engine  10  includes a core engine  30  which includes a high pressure compressor  34 , a combustor  36 , a high pressure turbine  38 , and a power turbine or a low pressure turbine  39  all arranged in a serial, axial flow relationship. In an alternative embodiment, core engine  30  includes a compressor, a detonation chamber, and a turbine arranged in a serial, axial flow relationship. Engine  10  also includes a bypass duct  44  that surrounds core engine  30 , and enables fluid flow to be routed downstream from core engine  30  rather than through core engine  30 . In an alternative embodiment, engine  10  includes a core fan assembly (not shown). An annular centerbody  50  extends downstream from core engine  30  toward a variable geometry exhaust nozzle  54 . 
     During operation, airflow enters engine  10  and fuel is introduced to core engine  30 . The air and fuel are mixed and ignited within core engine  30  to generate hot combustion gases. Specifically, pressurized air from high pressure compressor  34  is mixed with fuel in combustor  36  and ignited, thereby generating combustion gases. Such combustion gases drive high pressure turbine  38  which drives high pressure compressor  34 . The combustion gases are discharged from high pressure turbine  38  into low pressure turbine  39 . The core airflow is discharged from low pressure turbine  39  and directed aftward towards exhaust nozzle  54 . 
       FIG. 2  is a perspective view of an exemplary low pressure turbine nozzle  200  that may be used with a gas turbine engine, such as gas turbine engine  10  (shown in  FIG. 1 ). Nozzle  200  includes a plurality of circumferentially-spaced airfoil vanes  202  coupled together by an arcuate radially outer band or platform  204 , and an arcuate radially inner band or platform  206 . More specifically, in the exemplary embodiment, each band  204  and  206  is integrally-formed with airfoil vanes  202 . 
     In the exemplary embodiment, each airfoil vane  202  includes a first sidewall  208  and a second sidewall  210 . First sidewall  208  is convex and defines a suction side of each airfoil vane  202 , and second sidewall  210  is concave and defines a pressure side of each airfoil vane  202 . Second sidewall  210  is joined to first sidewall  208  at a leading edge  212  and at an axially-spaced trailing edge (not shown) of each airfoil vane  202 . More specifically, each airfoil trailing edge is spaced chordwise and downstream from each respective airfoil leading edge  212 . 
     Second sidewall  210  and first sidewall  208  extend longitudinally, or radially outwardly, in span from radially inner band  206  to radially outer band  204 . Additionally, second sidewall  210  and first sidewall  208  define a cooling cavity (not shown) within each airfoil vane  202 . More specifically, the cooling cavity is bounded by an inner surface (not shown) of each airfoil sidewall, and extends through each band  204  and  206 . 
     In the exemplary embodiment, nozzle  200  is fabricated from a nickel-base superalloy. “Nickel-base” as used herein means that the alloy contains more nickel by weight than any other element, for example, but not limited to, nickel-base superalloy, Rene  80 . In alternative embodiments, other materials such as iron-base, cobalt-base or titanium-base alloys may be used. 
     Nozzle  200  may be of any operable shape, such as, for example, a gas turbine blade, a gas turbine vane, a gas turbine nozzle, a piece of tubing, a tool shape, a pump impeller, a pump rotor, a fan blade, or an element of electronic hardware. Nozzle  200  may be prepared by any operable approach known in the art, such as casting or forging. Nozzle  200  may be furnished in substantially its final shape and dimensions as the aluminide coating is thin and adds little to the dimensions of the article. In some cases, the article may instead be furnished slightly undersized to account for the thickness of the applied coating. A first portion of nozzle  200  may be masked with a second portion of nozzle  200  unmasked and exposed. 
     In the exemplary embodiment, nozzle  200  is illustrated partially masked for a aluminide coating process. An arcuate band of a plurality of outer maskant segments  214  overlay and are in contact with an outer periphery of outer band  204 . Each outer maskant segment  214  includes a first circumferential end  216  and a second, opposite circumferential end  218 . Each end  216  and  218  includes an interlocking tab  220  and an interlocking recess  222 . In the exemplary embodiment, a radially inner surface (not shown) of segment  214  is machined to conform dimensionally to a radial outer surface (not shown) of outer band  204 . In an alternative embodiment, the radially inner surface of segment  214  is molded to conform to the radially outer surface of outer band  204 . 
     An arcuate band of inner maskant segments  224  overlay and contact an inner periphery  226  of inner band  206 . Each inner maskant segment  224  includes a first circumferential end  228  and a second, opposite circumferential end  230 . Each end  228  and  230  includes an interlocking tab  232  and an interlocking recess  234 . In the exemplary embodiment, a radially outer surface  236  of segment  224  is machined to conform dimensionally to a radially inner surface  226  of inner band  206 . In an alternative embodiment, the radially outer surface  236  of segment  224  is molded to conform to the radially inner surface  226  of inner band  206 . 
     During the coating process, segments  214  and  224  are assembled to overlay and contact outer band  204  and inner band  206  respectively. Each surface of segments  214  and  224  that contacts inner and outer band  204  and  206  respectively is formed to conform to band surface to facilitate preventing the coating atmosphere from contacting the portions of the bands  204  and  206  that are in contact with mask segments  214  and  224 . The machined surfaces of the segments that conform to the surfaces of bands  214  and  224  obviate the need to seal the edges of the contact surfaces to facilitate preventing coating atmosphere from reaching masked surfaces of band  204  and  206 . Interlocking tabs  220  and recesses  222  of segments are engaged to facilitate providing lateral support to each adjacent segment and to provide a torturous path past ends  216  and  218 . After coating and/or diffusion, the coated nozzle  200  may be cooled, and segments  214  and  224  may be removed and later reused. 
     Each graphite piece may be formed by molding or extruding and may be machined from a monolithic block of graphite that is formed in any manner as is known in the art. 
       FIG. 3  is a perspective view of an exemplary inner mask segment  224  that may be used with nozzle  200  (shown in  FIG. 2 ). Segment  224  includes first end  228  and second end  230 . Each end  228  and  230  includes interlocking tab  232  and interlocking recess  234 . Tab  232  and recess  234  are configured to engage and interlock with a tab and recess on each end of adjacent segments. Radially outer surface  236  of segment  224  is formed to conform to a respective mating face on inner band  206  (shown in  FIG. 2 ). 
     Although segments  214  and  224  are illustrated in association with a process for masking a turbine nozzle, it should be understood that the methods and apparatus described above may be used to mask articles of shapes and orientations different than those describe herein. It is anticipated that freestanding and/or formed graphite maskants provide benefits that would accrue to articles of various shapes and orientations. 
     The above-described methods and systems for applying diffusion aluminide coating on a selective area of a turbine engine component is cost-effective and highly reliable for facilitating coating a portion of a component where a coating is desired and for facilitating preventing the coating atmosphere from contacting a portion of the component where a coating is not desired. Specifically, the freestanding, dimensionally stable mask segments are reusable and easily handled and positioned to protect the portion desired to be free of coating. As a result, the methods and apparatus described herein facilitate fabrication and maintenance of components in a cost-effective and reliable manner. 
     Exemplary embodiments of combinations of gas turbine engine components and coating masks are described above in detail. The combinations are not limited to the specific embodiments described herein, but rather, components of each combination may be utilized independently and separately from other components described herein. Each combination component can also be used in combination with other system components. 
     While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.

Technology Category: 2