Patent Publication Number: US-8523524-B2

Title: Airfoil cooling hole flag region

Description:
BACKGROUND OF THE INVENTION 
     The subject matter disclosed herein relates to an airfoil having a cooling hole with a flag region. 
     In turbine engines, such as gas turbine engines or steam turbine engines, fluids at relatively high temperatures contact blades that are configured to extract mechanical energy from the fluids to thereby facilitate a production of power and/or electricity. While this process may be highly efficient for a given period, over an extended time, the high temperature fluids tend to cause damage that can degrade performance and increase operating costs. 
     Accordingly, it is often necessary and advisable to cool the blades to at least prevent or delay premature failures. This can be accomplished by delivering relatively cool compressed air to the blades. In many traditional gas turbines, in particular, this compressed air enters the bottom of each of the blades and flows through one or more round machined passages in the radial direction to cool the blade through a combination of convection and conduction. 
     In these traditional gas turbines, as the temperature of the fluids increase, it becomes necessary to increase the amount of cooling flow through the blades. This increased flow can be accomplished by an increase in a size of the cooling holes. However, as the cooling holes increase in size, the wall thickness of each hole to the external surface of the blade decreases and eventually reaches a minimum wall thickness required to maintain manufacturability and structural integrity of the blade. 
     BRIEF DESCRIPTION OF THE INVENTION 
     According to one aspect of the invention, an airfoil is provided and includes a body formed to define a substantially radially extending cooling hole therein, which is configured to be receptive of a supply of a coolant for removing heat from the body, and a flag region therein, which is fluidly communicative with the cooling hole and thereby configured to be receptive of a portion of the supply of the coolant such that the coolant portion is directed to form a vortex within the flag region to increase heat removal from the body beyond that provided by the coolant flow through the cooling hole. 
     According to another aspect of the invention, an airfoil of a turbine bucket is provided and includes a body having opposing pressure and suction surfaces extending axially between opposing leading and trailing edges and radially between inward and outward portions, the body being formed to define a substantially radially extending cooling hole therein, which is configured to be receptive of a supply of a coolant such that the coolant is forced to flow along a length thereof to remove heat from the body, and the body being further formed to define a flag region therein, which is fluidly communicative with the cooling hole and thereby configured to be receptive of a portion of the supply of the coolant such that the coolant portion is directed to form a vortex within the flag region to increase heat removal from the body beyond that provided by the coolant flow through the cooling hole. 
     These and other advantages and features will become more apparent from the following description taken in conjunction with the drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWING 
       The subject matter, which is regarded as the invention, is particularly pointed out and distinctly claimed in the claims at the conclusion of the specification. The foregoing and other features, and advantages of the invention are apparent from the following detailed description taken in conjunction with the accompanying drawings in which: 
         FIG. 1  is a perspective view of an airfoil; 
         FIGS. 2 and 3  are perpendicular plan views of the airfoil of  FIG. 1 ; and 
         FIGS. 4 and 5  are perspective views of an airfoil according to further embodiments. 
     
    
    
     The detailed description explains embodiments of the invention, together with advantages and features, by way of example with reference to the drawings. 
     DETAILED DESCRIPTION OF THE INVENTION 
     With reference to  FIGS. 1-3 , an airfoil  10  of a turbine bucket is provided. The airfoil  10  includes coolant  11  and a body  20  having opposing pressure and suction surfaces  21  and  22  extending axially between opposing leading and trailing edges  23  and  24  and radially between inward and outward portions  25  and  26 . 
     The body  20  may be an airfoil blade body and is formed to define a substantially radially extending cooling hole  30  therein, which is configured to be receptive of a supply of a coolant  11  such that the coolant  11  is forced to flow along a length thereof to remove heat from the body  20 . The cooling hole  30  may be of ovoid or round or non-ovoidal or non-round shapes such as, for example, elliptical, race track, rectangular etc. The body  20  is further formed to define a flag region  40  therein. The flag region  40  is fluidly communicative with the cooling hole  30  and thereby configured to be receptive of a portion of the supply of the coolant  11  such that the coolant  11  portion is directed to form a vortex  12  within the flag region  40 . The vortex formation increases heat removal from the body  20  beyond that which is provided by the flow of the coolant  11  through the cooling hole  30 . 
     A width, W, of the flag region  40  may be substantially similar to that of the cooling hole  30  in the circumferential direction. The flag region  40  may tangentially extend in an axial direction from a location of maximum circumferential width of the cooling hole  30 . A corner  41  of the flag region  40  may be defined with a right angle and, in some cases, the flag region  40  may be formed to have a substantially rectangular or square cross-section in at least one of radial and axial directions. 
     With reference to  FIGS. 4 and 5 , although the flag region  40  is described above as having a substantially rectangular shape, it is to be understood that this is merely exemplary and that other shapes and configurations are possible. For example, as shown in  FIG. 4 , the flag region  40  may, in some cases, have a non-rectangular shape  401  with edges at right or non-right angles, and which are rounded or non-rounded. Similarly, as shown in  FIG. 5 , the flag region  40  may also have a symmetrical shape or a non-symmetrical shape  402 . In each case, as will be described below, the shapes and radial spacing between a flag region  40  and another flag region  40  may vary along the length of cooling hole  30 . 
     The flag region  40  may be plural in number, as shown in  FIG. 1 . The plural flag regions  40  may be arrayed along the cooling hole  30  in a radial direction. In some embodiments, the plural flag regions  40  may be arrayed along an entire length of the cooling hole  30  in the radial direction. Conversely, the plural flag regions  40  may be arrayed along only a portion of the cooling hole  30  length. 
     The plural flag regions  40  may each have similar or, in some cases, differing shapes and may be aligned with or offset from one another. Where the flag regions  40  are offset, a degree of the offset is set to in accordance with a twist of the body  20 . However, even where the flag regions  40  are offset from one another, they may still be aligned in at least one dimension. For example, as shown in  FIG. 2 , even if the body  20  is twisted in a manner not evident from  FIG. 2 , the flag regions  40  are aligned in the radial direction. 
     The plural flag regions  40  may also be radially discrete in that the flag regions  40  are aligned with one another in the radial direction and separated by areas of airfoil material. Here, the radially discrete plural flag regions  40  may be spaced from one another by either a uniform radial distance or a variable radial distance that is established based on a known heating profile of the airfoil  10 . 
     As shown in  FIG. 3 , the flag regions  40  may be substantially equidistant from the pressure and suction surfaces  21  and  22  and closer to the trailing edge  24  than the leading edge  23  although this is not required. At least one sidewall  42  delimiting the flag region  40  may be substantially or nearly parallel with a local portion  43  of at least one of the pressure and suction surfaces  21  and  22 . In any case, however, a wall thickness, T w , between the flag region  40  and the pressure and suction surfaces  21  and  22  is at least a predefined minimum thickness. This predefined minimum thickness should be a minimum thickness that preserves the operability and manufacturability of the airfoil  10 . 
     In accordance with further aspects of the invention, the airfoil  10  may be defined with multiple cooling holes  30  with each cooling hole  30  being associated with zero, one or more flag regions  40 . For example, a series of cooling holes  30  may be arrayed axially along the camber line of the airfoil  10  with only the most downstream one or two cooling holes  30  having flag regions  40 . 
     In accordance with still further aspects of the invention, the cooling holes  30  and the flag regions  40  may be formed within the airfoil  10  by machining processes, such as electro-chemical machining (ECM) or the like. In particular, a heating profile of the airfoil  10  may be determined through testing to illustrate where the airfoil  10  is most likely to be heated beyond safe levels. Then, the cooling holes  30  and the flag regions  40  can be machined in those regions to thereby maintain a lower temperature therein. 
     Additionally, if it is found that only a small portion of the airfoil tends to be heated beyond the safe levels, the machining of the cooling holes  30  and the flag regions  40  can be strictly limited to that small portion. As such, a structural impact of the cooling holes  30  and the flag regions  40 , in terms of local areas of high stress, for example, can be substantially reduced. 
     While the invention has been described in detail in connection with only a limited number of embodiments, it should be readily understood that the invention is not limited to such disclosed embodiments. Rather, the invention can be modified to incorporate any number of variations, alterations, substitutions or equivalent arrangements not heretofore described, but which are commensurate with the spirit and scope of the invention. Additionally, while various embodiments of the invention have been described, it is to be understood that aspects of the invention may include only some of the described embodiments. Accordingly, the invention is not to be seen as limited by the foregoing description, but is only limited by the scope of the appended claims.