Patent Publication Number: US-7708525-B2

Title: Industrial gas turbine blade assembly

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
   This application claims the benefit of U.S. Provisional Application No. 60/654,770, filed on Feb. 17, 2005, the disclosure of which is herein incorporated by reference in its entirety. 

   FIELD OF THE INVENTION 
   This invention relates generally to gas turbine engines, and more particularly to systems for cooling platforms and preventing cracking of the platforms of industrial gas turbine blades. 
   BACKGROUND OF THE INVENTION 
   Concave platforms of cooled industrial gas turbine (IGT) blades experience high metal temperature and thermal strain during operation. For example, the GE 7FA+e 1 st  stage turbine blade experiences severe thermo mechanical fatigue (TMF) initiated cracking at a leading edge and trailing edge of the platform that leads to high scrap rates and possible platform separation during operation. The crack results from a large, thin uncooled concave platform constrained by a relatively cooler airfoil and buttress structure that puts the platform in a state of high compressive strain at steady state operating conditions. The transient start-up condition results in a more severe compressive strain than steady state because of the large mass difference between the platform web and the rest of the component. Because of the mass difference the platform heats up more rapidly. Similarly the platform cools down more rapidly upon shutdown putting the platform into a tensile loading condition. The field parts also experience platform thermal barrier coating (TBC) spallation and significant platform oxidation. 
   Accordingly, it is an object of the present invention to provide a gas turbine blade assembly which overcomes the above-mentioned drawbacks and disadvantages. 
   SUMMARY OF THE INVENTION 
   In an aspect of the present invention, a gas turbine blade assembly includes a neck defining a neck cavity, and has a first end and a second end at an opposite side relative to the first end. The assembly further includes a platform having first and second sides. The first side of the platform is disposed on and faces the second end of the neck. An airfoil is supported on the second side of the platform. The neck, platform and airfoil define at least one inner cooling passage extending from the first end to the second end of the neck and through the platform and into the airfoil. The neck defines at least one core channel extending between the cooling passage and the neck cavity. The platform defines at least one film cooling channel extending from a portion of the first side facing the neck cavity to a portion of the second side disposed exterior to the airfoil to permit cooling air to flow through the inner cooling passage into the neck cavity and through a portion of the platform exterior to the airfoil. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS 
       FIG. 1A  is a perspective view of a platform for a gas turbine blade assembly in accordance with the present invention. 
       FIG. 1B  is a cross-sectional view of the platform of  FIG. 1A  taken along the line B-B. 
       FIG. 2  is a cross-sectional view of the platform of  FIG. 1A . 
       FIG. 3  is a perspective view of the platform of  FIG. 1A  showing inner cooling passages. 
       FIG. 4  is an enlarged perspective view of the platform of  FIG. 1A . 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
   With reference to  FIGS. 1A and 1B , an industrial gas turbine engine blade assembly is indicated generally by the reference number  10 . The assembly  10  includes a neck  12  defining a neck cavity  13 , and has a base or first end  14  and a second end  16  at an opposite side relative to the base. The assembly  10  includes a concave platform  18  disposed along an upper portion of the neck  12 , and has a first side  20  facing the second end  16  of the neck. The assembly  10  further includes an airfoil  22  supported on a second or opposite side  24  of the platform  18  relative to the neck  12  and extending outwardly from the platform. The airfoil  22  includes a concave side  26  and an oppositely facing convex side  28 . The platform  18  has a rail structure  30  and includes a leading edge  32  and a trailing edge  34 . 
   The neck  12 , the platform  18  and the airfoil  22  cooperate to define at least one inner cooling passage—preferably a plurality of inner cooling passages  36  including leading edge and trailing edge cooling passages as shown in FIG.  3 —extending therethrough from the base or first end  14  of the neck to the second end  16  and through the platform  18  and into the airfoil  22 . The neck  12  also defines at least one and preferably a plurality of core channels  38  extending between the inner cooling passages  36  and the neck cavity  13 . The core channels  38  are disposed on either the concave side  26  or the convex side  28  of the neck  12 . A portion of the platform  18  disposed exterior and adjacent to either the concave side  26  or the convex side  28  of the airfoil  22  defines a plurality of film cooling channels  40  extending from a portion of the first side  20  of the platform  18  facing the neck cavity  13  to a portion of the second side  24  of the platform disposed exterior to the airfoil  22  to permit cooling air to flow through the inner cooling passages  36  into the neck cavity  13  and through a portion of the platform exterior to the airfoil. 
   In operation, the gas turbine blade assembly  10  in accordance with the present invention reduces the metal temperature and thermal strain in the platform  18  of the airfoil  22 . The neck cavity  13  is pressurized via the core channel  38 . The pressurized neck cavity  13  feeds the film cooling channels  40  to cool the platform  18 . The cooled platform  18  also reduces platform oxidation and thermal barrier coating (TBC) spallation. This active platform cooling can be implemented to repair used industrial gas turbine blades and to extend the usable life of such blades by an additional overhaul cycle. The assembly  10  in accordance with the present invention can also be included as a beneficial feature in new or re-engineered industrial gas turbine blades. 
   In addition to actively cooling the platform  18 , casting grain control can be employed to reduce the strain level in the platform. Industrial gas turbine blade directionally solidified (DS) castings tend to have a large single crystal (SC) grain for the entire platform area. This single crystal platform grain significantly increases the limiting strain level in the platform and the likelihood for thermo mechanical fatigue (TMF) crack initiation. The cracking also propagates along the large grain boundary. Casting parameters and processes can be used to control the platform grain and produce a more beneficial equiax grain state in the platform region without sacrificing the benefits of a directionally solidified grain in the airfoil. Grain control in accordance with the present invention can only apply to new or re-engineered industrial gas turbine blades. 
   The orientation of the core channel  38  preferably directs the flow of cooling air to impinge on an underside of the platform  18 . A tube brazed into the core channel  38  and laid against the neck  12  could be used to direct core flow to impinge more effectively upon the underside of the platform  18 . The core channels  38  could be created by machining or casting methods. In a particular GE 7FA+e 1 st  blade repair application, the core channel  38  is preferably 0.175 inches in diameter, pulls air from the inner cooling passage  36 , has a circular shape, and extends between a trailing edge cooling passage  36  and the neck cavity  13  as shown in  FIG. 3 . 
   In an exemplary embodiment, the film cooling channels  40  defined by the platform  18  are an array of .015 inch-.050 inch diameter holes oriented to provide maximum convective and film cooling while minimizing stress concentrations. The number of film cooling channels  40  varies preferably from three to fifteen. The film cooling channels  40  extend through the concave platform  18  entering on an underside (the first side  20 ) of the platform and exiting at the platform flow path at the second side  24  thereof. An alternate location for the film cooling channels is through a rail  42  on a forward edge  44  of the concave platform, entering on a back side  43  of the rail and exiting on the edge of the concave platform (inside platform gap of assembled blades). With respect to a particular GE 7FA+e 1 st  blade repair application, the array of film cooling channels  40  includes seven .035 inch diameter holes extending through the platform  18  and oriented at an acute angle of about 30 degrees from a surface  46  of the platform and at an acute angle of about 30 degrees from the edge  44  of the platform. In another example, the acute angle is 43 degrees from the edge  44  (see  FIG. 4 ). The angles shown relative to the dotted line represent the angle between the film cooling channels  40  and the primary gas flow (dotted lines). 
   Pressurized air from the neck cavity  13  can also be used to feed the film cooling channels exiting on the convex side  28  in order to cool other platform locations. A film cooling channel into the pressurized neck cavity could be used to purge a trailing edge undercut in a new or re-engineered industrial gas turbine blade as disclosed more fully in U.S. Ser. No. 10/738,288 filed on Dec. 17, 2003, the disclosure of which is herein incorporated by reference in its entirety. 
   An exemplary embodiment of the platform  18  is illustrated in  FIG. 4 . The platform  18  defines seven film cooling channels  40   a ,  40   b ,  40   c ,  40   d ,  40   e ,  40   f  and  40   g . The film cooling channels are each about 0.035 inches in diameter, and are about 0.285 inches long (Length/Diameter=8.143). The surface angle of the film cooling channels  40  is about 30 degrees. The exit angle of the film cooling channels is about −30 degrees relative to the edge  44  of the platform  18 . There are no diffusers at a film cooling channel exit, but diffusers could be used to improve cooling film effectiveness. The angles shown in  FIG. 4  represent the angle between the hole injection angle and the angle of the primary gas flow (dotted lines). 
   As will be recognized by those of ordinary skill in the pertinent art, numerous modifications and substitutions can be made to the above-described embodiment of the present invention without departing from the scope of the invention. Accordingly, the preceding portion of this specification is to be taken in an illustrative, as opposed to a limiting sense.