Patent Document

FIELD OF THE INVENTION 
     The present invention relates to seal structure for preventing leakage of gases across a gap between first and second components in a turbine engine. 
     BACKGROUND OF THE INVENTION 
     U.S. Pat. No. 5,934,687 discloses a gas-path leakage seal for sealing a gap between first and second members of a gas turbine. The seal comprises a flexible metal sheet assembly, and first and second cloth layer assemblies. 
     U.S. Pat. No. 6,733,234 discloses a gas-path leakage seal assembly for sealing a gap between turbine members comprising a shim and a spring for urging the shim into contact with the turbine members. The shim may comprise a protection material for contacting the turbine components. 
     SUMMARY OF THE INVENTION 
     In accordance with a first aspect of the present invention, a seal structure is provided for preventing leakage of gases across a gap between first and second components in a turbine engine. The seal structure is adapted to be received in first and second slots provided in the first and second components. The seal structure may comprise: a wear resistant layer; and a deformable layer defined by a material having one of a varying density and a varying porosity. 
     The seal structure may further comprise a core layer positioned between the wear resistant layer and the deformable layer. The core layer may comprise a metal core layer. 
     The wear resistant layer may be formed from one of a metal powder and a ceramic powder. Preferably, the wear resistant layer is slightly harder than the first and second turbine engine components. 
     The deformable layer may be formed from one of a metal powder and a ceramic powder. Preferably, the deformable layer is softer than the first and second turbine engine components. 
     The deformable layer includes an outer surface and an inner surface and may comprise a density which increases from the outer surface to the inner surface. 
     The deformable layer includes an outer surface and an inner surface and may comprise a porosity which decreases from the outer surface to the inner surface. 
     In accordance with a second aspect of the present invention, a turbine engine is provided comprising a first component having a first slot; a second component having a second slot; and a seal structure. The second component is positioned adjacent to the first component such that the first and second slots are positioned generally opposed to one another. The seal structure is provided in the first and second slots for preventing leakage of gases across a gap between the first and second components. The seal structure comprises a wear resistant layer, and a deformable layer defined by a material having one of a varying density and a varying porosity. The seal structure may further comprise a core layer positioned between the wear resistant layer and the deformable layer. 
     The first slot may be defined by first and second inner surfaces of the first component and the second slot may be defined by third and fourth inner surfaces of the second component. The second and fourth inner surfaces may have surface imperfections. 
     The second and fourth surfaces may have a surface roughness Ra falling within a range of from about 0.8 micrometers to about 12.5 micrometers. 
     The deformable layer may contact the second and fourth inner surfaces of the first and second components and permanently deform during operation of the engine so as to correspond in shape to the surface imperfections on the second and fourth inner surfaces. 
     The deformable layer may be exposed to hot working gases and the wear resistant layer may be exposed to cooling gases. The cooling gases may have a pressure greater than that of the hot working gases. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a perspective view of first and second vanes with a seal structure constructed in accordance with the present invention; 
         FIG. 2  is an enlarged view of a portion of the first and second vanes and the seal structure illustrated in  FIG. 1 ; 
         FIG. 3  is an enlarged view of a recess provided in the second vane with a seal structure constructed in accordance with a first embodiment of the present invention just after it has been inserted into the recess; 
         FIG. 4  is a view similar to  FIG. 3 , but after the seal has been in the recess for some period of time; and 
         FIG. 5  is an enlarged view of a recess provided in the second vane with a seal structure constructed in accordance with a second embodiment of the present invention just after it has been inserted into the recess. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     The embodiments of the present invention provide a gas-path leakage seal structure for use in a turbine engine. 
       FIG. 1  illustrates first and second turbine engine components comprising first and second adjacent stationary vanes  10  and  12 . The first vane  10  comprises a first airfoil  10 A and a first platform  10 B. The second vane  12  comprises a second airfoil  12 A and a second platform  12 B. The vane airfoils  10 A and  12 A function to guide hot combustion gases to rotatable blades (not shown) coupled to a rotor to effect rotation of the rotor. As is apparent from  FIGS. 1 and 2 , the first and second vane platforms  10 B and  12 B are positioned adjacent to one another. 
     In accordance with a first embodiment of the present invention, a seal structure  20  is provided between the adjacent first and second vane platforms  10 B and  12 B to seal a gap G between the first and second platforms  10 B and  12 B, see  FIGS. 1-4 . The first platform  10 A is provided with first and second circumferentially spaced apart slots  10 C and  10 D and the second platform  12 B is provided with third and fourth circumferentially spaced apart slots  12 C and  12 D. The second and third slots  10 D and  12 C are adjacent to one another and are open to the gap G, see  FIGS. 1 and 2 . The seal structure  20  fits into the second and third slots  10 D and  12 C and spans across the gap G so as to seal the gap G to prevent the hot working gases moving past the vane airfoils  10 B and  12 B from passing through the gap G. The seal structure  20  also prevents cooling gases or air exposed to lower surfaces  100 A and  120 A of the platforms  10 B and  12 B from passing through gap G. 
     It is also contemplated that the seal structure  20  may be used to seal gaps between other turbine engine components such as blades and ring segments (not shown). 
     The first and second vanes  10  and  12  may be formed from a metal alloy via a casting operation. The first, second, third and fourth slots  10 C,  10 D,  12 C and  12 D in the vane platforms  10 B and  12 B may be formed via a conventional electro-discharge machining (also referred to as electric discharge machining) operation. The second slot  10 D is defined by first and second inner surfaces  100 C and  100 D in the first vane platform  10 A and the third slot  12 C is defined by third and fourth inner surfaces  120 C and  120 D in the second vane platform  12 B, see  FIG. 2 . The first, second, third and fourth inner surfaces  100 C,  100 D,  120 C and  120 D of the first and second vane platforms  10 B and  12 B, because they are formed via an electro-discharge machining operation, have irregular surfaces S I  or non-smooth topologies, see  FIG. 3 , which is an enlarged schematic view of portions of the third and fourth surfaces  120 C and  120 D in the second vane platform  12 A. The inner surfaces  100 C,  100 D,  120 C and  120 D my have a surface roughness Ra falling within a range of from about 0.8 micrometer to about 12.5 micrometers. 
     In a first embodiment illustrated in  FIGS. 2-4 , the seal structure  20  comprises a wear resistant layer  22 , a core layer  24  and a deformable layer  26 , wherein the core layer  24  is positioned between the wear resistant layer  22  and the deformable layer  26 . In the illustrated embodiment, the wear resistant layer  22  is positioned adjacent to the first and third surfaces  100 C and  120 C of the first and second vane platforms  10 B and  12 B. Hence, the wear resistant layer  22  is exposed to cooling gases, which cooling gases also contact the lower surfaces  100 A and  120 A of the platforms  10 B and  12 B, as noted above. Since the wear resistant layer  22  is preferably harder than the first and third surfaces  100 C and  120 C of the first and second vane platforms  1013  and  12 B, the wear resistant layer  22  will experience minimal wear during turbine engine operation. 
     The wear resistant layer  22  may be formed via a conventional laser cladding operation from one of a metal powder, e.g., nickel alloys, and a ceramic powder. Such a laser cladding operation may involve injecting a metal or ceramic powder towards a laser beam, such that the laser beam melts the powder, which melted powder is then deposited onto a substrate, i.e., the core layer  24 . Preferably, the wear resistant layer  22  is slightly harder than the first and second vane platforms  10 B and  12 B. Hardness of the wear resistant layer  22  can be defined by selecting a metal powder or ceramic powder having a desired hardness, which, preferably, exceeds that of the first and second vane platforms  10 B and  12 B. 
     The core layer  24  may be formed from a metal such as a Nickel or Cobalt based Alloy and functions to provide load carrying strength and/or provide a spring function to the seal structure  20 . 
     In the illustrated embodiment, the deformable layer  26  is positioned adjacent to the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B. Hence, the deformable layer  26  is exposed to the hot working gases, which hot gases also contact the airfoils  10 A and  12 A, as noted above. The deformable layer  26  may also be formed via a conventional laser cladding operation from one of a metal powder, e.g., nickel alloys, and a ceramic powder. Preferably, the deformable layer  26  is softer, i.e., less hard, than the first and second vane platforms  10 B and  12 B. Softness/hardness of the deformable layer  26  can be selected based on the softness/hardness of the metal powder or ceramic powder used in forming the deformable layer  26 . Softness/hardness can also be varied based on the density of the deformable layer  26 , which density can be varied with metal or ceramic powder feed rate as well as by selecting an appropriate laser power. For example, as laser power is decreased, the resulting layer may comprise less densely packed powder particles with more voids between the powder particles, thereby resulting in a less hard and/or more deformable layer  26 . Softness/hardness can further be varied based on porosity of the deformable layer  26 , which porosity can be varied based on metal or ceramic powder particle size and/or laser power. For example, as laser power is decreased, the resulting layer may comprise less densely packed powder particles with more voids between the powder particles. 
     Preferably, the deformable layer  26  includes an outer surface  260 A, near the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B, and an inner surface  260 B, adjacent the core layer  24 , see  FIG. 3 . The deformable layer  26  preferably comprises a density which increases gradually from the outer surface  260 A to the inner surface  260 B. Alternatively, the deformable layer  26  may comprise a porosity which decreases gradually from the outer surface  260 A to the inner surface  260 B.  FIG. 3  schematically illustrates the seal structure  20  just after it is first inserted into the second and third slots  10 D and  12 C in the vane platforms  10 B and  12 B. 
     During operation of the engine turbine, the cooling gases have a greater pressure than that of the hot working gases. Hence, the cooling gases apply a force on the wear resistant layer  22  so as to force the deformable layer  26  against the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B. Hence, the deformable layer  26  may permanently deform, i.e., powder or metal particles of the deformable layer  26  may break off from adjacent particles, such that the layer  26  corresponds in shape to the surface imperfections on the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B. Because the deformable layer  26  conforms to the irregular surfaces S I  of the second and fourth surfaces  100 D and  120 D, an enhanced seal is made between the seal structure  20  and the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B so as to limit or minimize leakage of hot working gases and/or cooling gases through the gap G. 
     In a second embodiment illustrated in  FIG. 5 , the seal structure  20 ′ comprises a wear resistant layer  22 ′ and a deformable layer  26 ′. No metal core layer is provided in this embodiment. The wear resistant and deformable layers  22 ′ and  26 ′ may be formed in the same manner as the wear resistant and deformable layers  22  and  26  illustrated in  FIGS. 3 and 4 . During operation of the engine turbine, the cooling gases apply a force on the wear resistant layer  22 ′ so as to force the deformable layer  26 ′ against the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B. Hence, the deformable layer  26 ′ may permanently deform, i.e., powder or metal particles of the deformable layer  26 ′ may break off from adjacent particles, such that the layer  26 ′ corresponds in shape to the surface imperfections on the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B. Because the deformable layer  26 ′ conforms to the irregular surfaces S I  of the second and fourth surfaces  100 D and  120 D, an enhanced seal is made between the seal structure  20 ′ and the second and fourth surfaces  100 D and  120 D of the first and second vane platforms  10 B and  12 B so as to limit or minimize leakage of hot working gases and/or cooling gases through the gap G. 
     While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.

Technology Category: f