Patent Document

CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    Not applicable. 
       TECHNICAL FIELD 
       [0002]    The present invention relates to gas turbine engines. More particularly, embodiments of the present invention relate to a gas turbine vane having a platform shaped in order to reduce ingestion of hot combustion gases into joints between adjacent vanes of a vane assembly. 
       BACKGROUND OF THE INVENTION 
       [0003]    A gas turbine engine operates to produce mechanical work or thrust. For a land-based gas turbine engine, a generator is typically coupled to the engine through an axial shaft, such that the mechanical work of the engine is harnessed to generate electricity. A typical gas turbine engine comprises a compressor, at least one combustor, and a turbine, with the compressor and turbine coupled together through the axial shaft. In operation, as air passes through multiple stages of axially-spaced rotating blades and stationary vanes of the compressor, its pressure increases. The compressed air is then mixed with fuel in the combustion section, which can comprise one or more combustion chambers. The fuel-air mixture is ignited in the combustion chamber(s), producing hot combustion gases, which pass into the turbine causing the turbine to rotate. The turning of the shaft also drives the generator. 
         [0004]    A prior art turbine vane  100  is shown in  FIG. 1 . The turbine vane  100  includes an inner platform  102 , an outer platform  104  spaced a distance radially outward relative to an engine centerline. Positioned between and connected to the platforms  102  and  104  is at least one airfoil  106 . In operation hot combustion gases pass through the channels created between the airfoils  106 . 
         [0005]    The turbine comprises a plurality of rotating and stationary stages of airfoils. For the turbine vanes, the leading edge region of the airfoil and vane platform is subjected to the aerodynamic loads from the preceding stage of turbine blades or the exit flow of a combustor. The combustion gases then pass around the airfoil, beginning at the airfoil&#39;s leading edge. Depending on the shape of the airfoil and the angle at which the flow of hot gases are imparted onto the leading edge of the airfoil, a bow wave can be created, which is an area of high pressure combustion gases extending a distance away from the airfoil leading edge. This wave of combustion gases is often forced into the region between adjacent turbine vanes in a vane assembly. Depending on the supply pressure of the cooling air within the platform region and the strength of the bow wave, the hot combustion gases of the bow wave may penetrate into the joint between adjacent vanes, causing overheating and erosion of the platform. 
       SUMMARY 
       [0006]    Embodiments of the present invention are directed towards gas turbine vanes and a gas turbine vane assembly. In an embodiment of the present invention, a gas turbine vane comprises an inner arc-shaped platform, an outer arc-shaped platform, and an airfoil extending therebetween. The inner arc-shaped platform has a pressure side radial face and a suction side radial face where the pressure side radial face is formed in two intersecting portions and includes a relief cut at the intersection of the two portions. The outer arc-shaped platform, which is spaced a radial distance from the inner platform also has a pressure side radial face and suction side radial face where the pressure side radial face is also formed having two intersecting portions and includes a relief cut at the intersection of the two portions. The outer arc-shaped platform is separated from the inner arc-shaped platform by at least one airfoil. The suction side radial faces each have a generally planar wall. 
         [0007]    In an alternate embodiment of the present invention, a gas turbine vane comprises an inner arc-shaped platform, an outer arc-shaped platform, and an airfoil extending between. The inner arc-shaped platform has a pressure side radial face and a suction side radial face where the suction side radial face is formed in two intersecting portions. The outer arc-shaped platform, which is spaced a radial distance from the inner platform also has a pressure side radial face and suction side radial face where the suction side radial face is also formed in two intersecting portions. The outer arc-shaped platform is spaced radially from the inner arc-shaped platform by at least one airfoil. 
         [0008]    In yet another alternate embodiment of the present invention a gas turbine vane assembly is disclosed comprising a first vane assembly, a second vane assembly, and a fastener mechanism. The first vane assembly has a first inner platform with a pressure side radial face having a first portion and a second portion, a first outer platform with a pressure side radial face also having a first portion and second portion, and a first airfoil extending between the first inner platform and first outer platform. The second vane assembly has a second inner platform with a suction side radial face having a first portion and a second portion, a second outer platform with a suction side radial face also having a first portion and second portion. The first vane assembly and second vane assembly are fastened together along the surfaces opposite of the multi-surface platform faces by a fastener mechanism. 
         [0009]    Additional advantages and features of the present invention will be set forth in part in a description which follows, and in part will become apparent to those skilled in the art upon examination of the following, or may be learned from practice of the invention. 
     
    
     
       BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS 
         [0010]    The present invention is described in detail below with reference to the attached drawing figures, wherein: 
           [0011]      FIG. 1  depicts a perspective view of a gas turbine vane assembly of the prior art; 
           [0012]      FIG. 2  depicts a perspective view of a gas turbine vane assembly in accordance with an embodiment of the present invention; 
           [0013]      FIG. 3  depicts a perspective view of the inner platform of the gas turbine vane assembly of  FIG. 2  in accordance with an embodiment of the present invention; 
           [0014]      FIG. 4  depicts a perspective view of components forming a gas turbine vane assembly in accordance with an embodiment of the present invention; 
           [0015]      FIG. 5  depicts a perspective view of the outer platforms of a gas turbine vane assembly in accordance with an embodiment of the present invention; 
           [0016]      FIG. 6  depicts a top view of a gas turbine vane assembly in accordance with an embodiment of the present invention; 
           [0017]      FIG. 7  depicts a graph of the gap pressure between the inner platforms of vanes of a gas turbine vane assembly of the prior art and an embodiment of the present invention; 
           [0018]      FIG. 8  depicts a graph of the gap pressure between the outer platforms of vanes of a gas turbine vane assembly of the prior art and an embodiment of the present invention; and 
           [0019]      FIG. 9  depicts a cross section view through a portion of the turbine vane assembly of the prior art within a turbine section showing the pressure isolines across the axial length of the turbine vane assembly. 
       
    
    
     DETAILED DESCRIPTION 
       [0020]    The subject matter of the present invention is described with specificity herein to meet statutory requirements. However, the description itself is not intended to limit the scope of this patent. Rather, the inventors have contemplated that the claimed subject matter might also be embodied in other ways, to include different components, combinations of components, steps, or combinations of steps similar to the ones described in this document, in conjunction with other present or future technologies. 
         [0021]    A gas turbine vane assembly  200  in accordance with an embodiment of the present invention is depicted in  FIGS. 2-6 . Referring to  FIG. 2 , the gas turbine vane assembly  200  comprises an inner arc-shaped platform  202  having a pressure side radial face  204  and a suction side radial face  206 . The pressure side radial face comprises a first portion  204 A, a second portion  204 B, and a relief cut  208  located at the intersection of the first portion  204 A and the second portion  204 B. 
         [0022]    The gas turbine vane assembly  200  further comprises an outer arc-shaped platform  210  spaced a distance radially outward of the inner arc-shaped platform  202 . The outer arc-shaped platform  210  has a pressure side radial face  212  and a suction side radial face  214 . The pressure side radial face  212  comprises a first portion  212 A, a second portion  212 B and a relief cut  216  at the intersection of the first portion  212 A and second portion  212 B. The gas turbine vane assembly  200  also comprises at least one airfoil  218  extending between the inner arc-shaped platform  202  and the outer arc-shaped platform  210 . Although a variety of manufacturing techniques can be used, for ease of manufacturing and structural integrity, it is preferred that the inner arc-shaped platform, airfoil, and outer arc-shaped platform are integrally cast together. 
         [0023]    The first portion  204 A of the inner arc-shaped platform  202  is generally co-planar with the first portion  212 A of the outer arc-shaped platform  210 . Further, the second portion  204 B of the inner arc-shaped platform  202  is also generally co-planar with the second portion  212 B of the outer arc-shaped platform  210 . Alignment of these surfaces is necessary to aid in assembly of the gas turbine vane assembly  200 , as discussed below. 
         [0024]    When the gas turbine vanes are assembled together in the turbine along their corresponding chevron portions, it is necessary to place one or more seals between adjacent platforms of the turbine vanes in order to prevent leakage between adjacent vanes. Referring again to  FIG. 2 , in an embodiment of the present invention, the first portion  204 A and the second portion  204 B of the inner arc-shaped platform  204  further comprises an inner seal slot  220 . In this embodiment, the first portion  212 A and second portion  212 B of the outer platform  210  also comprises an outer seal slot  222 . Therefore, one or more seals (not shown) can be placed in the slots  220  and  222  to seal the pressure side of the inner and outer platforms against an adjacent turbine vane. A variety of seal materials can be used, but one such material is a sheet metal seal, such as that disclosed in U.S. Pat. No. 7,334,800. 
         [0025]    Because of the extreme operating temperatures to which the turbine vane  200  is exposed, it is often necessary to provide additional measures to help protect the turbine vane. Therefore, an embodiment of the invention includes applying a thermal barrier coating to the gas path surfaces of the inner arc-shaped platform, the outer arc-shaped platform, and at least one airfoil. Also, in an embodiment of the invention, the vane assembly may be actively cooled by directing an air source to the airfoil  218  through the outer arc-shaped platform  210 . 
         [0026]    Referring to  FIGS. 2 ,  3 , and  5 , an alternate embodiment of the present invention is depicted. In the alternate embodiment, a gas turbine vane  400  comprises an inner arc-shaped platform  402  having a pressure side radial face  404  and a suction side radial face  406 , where the suction side radial face has a first portion  406 A and a second portion  406 B. The gas turbine vane  400  also comprises an outer arc-shaped platform  408  spaced a distance radially outward of the inner arc-shaped platform  402 . The outer arc-shaped platform  408  has a pressure side radial face  410  and a suction side radial face  412  where the suction side radial face  412  comprises a first portion  412 A and a second portion  412 B. At least one airfoil  414  extends between the inner arc-shaped platform  402  and the outer arc-shaped platform  408 . The inner arc-shaped platform  402 , outer arc-shaped platform  408 , and airfoil  414  are preferably integrally cast together. 
         [0027]    The first portion  406 A of the inner arc-shaped platform  402  is generally co-planar with the first portion  412 A of the outer arc-shaped platform  408 . While the second portion  406 B of the inner arc-shaped platform  402  is generally parallel to the second portion  412 B of the outer arc-shaped platform  408 . Furthermore, the first portion  406 A and the second portion  406 B of the inner arc-shaped platform  402  further comprises an inner seal slot  416 . Also, the first portion  412 A and second portion  412 B of the outer arc-shaped platform  412  further comprises an outer seal slot  418 . Similar to the first vane assembly  200 , one or more sheet metal seals can be placed in slots  416  and  418  to reduce leakage along the platform sidefaces between the first and second portions of adjacent radial faces. 
         [0028]    The gas turbine vane  400  also includes one or more alternatives for improving the thermal capability of the vane. One such alternative is a bond coating and thermal barrier coating. The bond coating and thermal barrier coating is applied to a portion of the inner arc-shaped platform, a portion of the outer arc-shaped platform and the at least one airfoil extending between the platforms. An additional way of improving thermal capability is through active cooling. The gas turbine vane  400  also comprises an airfoil  414  that is air cooled by a source of air entering the airfoil  414  through the outer arc-shaped platform  408  and passing along the walls of the airfoil and then through a plurality of openings  420  (see  FIGS. 3-5 ). 
         [0029]    A common prior art vane assembly configuration includes two parallel mate face surfaces, often times cut along an angle relative to the vane platform leading face, as depicted by A in  FIG. 9 . Depending on manufacturing and assembly tolerances, a small gap may be present between the adjacent platforms, thereby providing a way for hot combustion gases to enter the gap region. When this occurs, the hot gases, often upwards of approximately 2000 deg. F. can cause overheating and erosion to the platform leading edge region. One instance where this is known to occur, as shown in  FIG. 9 , is when there is a bow wave BW of hot gases extending away from the airfoil leading edge region  218 A and the airfoil leading edge  218 A is close enough to the vane platform edge A that because of the bow wave BW high pressure, it can enter the gap between adjacent platforms. As shown in  FIG. 9 , in the vane assembly the BW is shown coming off the leading edge region of each airfoil. 
         [0030]    In yet another embodiment of the present invention, a gas turbine vane assembly is disclosed. The gas turbine vane assembly  500  is shown in detail in  FIGS. 2-6 . The assembly  500  comprises a first vane assembly  200 , a second vane assembly  400 , and a fastener mechanism  600 . 
         [0031]    Through an embodiment of the present invention, where the first vane assembly  200  is secured to the second vane assembly  400  so as to form the gas turbine vane assembly  500 , significant improvements in eliminating injection of the bow wave gases is achieved, resulting in extended component life of the vane assembly  500 . Referring back to  FIG. 2 , the gas turbine vane assembly  500  comprises a first vane assembly  200  having a first inner platform  202  having a pressure side radial face  204  with first portion  204 A, second portion  204 B, and a relief cut  208  at their intersection. Inner platform  202  also includes a suction side radial face  206 , which is a generally straight surface. The first vane assembly  200  also comprises a first outer platform  210  that is spaced a radial distance from the first inner platform  202  and having a pressure side radial face  212  and a suction side radial face  214 , where the pressure side radial face  212  has a first portion  212 A and a second portion  212 B joined together at their intersection by a relief cut  216 . A first airfoil  218  extends between and joins together the first inner platform  202  and first outer platform  210 . 
         [0032]    As it can be seen from  FIGS. 2 ,  3 , and  5 , referring to the first vane assembly  200 , the first portion  204 A of the inner arc-shaped platform  202  is generally co-planar with the first portion  212 A of the outer arc-shaped platform  210 . Furthermore, the second portion  204 B of the inner arc-shaped platform  202  is also generally co-planar with the second portion  212 B of the outer arc-shaped platform  210 . Referring to  FIGS. 3 and 5 , a similar co-planar orientation of inner and outer platform mate face surfaces also exists for the suction side of the vane assemblies. 
         [0033]    The gas turbine vane assembly  500  further comprises a second vane assembly  400  that is secured to the first vane assembly  200 . Referring to  FIG. 3 , the second vane assembly  400  comprises a second inner platform  402  having a pressure side radial face  404  and a suction side radial face  406 , where the suction side radial face  406  has a first portion  406 A and a second portion  406 B. The second vane assembly  400  also comprises a second outer platform  408 , which is spaced a radial distance from the second inner platform  402  and also has a pressure side radial face  410  and a suction side radial face  412 , where the suction side radial face  412  has a first portion  412 A and a second portion  412 B. The second vane assembly  400  also includes a second airfoil  414  extending between the second inner platform  402  and the second outer platform  408 . 
         [0034]    The first vane assembly  200  and second vane assembly  400  is secured together by a fastener mechanism proximate the inner and outer platforms, as shown in  FIGS. 2 ,  4 , and  5 . An outer fastening mechanism  600  includes a first bracket  602  secured to a first vane assembly  200  and a second bracket  604  secured to a second vane assembly  400 . The first bracket  602  and second bracket  604  are held together by a removable fastener such as a pin or bolt (not shown). Referring to  FIG. 3 , an inner fastening mechanism  606  is shown and includes brackets  608  and  610  Like the outer fastening mechanism  600 , the inner fastening mechanism  606  is also utilizes removable fasteners, such as a pin or bolt (also not shown) for securing the brackets  608  and  610  together. Because of the vane assembly orientation, the outer fastening mechanism  600  utilizes three fasteners, where the inner fastening mechanism  606  utilizes one fastener. 
         [0035]    The vane assembly  500  is oriented such that the suction side radial face  214  of the outer platform  210  of the first vane assembly  200  is adjacent to the pressure side radial face  410  of the outer platform  408  of second vane assembly  400 , as shown in  FIG. 2 . The fastener mechanism  600  secures the first vane assembly  200  to the second vane assembly  400  so as to form a sealed interface. Because of the alternate platform configuration disclosed above, and the fastening mechanisms  600  and  606 , the joint between the platforms is better sealed than the prior art configurations and where the joint is not bolted together. For the non-bolted mateface, a gap exists and seals are used. With the chevron configuration disclosed herein, the gap has been moved further away from any bow wave coming off the airfoil leading edge, such that the hot combustion gases do not enter the gap between the vane assemblies. 
         [0036]    While the embodiments of the present invention improve the sealing between adjacent vanes of a vane assembly, any hot combustion gases that do leak between the platform surfaces can be minimized through alternate sealing arrangements. A plurality of flexible sheet metal seals (not shown) can be positioned in slots of the inner and outer platforms to prevent the flow of gases or compressed air in between any gaps of the platforms. More specifically, and as shown in  FIGS. 2 and 3 , the first vane assembly  200 , includes a series of slots in the platform mate faces, such as slots  220  in the first inner platform  202  and slots  222  in the outer platform  210 . 
         [0037]    Referring to  FIG. 7 , a chart showing static gap pressure along the mateface of the inner diameter platforms is shown. The solid line (without symbols) indicates the pressure of the cooling air under the platform. The line with the solid data points shows the gap pressure of the prior art vane of  FIG. 1 . As it can be seen from  FIG. 7 , the prior art vane assembly has a higher gap pressure than the underplatform pressure along part of the platform. This indicates that some of the hot combustion gases enters the platform gap (as caused by the bow wave and shape of vane platform), thereby causing erosion in the platform. The shaded data points indicate the gap pressure of the vane assembly  500  of the present invention. As it can be seen from  FIG. 7 , the inner vane platform gap pressure is lower than the under platform pressure. In this embodiment, because the inner vane platform gap pressure is lower than the under platform pressure, hot combustion gases do not enter the gap between adjacent platforms. 
         [0038]    While the present invention corrects an inflow problem along the inner arc-shaped platforms, no significant inflow problem exists at the outer platform for the prior art configuration, as shown in  FIG. 8 . However, the outer platform design of the present invention further increases the margin between the underplatform pressure side and the static pressure at the platform gap. 
         [0039]    The present invention has been described in relation to particular embodiments, which are intended in all respects to be illustrative rather than restrictive. Alternative embodiments will become apparent to those of ordinary skill in the art to which the present invention pertains without departing from its scope. 
         [0040]    From the foregoing, it will be seen that this invention is one well adapted to attain all the ends and objects set forth above, together with other advantages which are obvious and inherent to the system and method. It will be understood that certain features and sub-combinations are of utility and may be employed without reference to other features and sub-combinations. This is contemplated by and within the scope of the claims.

Technology Category: 2