Patent Publication Number: US-7708528-B2

Title: Platform mate face contours for turbine airfoils

Description:
This invention was made with government support under Contract No. F33615-03-D-235-0006 awarded by the United States Air Force. The government therefore has certain lights in this invention. 

   BACKGROUND OF THE INVENTION 
   This application relates to an improved airfoil, wherein mate faces between adjacent airfoils are contoured to optimize cooling air flow between the mate faces. 
   Various components in a gas turbine engine have an airfoil shape extending outwardly from a platform. One example is a turbine blade, which typically includes a platform, with an airfoil extending above the platform. The airfoil is curved, extending from a leading edge to a trailing edge, and between a pressure wall and a suction wall. 
   The turbine blade can become quite hot during operation of the gas turbine engine. Thus, cooling circuits are formed within the turbine blade to circulate cooling fluid, typically air. A number of cooling channels extend through the cross-section of the airfoil, and from the platform outwardly toward a tip. Air passes through these channels, and cools the turbine blade. 
   Many distinct types of cooling circuits are provided within the airfoil, and associated structures such as a platform, the root, etc. As known, a number of turbine blades are mounted to be circumferentially spaced. Leakage air is allowed to flow between a leading face of the platform, a trailing face of the platform, and on mate faces between adjacent platforms. This air cools the platforms, and allows the airfoils to better survive in the harsh environment of the gas turbine engine. 
   The platform has side edges that define mate faces. The cooling air flow between the mate faces has been directed by the gap between the mate faces. The gap is parallel to the mate faces, and the mate faces have traditionally been parallel to a groove within the root such that the blade can be more easily mounted to a rotor. 
   Applicant has determined that, for various reasons, providing cooling air flow from a gap between generally straight edges of a platform, does not optimize this cooling air flow. Instead, applicant has recognized that there are hot spots on the platform due to several features that are not best addressed by the prior cooling air flow. 
   In at least one prior art airfoil, the platform side edges are defined by a pair of straight sections. This was to allow the use of a platform having an edge extending on an angle that might otherwise intersect with the airfoil. This has not been utilized to address local hot spots. 
   SUMMARY OF THE INVENTION 
   In a disclosed method of this invention, an airfoil is studied, and heat stresses along the platform are identified. Localized hot spots are identified adjacent the approximate area of the mate faces. The mate faces are then designed to assure optimum cooling air flow from the gap between the mate faces over these hot spots. The present invention thus results in a platform for turbine airfoil components, which has better cooling characteristics due to the optimized direction of the cooling air between the mate faces. In addition, and flowing from the above-described benefit, internal cooling channels may be eliminated as not being necessary. Thus, the present invention not only improves operation, but may also reduce the complexity of manufacturing the turbine blade. 
   In a disclosed embodiment of this invention, the mate face has a curved portion near the hot spot, and then a second straight portion. 
   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 
       FIG. 1  shows a plurality of prior art gas turbine engine turbine blades. 
       FIG. 2  is a top view of the prior art blades. 
       FIG. 3  is a top view of an inventive blade made according to an inventive method. 
   

   DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
   Prior art turbine blades  20  are illustrated in  FIG. 1 , having airfoils  22 , and platforms  26 . As is known, a root portion  27  is utilized to mount the turbine blade  20  within a rotor. 
   The roots  27  have grooves  31  for being received in a mating structure on the rotor. Gaps  28  are formed between mating faces  40  and  42  on adjacent turbine blades. The airfoils  22  each have a leading edge  29  and a trailing edge  131 . Air flow leaks around the platform  26  at the leading edge as shown at  30 , and at the trailing edge as shown at  32 . Further, air flow leaks at  34  between a gap  28  between the mating faces  40  and  42 . These air flows assist in cooling the turbine blade  20 , and in particular along the platforms  26 . As known, the airfoils have a pressure wall  38  and a suction wall  36 . 
   In the prior art illustrated in  FIG. 1 , the mate faces  40  and  42  are defined by edges of the platform  26  that extend generally parallel to the grooves  31  in root  27 . 
   As shown in  FIG. 2 , the edges  40  and  42  on platforms  26  extend generally parallel to each other, and along a generally straight line. 
   Heat stress analysis shows hot spots  44  on the platforms  26 . The air flow from the gaps  28  flows along the platform, and as controlled by the movement of the turbine blades, etc. The air flow paths or streamlines can be mapped and studied. However, this air flow has never been controlled or designed to flow in a particular direction based upon the location of the hot spot  44 . Applicant has now considered the heat stress and air flow streamlines, and has identified an improved mate face to direct cooling air to the platform. As shown in  FIG. 2 , a path  45  extends along a curve, and is an optimum location of the air flow  34  for addressing the hot spot  44 . As mentioned, the design of the airfoil platform has never taken this path  45  into account. That is, path  45  is not part of the prior art. 
     FIG. 3  shows an inventive gas turbine blade  50 , wherein the airfoil  52  has a suction wall  54 , a leading edge  56 , a pressure wall  58  and a trailing edge  60 . One edge of the platform  61  has a straight portion  62  leading to a curved indent  64 . The opposed side of the platform  61  has a similar straight section  62  leading to a bulged section  66 . The sections  64  and  66  are formed along curves that may be optimally modeled on the path  45 . When the blades  50  are mounted within a gas turbine environment, air flow leaks between the gaps between turbine blades  50 . The air flow from the gap between sections  64  and  66  is directed to best address the local hot spots  44 . As is clear, the curved sections  64  and bulged sections  66  are spaced away from the trailing edge  60  relative to the straight portions  62 . 
   As can be appreciated in  FIG. 3 , the contours of the sections  64  or  66  generally can be said to have a leading edge section  80 , an intermediate section  82 , and a merging section  84 , which merges with the relatively straight portions  62 . The sections  80  and  84  extend along curves, but have a major component in their direction that is parallel to the path of the relatively straight sections  62 . The intermediate section  82  extends along a curve that has a larger component perpendicular to the direction of the relatively straight section  62 . It has been found that this general contour provides the best cooling air flow paths to address the hot spots  44 , at least in a number of turbine blade designs. 
   Further, while the term “relatively straight portions” has been utilized to define portion  62 , it should be understood that the part does extend along contours and curves in several directions, and thus, the surface may not be identically straight. 
   As can be appreciated from  FIG. 3 , the straight portions  62  are non-perpendicular to the leading edge face  100  and trailing edge  102  of the platform. 
   The present invention thus improves upon the prior art. 
   While the present invention is specifically disclosed in a turbine blade, it has application in the design of any gas turbine engine components having airfoils and platforms wherein the components are mounted to be adjacent to each other and cooling air flow is provided between the mating faces. As an example, static vanes would benefit from this invention, as would other components that meet this basic definition. 
   Although a preferred embodiment of this invention has 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.