Abstract:
A vane assembly for a gas turbine engine is disclosed having lower thermally induced stresses resulting in improved component durability. The stresses in the vane assembly airfoils are lowered by increasing the flexibility of the vane platform and reducing its resistance to thermal deflection. This is accomplished by placing an opening along the innermost vane assembly rail that reduces the effective stiffness of the platform, thereby lowering the operating stresses in the airfoils of the vane assembly. A removable seal is then placed in the opening in order to prevent undesired leakages, while maintaining the benefit of the increased platform flexibility.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS  
       [0001]     This application is a continuation-in-part of U.S. patent application Ser. No. 10/891,400, filed on Jul. 14, 2004, and assigned to the same assignee hereof. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT  
       [0002]     Not Applicable.  
       TECHNICAL FIELD  
       [0003]     The present invention relates generally to gas turbine engines and more specifically to a turbine vane configuration having reduced airfoil stresses.  
       BACKGROUND OF THE INVENTION  
       [0004]     A gas turbine engine typically comprises a multi-stage compressor, which compresses air drawn into the engine to a higher pressure and temperature. A majority of this air passes to the combustors, which mix the compressed heated air with fuel and contain the resulting reaction that generates the hot combustion gases. These gases then pass through a multi-stage turbine, which, in turn drives the compressor, before exiting the engine. A portion of the compressed air from the compressor bypasses the combustors and is used to cool the turbine blades and vanes that are continuously exposed to the hot gases of the combustors. In land-based gas turbines, the turbine is also coupled to a generator for generating electricity.  
         [0005]     Turbines are typically comprised of alternating rows of rotating and stationary airfoils. The stationary airfoils, or vanes, direct the flow of hot combustion gases onto the subsequent row of rotating airfoils, or blades, at the proper orientation such as to maximize the output of the turbine. As a result of the hot combustion gases passing through the vanes, the vanes operate at a very high temperature, typically beyond the capability of the material from which they are made. In order to lower the operating temperatures of the vane material to a more acceptable level, vanes are often cooled, either by air or steam. Typically, turbine vanes are configured in multiple segments, with each segment including a plurality of vanes. This configuration is well known in order to minimize hot gas leakage between adjacent vanes, thereby lowering turbine performance. While this configuration is advantageous from a leakage perspective, it has inherent disadvantages as well, including an increased stiffness along the platform that connects the adjacent vanes, relative to a single vane configuration.  
         [0006]     A vane assembly  10  of the prior art, is shown in  FIG. 1 , and comprises an inner platform  11 , inner rail  12 , outer platform  13 , and vanes  14  extending between inner platform  11  and outer platform  13 . While the inner rail serves as a means to seal the rim cavity region from leakage of the cooling air into the hot gas path instead of passing to the designated vanes, inner rail  12  also stiffens inner platform  11 . Inner rails  12 , which can be rather large in size, are located proximate the plenum of cooling air and are therefore operating at approximately the temperature of the cooling air. As a result, hot combustion gases passing around vanes  14  and between inner platform  11  and outer platform  13  cause the vanes and platforms to operate at an elevated temperature relative to the inner rail. This sharp contrast in operating temperatures creates regions of high thermally induced stresses in vanes  14  and along inner platform  11  that has been known to cause cracking of the vane assembly requiring premature repair or replacement.  
         [0007]     What is needed is a vane assembly configuration that lowers the operating stresses in the vane and platform for a vane assembly having an inner rail portion that is exposed to lower operating temperatures than the platform or vane.  
       SUMMARY OF THE INVENTION  
       [0008]     A turbine vane assembly for use in a gas turbine engine is disclosed having lower thermally induced stresses in the airfoil and platform region resulting in improved component durability. In an embodiment of the invention, the vane assembly comprises a first platform, a second platform positioned radially outward of the first platform, and at least one airfoil extending therebetween. The source of cracking in prior art vane assemblies related to the significant temperature differences over a short radial distance between the vane, platform, and first rail, located along the first platform, opposite to the airfoil. In the present invention, the first platform further comprises a first rail having a first rail length, a first rail height, a first rail thickness, a first rail wall, and at least one opening extending from the first rail wall and through the first rail thickness. The at least one opening is sized to allow the first platform to have reduced resistance to thermal deflections while not compromising the structural integrity of the first platform nor allowing leakage of vane cooling fluid.  
         [0009]     It is an object of the present invention to provide a turbine vane assembly having reduced thermal stresses in the airfoil and platform regions.  
         [0010]     It is another object of the present invention to provide a turbine vane assembly having increased flexibility along the first platform region.  
         [0011]     In accordance with these and other objects, which will become apparent hereinafter, the instant invention will now be described with particular reference to the accompanying drawings.  
         [0012]     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 DRAWING  
       [0013]     The present invention is described in detail below with reference to the attached drawing figures, wherein:  
         [0014]      FIG. 1  is a perspective view of a turbine vane assembly of the prior art;  
         [0015]      FIG. 2  is a cross section view of a portion of a gas turbine engine in which an embodiment of the present invention operates;  
         [0016]      FIG. 3  is a detailed cross section view of a portion of a turbine section of a gas turbine engine in which an embodiment of the present invention operates;  
         [0017]      FIG. 4  is a partial end view of a portion of the turbine taken generally perpendicular to the view of  FIG. 3  in accordance with an embodiment of the present invention;  
         [0018]      FIG. 5  is a perspective view of a turbine vane assembly in accordance with an embodiment of the present invention;  
         [0019]      FIG. 6  is a detailed perspective view of a portion of a turbine vane assembly in accordance with an embodiment of the present invention; and  
         [0020]      FIG. 7  is an end view of a portion of a turbine vane assembly in accordance with the preferred embodiment of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0021]     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 steps or combinations of steps similar to the ones described in this document, in conjunction with other present or future technologies. Moreover, although the terms “step” and/or “block” may be used herein to connote different elements of methods employed, the terms should not be interpreted as implying any particular order among or between various steps herein disclosed unless and except when the order of individual steps is explicitly described.  
         [0022]     The present invention is shown in detail in  FIGS. 2-7 . Referring initially to  FIG. 2 , a partial cross section of a typical gas turbine engine  15  is shown. The engine includes an air inlet  16 , a compressor  17 , a combustion system  18 , a turbine  19 , with the compressor  17  and turbine  19  coupled along a longitudinal axis, denoted as A-A, that extends through the engine and is the axis about which the plurality of blades and vanes in the compressor  17  and turbine  19  are positioned circumferentially. Note that the airfoils extend outward in a radial direction. A more detailed view of a portion of the turbine  19  is shown in cross section in  FIG. 3 , in which alternating rows of rotating airfoils (blades)  40  and stationary airfoils (vanes)  30  are shown.  
         [0023]     Referring now to  FIG. 4 , an elevation view looking aft is shown in which a plurality of vane assemblies  30  are shown assembled in an array.  FIG. 4  is taken generally perpendicular to  FIG. 3 .  
         [0024]     Referring now to  FIGS. 4 and 5 , a vane assembly for a gas turbine engine in accordance with an embodiment of the present invention is shown. Vane assembly  20  comprises a first arc-shaped platform  21  having a first thickness  22 , a forward wall  34  and an aft wall  35 , and a first rail  23  extending generally circumferentially along the non-flowpath side of the first arc-shaped platform  21 . The first rail  23 , which is shown in greater detail in  FIGS. 6 and 7 , further comprises a first rail length  24 , a first rail height  25 , a first rail thickness  26 , a first rail wall  27 , and at least one opening  28  that is substantially cylindrical in shape. The specific dimensions of rail length  24 , rail height  25 , and rail thickness  26  can vary depending on the turbine vane configuration and location in the engine. The at least one opening  28  extends through the first rail thickness  26  and has a slot  36  initiating at the first rail wall  27  and extends radially outward to the opening  28 . As previously mentioned, the greatest temperature gradient and corresponding highest thermal stress is at the region of the hottest portion of the airfoil  30  and the rail  23  intersect. The opening is preferably positioned along the first rail  23  at the location of highest thermal stress between the first rail  23  that operates at a lower temperature than the adjacent platform and airfoil. While the exact location of the opening  28  can vary, it is often located radially beneath an airfoil  30 .  
         [0025]     As it can be seen from  FIGS. 4 and 5 , vane assembly  20  also comprises a second arc-shaped platform  29  that is positioned radially outward of the first arc-shaped platform  21 . The second platform  29  also has at least one second rail  32  that extends generally circumferentially along the second arc-shaped platform  29 . For the embodiment disclosed in the figures, it can be understood that the first rail  23  and at least one second rail  32  are both arc-shaped with the arcs corresponding to their associated arc-shaped platform. The rails are located along the side of the sides of the platforms opposite of the airfoil  30 . As one skilled in the art will understand, with both the first platform  21  and the second platform  29  each having an arc-shape and separated by at least one radially extending airfoil  30 , then for a given number of vane segments about the engine axis, the second rail  32  will have a length  33  that is greater than the first rail length  24 . This difference in length can be seen in  FIG. 4 . In one embodiment of the invention, a total of 24 vane assemblies comprise a stage of the turbine (as previously discussed). The second rail  32  for this vane assembly, is located approximately 49 inches from the longitudinal axis A-A while the first rail  23  is located approximately 38 inches from the same longitudinal axis A-A. Therefore, for this vane assembly  20 , the first rail  23  has a rail length  24  of approximately 9.95 inches while the second rail length  33  for the second rail  32  is approximately 12.83 inches.  
         [0026]     As previously discussed, extending radially outward to the second arc-shaped platform  29  from the first arc-shaped platform  21  is at least one airfoil  30 . The airfoil  30  extends from the first arc-shaped platform  21 , opposite from the first rail  23 . For the embodiment shown in the figures, two airfoils are present in each vane assembly  20 . However, it is important to note that the present invention can be applied to a vane assembly having fewer or greater number of airfoils  30 . As one skilled in the art will understand, turbine blades and vanes operate at extremely high temperatures, often times at temperatures that would ordinarily exceed the capability of the material. As such, the vane assemblies  20  of the present invention pass a cooling fluid through the airfoils  30  for lowering the operating temperatures. The cooling fluid is typically air, but can also be steam.  
         [0027]     The vane assembly  20  further comprises a seal  31  as shown in  FIG. 6 . The seal  31 , which is preferably a metal plate, is placed into the slot  36  that extends radially outward from first rail wall  27  such that the seal  31  closes off the opening  28  in first arc-shaped rail  23 . The seal  31  prevents the leakage of any fluids through the now more pliable first arc-shaped rail  23 . The seal can be secured to the first rail  23  by a variety of means including tack welding, peening, or any other method by which the seal can be removed if desired, such that the structural freedom achieved by opening  28  is maintained.  
         [0028]     The focus of the present invention is directed towards the first rail  23  and at least one opening  28  located therein, which is shown in the figures is the inner rail closest to the axis A-A. The stress relief provided to the first rail  23  by the opening  28  could be applied to a variety of vane assemblies and is not limited to the embodiment disclosed. The opening  28  is configured to allow the first arc-shaped platform  21  to have increased flexibility while not compromising the structural integrity of the platform. For example, in the preferred embodiment of the present invention, the opening  28  comprises a slot having a generally circular end, as shown in  FIGS. 4-7 . This opening configuration reduces the platform effective stiffness thereby increasing platform flexibility and reducing the resistance to thermal deflections imposed by a multiple airfoil vane assembly. Reducing the resistance to thermal deflections allows for release of the thermal stresses in the first arc-shaped platform  21  and airfoil  30  due to their differing thermal gradients. For the particular embodiment shown in  FIGS. 4-7 , the configuration of opening  28  resulted in approximately 14% reduction in airfoil stresses. The quantity of openings  28 , their respective location along the first rail  23 , and their respective configuration depends on the stress levels of the vane assembly configuration, which in turn is a function of at least the quantity of airfoils, aerodynamic shape of the airfoils, operating temperatures, and material composition, etc. It is important for opening  28  to include a rounded end so as to not introduce any locations having a concentrated stress that could result in potential crack initiation.  
         [0029]     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. While the invention has been described in what is known as presently the preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment but, on the contrary, is intended to cover various modifications and equivalent arrangements within the scope of the following claims.