Abstract:
The present invention provides a turbine blade having a revised under-platform structure including a unique coating combination that reduces mechanical stress factors within the turbine blade. The turbine blade includes a platform with an airfoil extending upwardly from the airfoil and a root portion extending downwardly from the platform. Two suction side tabs extend a first distance outward from a suction side of the root potion. Two pressure side tabs extend outward from a pressure side of the root portion. One of the two pressure side tabs extends outward a distance similar to the first distance, however, the other of the two pressure side tabs extends outward a distance much smaller than the first distance, which reduces stresses acting on the turbine blade. In addition, a plurality of coatings are systematically applied to the turbine blade to further reduce mechanical stress factors and improve cooling.

Description:
BACKGROUND OF THE INVENTION  
       [0001]     This application relates generally to a turbine blade for a gas turbine engine wherein a tab structure under the platform is modified.  
         [0002]     Conventional gas turbine engines include a compressor, a combustor and a turbine assembly that has a plurality of adjacent turbine blades disposed about a circumference of a turbine rotor. Each turbine blade typically includes a root that attaches to the turbine rotor, a platform, and a blade that extends radially outwardly from the turbine rotor.  
         [0003]     The compressor receives intake air. The intake air is compressed by the compressor and delivered primarily to the combustor where the compressed air and fuel are mixed and burned in a constant pressure process. A portion of the compressed air is bled from the compressor and fed to the turbine to cool the turbine blades.  
         [0004]     The turbine blades are used to provide power in turbo machines by exerting a torque on a shaft that is rotating at a high speed. As such, the turbine blades are subjected to a myriad of mechanical stress factors. In addition, the turbine blades are typically cooled using relatively cool air bled from the compressor resulting in temperature gradients being formed, which can lead to additional elements of thermal-mechanical stress within the turbine blades.  
         [0005]     Further, because the turbine blades are located downstream of the combustor where fuel and air are mixed and burned in a constant pressure process, they are required to operate in an extremely harsh environment. Traditionally, a chromium-based coating is applied to the entire turbine blade to resist the corrosive effects associated with this harsh environment. The traditional coating protects primarily against stress corrosion in areas of low stress concentration, however, the traditional coating does not provide adequate protection against stress corrosion in areas of high stress concentration, for example, under the platform.  
         [0006]     As such, it is desirable to provide a turbine blade that is optimized to reduce the effects of the mechanical and environmental stress factors.  
       SUMMARY OF THE INVENTION  
       [0007]     The present invention provides a turbine blade having a revised under-platform structure, including a novel coating process and a configuration that reduces mechanical and environmental stress factors within the turbine blade.  
         [0008]     The turbine blade includes a platform with an airfoil extending upwardly from the platform and a root portion extending downwardly from the platform. The turbine blade has a pressure side and a suction side. Two suction side tabs extend a first distance outwardly from the suction side of the root portion below the platform. Two pressure side tabs extend outwardly from the pressure side of the root portion below the platform. One of the two pressure side tabs extends outwardly a distance similar to the first distance, however, the other of the two pressure side tabs extends outwardly a second distance that is significantly less than the first distance. The shorter of the two pressure tabs regionally decreases mechanical stress factors within the turbine blade.  
         [0009]     In addition, a plurality of coatings are systematically placed and layered to reduce mechanical and environmental stress factors. A first coating is applied to substantially cover the turbine blade on both sides of the platform. The first coating protects against corrosion in areas of low stress concentration. However, the area under the platform of the turbine blade at the root portion is subjected to much higher stress concentrations than other areas of the turbine blade. Therefore, a second coating is applied over the first coating only under the platform. The second coating is added to resist corrosion cracking in areas of high stress concentration. The second coating is applied using a line-of-sight coating process through an access area that is created as a result of the shortened pressure side tab. The second coating is applied underneath the platform by spraying the coating directly at the shorter of the two pressure side tabs. Additional coatings are applied to the turbine blade to further reduce the effects of stress.  
         [0010]     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0011]      FIG. 1  is a schematic illustration of an example gas turbine engine;  
         [0012]      FIG. 2  illustrates a prior art turbine blade;  
         [0013]      FIG. 3  illustrates a pair of prior art turbine blades;  
         [0014]      FIG. 4  illustrates an example turbine blade according to one embodiment of the present invention;  
         [0015]      FIG. 5A  shows a cross-sectional illustration of a pair of prior art tabs; and  
         [0016]      FIG. 5B  shows a cross-section illustration of a pair of tabs according to one embodiment of the present invention.  
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT  
       [0017]      FIG. 1  is a schematic illustration of an example gas turbine engine  10  circumferentially disposed about an engine centerline, or axial centerline axis  12 . The example gas turbine engine  10  includes a fan  14 , a compressor  16 , a combustor  18 , and a turbine assembly  20 . As is known, intake air from the fan  14  is compressed in the compressor  16 , the compressed air is mixed with fuel that is burned in the combustor  18  and expanded in the turbine assembly  20 . The turbine assembly  20  includes rotors  22  and  24  that, in response to the expansion, rotate, driving the compressor  16  and the fan  14 . The turbine assembly  20  includes alternating rows of rotary blades  26  and static airfoils or vanes  28 , which are mounted to the rotors  22  and  24 . The example gas turbine engine  10  may, for example, be a gas turbine used for power generation or propulsion. However, this is not a limitation on the present invention, which may be employed on gas turbines used for electrical power generation, in aircraft, etc.  
         [0018]      FIG. 2  schematically illustrates a prior art turbine blade  30  having a platform  32 , with an airfoil  34  extending upwardly from the platform  32  and a root  36  extending downwardly from the platform  32 . The turbine blade  30  includes a pressure side  38  and a suction side  40 . A first set of tabs  42  is disposed on the root  36  on the pressure side  38  of the turbine blade  30  below the platform  32 . A second set of tabs  43  is disposed on the root  36  on the suction side  40  of the turbine blade  30  below the platform  32 . Notably in  FIG. 2 , only one of each set of tabs  42  and  42  are shown. However, it should be understood a second tab is disposed behind the one illustrated tab.  
         [0019]     The second set of tabs  43  extends outwardly from the root  36  on the suction side  40  in a first direction that is substantially parallel to the platform  32 . The first set of tabs  42  extends outwardly from the root  36  on the pressure side  38  in a second direction, substantially opposite the first direction. The second direction is also substantially parallel to the platform  32 .  
         [0020]      FIG. 3  schematically illustrates a pair of adjacent prior art turbine blades  30 A and  30 B. Each turbine blade,  30 A and  30 B, includes a root  36 , a platform  32  and an airfoil  34  as described previously in  FIG. 2 . A damper  44  is disposed between the adjacent turbine blades  30 A and  30 B, below the adjacent platforms  32 A and  32 B. The damper  44  is positioned between a first set of tabs  45  disposed on the suction side  40  of root  36 A of the turbine blade  30 A and a second set of tabs  47  disposed on the pressure side  38  of the root  36 B of the turbine blade  30 B. Notably, as in  FIG. 2 , only one of each set of tabs  45  and  47  are shown. However, it should be understood a second tab is disposed behind the one illustrated tab.  
         [0021]      FIG. 4  illustrates a turbine blade  60  according to one embodiment of the present invention. The turbine blade  60  includes an airfoil  62  extending upwardly from one side of a platform  64  and a root  66  extending downwardly from the platform  64 . The turbine blade  60  includes a leading edge  63  and a trailing edge  65  and has a pressure side  68  and a suction side  70 . The root  66  includes a front face  78  adjacent to the leading edge  63  and a rear face  74  adjacent to the trailing edge  65 . A first tab  72  is disposed on the pressure side  68  of the root  66  below the platform  64  and closest to the rear face  74  of the root  66 . A second tab  76  is disposed on the pressure side  68  of the root  66  below the platform  64  and closest to the front face  78  of the root  66 .  
         [0022]     The first tab  72  and the second tab  76  extend outwardly from the pressure side  68  of the root  66  in a direction substantially parallel to the platform  64 . The second tab  76  is significantly shorter than the first tab  72 . A third tab and a fourth tab are positioned on the suction side  70  of the root  66 , similar to the prior art, and have lengths that are similar to the first tab  72 . The tabs are used to position the damper as shown in  FIG. 3 .  
         [0023]     The first tab  72 , the third tab and the fourth tab respectively include a base portion  72 A and a post portion  72 B. The second tab  76  includes only a base portion  76 A. By only using the base portion  76 A in this region, an amount of mechanical stress imposed on the turbine blade  60  in this region is reduced. While the inventive turbine blade  60  is disclosed for use in a first stage turbine assembly, the inventive turbine blade  60  may be used in any stage.  
         [0024]     To further reduce the effects of stress on the turbine blade  60 , a plurality of coatings are applied to specified portions of the turbine blade  60 . A first coating, which in this example is a chromium-based coating, is applied to substantially cover the turbine blade  60  for corrosion protection. The first coating is applied to resist stress corrosion in areas of low stress concentration. Any type of chromium-based coating may be used.  
         [0025]     A second coating is applied over the first coating to address high stress areas on the turbine blade  60 . One high stress area is an area under the platform  64 , more specifically a region surrounding the base portion  72 A of the first tab  72  and including the first tab  72 . This area is subjected to much higher stress concentrations than the remainder of the turbine blade  60 . Further, the area under the platform  64  is susceptible to a different type of corrosion, that is, corrosion that occurs as a result of the high stress concentration. As such, the second coating, which is also chromium-based, is applied only under the platform  64  to resist stress corrosion is areas of high stress concentrations. This second coating is applied using a line-of-sight application process in which a sprayer, shown schematically at  200  in  FIG. 5B , is positioned to deliver the second coating through an access area created as a result of the second tab  76  only having a base portion  76 A. The second coating is sprayed underneath the platform by directing spray directly at the second tab  76 . The application of the second coating may include heat treating prior to application to prepare the surface by removing oxidation to ensure proper adhesion of the second coating.  
         [0026]     A third coating is applied over the first coating only on the airfoil  62 . In this example, the third coating is a metallic-bond coating which assists in adherence of a fourth coating applied over the third coating only on the airfoil  62 . This improves adhesion of a fourth coating, which in this example is a ceramic coating. The combination of coatings used on the airfoil  62  may include a heat treat process to ensure adhesion. Further, the combination of coatings reduces the effects of the harsh environment on the turbine blade  60 .  
         [0027]     Finally, a fifth coating is applied over the fourth coating only to a tip  80  of the turbine blade  60  to facilitate blade cutting. The fifth coating is a cubic boron nitride (CBN) coating. To ensure the tight clearances required by the turbine engine, the tips of the turbine blades are required to cut-in to the case surrounding the turbine engine. As such, the fifth coating is sacrificial, maintaining its integrity only long enough to ensure adequate run-in.  
         [0028]     The types of coatings discussed above are examples of each coating and other types of coatings could also be used to provide the desired characteristics.  
         [0029]     A comparison of the geometries of the tabs of the prior art and the present invention is more clearly illustrated in  FIGS. 5A and 5B , which show cross-sectional comparison of the tabs in the prior art and in one embodiment of the present invention respectively.  
         [0030]      FIG. 5A  illustrates a cross-sectional view of prior art tabs  42 . Each tab includes a base portion  42 A and a post portion  42 B. Each base portion  42 A extends outwardly from a pressure side  38  along a first distance D 1 . Each post portion  42 B extends outwardly from the base portion  42 A along a second distance D 2 , which is greater than the first distance D 1 . Therefore, the overall length L of the prior art tabs  42  is the same, that is, L=D 1 +D 2 .  
         [0031]      FIG. 5B  illustrates a cross-sectional view of tabs  72  and  76  according to one embodiment of the present invention. The first tab  72  includes a first base portion  72 A and a first post portion  72 B. The first base portion  72 A extends outwardly from a pressure side  68  along a first distance D 1 . The first post portion  74 B extends outwardly from the first base portion  72 A along a second distance D 2 , which is greater than the first distance D 1 .  
         [0032]     The second tab  76  includes only a base portion  76 A. This base portion  76 A extends outwardly from the pressure side  68  along a third distance D 3 , which is approximately equal to D 1 . The overall length L of the first tab  72  is D 1 +D 2 , which is significantly greater than D 3 .  
         [0033]     Because the second tab  76  only includes the base portion  76 A, the mechanical stress in the region surrounding the base portion  76 A under the platform  64  is reduced. That is, because the second tab  76  of the present invention is shorter than the prior art tab  47 , it does not extend into the cavity created between two adjacent turbine blades  30 A and  30 B to support the damper  44 . As such, the mechanical stress, more specifically, the torsional stress induced by the damper  44  into the region under the platform  64  through the length of the prior art tab  47  no longer exists in the present invention.  
         [0034]     Further, as discussed above, because the second tab  76  only includes the base portion  76 A, the shorter second tab  76  provides an access area for coating application. This access provides an unimpeded line-of-sight for application of the second coating under the platform  64 , which ensures complete coverage of the area of highest stress concentration including the first tab  72 .  
         [0035]     While the present invention is illustrated in a turbine blade, it should be understood that the invention would also be beneficial in a static structure such as a stator or a vane.  
         [0036]     Although preferred embodiments of this invention have been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.