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 . fig1 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 . fig2 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 fig1 ). 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 . fig3 is a perspective view of an exemplary inner mask segment 224 that may be used with nozzle 200 ( shown in fig2 ). 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 fig2 ). 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 .