Abstract:
A gas turbine airfoil ( 10 ) includes a serpentine cooling path ( 32 ) with a plurality of channels ( 34,42,44 ) fluidly interconnected by a plurality of turns ( 38,40 ) for cooling the airfoil wall material. A splitter component ( 50 ) is positioned within at least one of the channels to bifurcate the channel into a pressure-side channel ( 46 ) passing in between the outer wall ( 28 ) and the inner wall ( 30 ) of the pressure side ( 24 ) and a suction-side channel ( 48 ) passing in between the outer wall ( 28 ) and the inner wall ( 30 ) of the suction side ( 26 ) longitudinally downstream of an intermediate height ( 52 ). The cross-sectional area of the pressure-side channel ( 46 ) and suction-side channel ( 48 ) are thereby controlled in spite of an increasing cross-sectional area of the airfoil along its longitudinal length, ensuring a sufficiently high mach number to provide a desired degree of cooling throughout the entire length of the airfoil.

Description:
STATEMENT REGARDING FEDERALLY SPONSORED DEVELOPMENT 
     Development for this invention was supported in part by Contract No. DE-FC26-05NT42644, awarded by the United States Department of Energy. Accordingly, the United States Government may have certain rights in this invention. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates to the field of turbine vanes, and more particularly, the present invention relates to turbine vanes having cooling channels for passing cooling fluids to cool the turbine vanes. 
     BACKGROUND OF THE INVENTION 
     Gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine vane and blade assemblies to these high temperatures. As a result, turbine vanes and blades must be made of materials capable of withstanding such high temperatures. In addition, turbine vanes and blades often contain cooling systems for additional thermal protection. 
     Typically, turbine vanes are formed from an elongated portion forming an airfoil having one end configured to be coupled to a vane carrier and an opposite end configured to be movably coupled to an inner endwall. The turbine vane is ordinarily composed of a leading edge, a trailing edge, a suction side, and a pressure side. Additionally, the turbine vane includes an outer diameter endwall at a first end and an inner diameter endwall at a second end. The inner aspects of most turbine vanes typically contain an intricate maze of cooling circuits forming a cooling system. These cooling circuits in the vanes receive air from the compressor of the turbine engine and pass the air through the ends of the vane adapted to be coupled to the vane carrier. The cooling circuits often include multiple flow paths that are designed to maintain all areas of the turbine vane at a relatively uniform temperature. At least some of the air passing through these cooling circuits may be exhausted through orifices in the wall of the vane. 
     U.S. Pat. No. 6,955,523 to McClelland discloses such a cooling circuit including a serpentine network of channels passing between the suction and pressure sides of the turbine vane, where each channel extends between turns of the serpentine network positioned at the inner diameter and outer diameter endwalls. 
     U.S. Patent Application Publication No. 2005/0244270 to the inventor of the present invention, discloses a cooling circuit for a turbine blade including channels within the suction and pressure sides for passing cooling fluid toward the turbine blade tip at the first end for creating a counterflow to a leakage flow of combustor gases between the blade tip and an outer seal. 
     An additional cooling system for a turbine blade is disclosed in U.S. Patent Application Publication No. 2005/0031452, also to the inventor of the present invention, and discloses directing cooling fluid into a center cavity between the pressure and suction sides, after which the cooling fluid flows through supply orifices and into cavities within the suction and pressure walls for spiral fluid flow before exiting the turbine blade through exhaust orifices in the outer surface of the pressure and suction sides. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained in the following description in view of the drawings that show: 
         FIG. 1  is a perspective view of a turbine vane according to one embodiment of the present invention. 
         FIG. 2  is a cross-sectional view of the turbine vane of  FIG. 1  taken along the line  2 - 2 . 
         FIG. 3  is a cross-sectional view of the turbine vane of  FIG. 2  taken along the line  3 - 3 . 
         FIG. 4  is a cross-sectional view of the turbine vane of  FIG. 1  taken along the line  4 - 4 . 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     For certain airfoil designs having an increasing cross-sectional area along a longitudinal axis extending from an inside diameter portion to an outside diameter portion, the cooling channel structure of known serpentine cooling networks includes a large cross-sectional area increase from the inner diameter endwall to the outer diameter endwall. The present inventor has recognized that this results in a reduced cooling fluid flow rate toward the outer diameter portion of the airfoil, and that such a reduction of the fluid flow rate necessitates an over-cooling of radially inward portions of the airfoil in order to ensure adequate cooling of the radially outward portions of the airfoil. 
     Referring to  FIGS. 1-4 , a turbine vane  10  in accordance with the present invention will now be described that addresses the shortcomings of the prior art designs. The turbine vane  10  includes a cooling system  11  in inner aspects of the turbine vane  10  for use in turbine engines. While the description below focuses on a cooling system  11  in a stationary turbine vane  10 , the cooling system  11  may also be used in a rotating turbine blade. The present invention is particularly useful for turbine airfoils wherein the cross-sectional area of the airfoil increases from the inside diameter endwall to the outside diameter endwall by a factor of at least 1.5:1. 
     The turbine vane  10  illustratively includes a leading edge  12 , a trailing edge  14 , an outer diameter endwall  16  at a first end  18 , and an inner diameter endwall  20  at a second end  22  longitudinally opposite the first end. The turbine vane  10  further includes a generally concave shaped pressure side  24  coupling the leading edge  12  and the trailing edge  14  and a generally convex shaped suction side  26  positioned opposite from the pressure side. The pressure side  24  and the suction side  26  extend radially outward from an inner diameter at the second end  22  to an outer diameter at the first end  18 . An outer wall  28  defines at least a portion of the outer surfaces of the pressure side  24  and suction side  26 . An inner wall  30  is positioned relative to the outer wall on both the pressure side  24  and suction side  26 . 
     The cooling system  11  includes a serpentine cooling path  32  including a plurality of channels longitudinally extending from adjacent the first end  18  to adjacent the second end  22 . Additionally, the serpentine cooling path  32  includes a plurality of turns  38 , 40  with each turn positioned adjacent to the first or second end  18 ,  22  for coupling consecutive channels. The plurality of channels illustratively includes an inflow channel  34  longitudinally extending adjacent the leading edge  12  from an inlet  36  adjacent the first end  18  to a first turn  38  adjacent the second end  22 . Further, the plurality of channels includes a plurality of intermediate channels  42  passing in between the outer wall  28  and inner wall  30 , including a first intermediate channel  42  extending between the first turn  38  and a second turn  40  adjacent the first end  18 . Additionally, subsequent intermediate channels  42  similarly extend between consecutive turns  38 , 40  at the respective second and first end  22 , 18  of the turbine vane  10 . The plurality of channels further include an outflow channel  44  extending adjacent the trailing edge  14  from a last turn  40  to an outlet  70  adjacent the second end  22 . A rib  64  may longitudinally extend from adjacent the first end  18  to adjacent the second end  22  for separating consecutive channels of the plurality of channels. Although  FIG. 2  illustrates one inflow channel  34 , a plurality of intermediate channels  42  and one outflow channel  44 , other arrangements may be used such as a plurality of inflow channels and outflow channels, and a single intermediate channel may be utilized in the serpentine cooling path  32 . Additionally, an additional outlet  71  may be positioned adjacent the first turn  38  between the inflow channel  34  and the first intermediate channel  42 . 
     As may be best appreciated by viewing  FIG. 3 , each intermediate channel  42  extends from the second end  22  and bifurcates into a pair of intermediate channels at an intermediate height  52 . The pair of intermediate channels includes a pressure-side channel  46  passing in between the outer wall  28  and the inner wall  30  of the pressure side  24  and a suction-side channel  48  passing in between the outer wall  28  and the inner wall  30  of the suction side  26 . The pressure-side channel  46  and the suction-side channel  48  mutually diverge in extending to adjacent the first end  18 . A splitter component  50  is positioned within each of the intermediate channels  42 , and longitudinally extends from an intermediate height  52  to adjacent the first end  18 . The splitter  50  divides the intermediate channel  42  into respective pair of diverging channels  46 ,  48 . The splitter component  50  includes a pressure face  54  and suction face  56  respectively aligned with the pressure side  24  and the suction side  26 . The pressure face  54  and suction face  56  mutually diverge parallel with the pressure-side channel  46  and the suction-side channel  48  from a common diverging point at the intermediate height  52  along the radial length of the vane to adjacent the first end  18 . The pressure face  54  and suction face  56  bifurcate each intermediate channel  42  into the pair of intermediate channels including the pressure-side channel and suction-side channel  46 ,  48 , thus providing a near wall cooling fluid flow along each of the pressure and suction sides at locations downstream of the intermediate height  52 . The splitter component  50  may include a hollow or solid center portion between the pressure face  54  and suction face  56 . 
     The cross-sectional flow area of each intermediate channel  42  from the second end  22  to the first end  18  is reduced by inserting the splitter component  50  into the intermediate channel. The splitter component  50  may be sized to control and regulate the cross-sectional area of the pressure-side channel  46  and the suction-side channel  48 . The splitter component may be sized to minimize the variation in cross-sectional area of the pressure-side channel  46  and suction-side channel  48  along its longitudinal length. The cross-sectional flow area of the channels  46 ,  48  may be approximately constant from the intermediate height diverging point  52  to their respective ends, and the sum of these two flow areas may remain approximately equal to the cross-sectional flow area of the intermediate channel at the diverging point  52 . A typical mach number variation of the cooling fluid flow rate through a turbine vane of the prior art may be from 0.06 to 0.02 along the length of the airfoil. Selection of the size, geometry and location of the splitter component  50  enables a designer of an airfoil of the present invention to control the variation in mach number to any desired limited range, such as from 0.06 to 0.08. 
     At incremental positions between the leading edge  12  and the trailing edge  14 , an intermediate channel  42  is passed through the turbine vane  10  and bifurcated into a pair of intermediate channels, a pressure-side and suction-side channel  46 ,  48 . Cooling fluid passes through the pressure-side channels and suction-side channels of adjacent incremental positions in an opposite flow direction. The number and positioning of such incremental positions of the pressure and suction-side channels  46 ,  48  between the leading and trailing edges  12 ,  14  is selectively determined so to maintain a minimum threshold flow rate of the cooling fluid through each pressure and suction-side channel so to maintain a desired cooling efficiency for the turbine vane cooling system. In an exemplary embodiment of the present invention, the minimum threshold flow rate of the cooling system may be a mach number of 0.08, for example. 
     Consecutive turns  38 , 40  for an intermediate channel  42  are positioned adjacent an inner diameter cavity  60  along the inner diameter endwall  20  and adjacent an outer diameter cavity  62  along the outer diameter endwall  16 . The inner diameter cavity  60  and the outer diameter cavity  62  respectively extend adjacent the second end  22  and the first end  18  of the turbine vane  10 . 
     A portion of the inner surface of the channels may include at least one skew trip strip  66  for increasing the heat transfer coefficient by causing turbulent flow through the respective channel. 
     The outflow channel  44  may include one or more cooling holes  68  along the trailing edge  14 , where each of the cooling holes extends from the inner surface of the outflow channel to the outer surface of the trailing edge. The outflow channel  44  may further include one or more outlets  70  adjacent the inner diameter cavity  60 , where each outlet extends from the inner surface of the inner diameter cavity to the outer surface of the inner diameter endwall. Each outlet  70  may direct used cooling fluid to a rim cavity (not shown) positioned external to the turbine vane  10 . 
     During operation, the cooling fluid flows through the inlet  36  and into the inflow channel  34 , around the first turn  38 , and into a first intermediate channel  42 . The cooling fluid flows toward the first end  18  and upon reaching the intermediate height  52  within the intermediate channel  42 , the cooling fluid is bifurcated into a pressure-side channel  46  and a suction-side channel  48 . Each of the suction-side channel and pressure-side channel  46 , 48  then extend to the outer diameter cavity  62  adjacent the first end  18 . Within the outer diameter cavity  62 , the cooling fluid traverses toward the trailing edge  14 , before taking a second turn  40  into a pressure-side channel  46  and suction-side channel  48  of an adjacent intermediate channel  42 . The cooling fluid passes through each of the pressure-side channel  46  and suction-side channel  48  in the direction of the second end  22 , before merging at the intermediate height  52  where the splitter component  50  ends. The cooling fluid then flows within the intermediate channel  42  to the inner diameter cavity  60  adjacent the second end  22 . The cooling fluid continues through the serpentine cooling path  32  in this fashion and upon taking a last turn adjacent the first end  18 , enters the outflow channel  44 . The cooling fluid flows toward the second end  22  and partially diffuses out the trailing edge  14  through cooling holes  68  in the trailing edge. Further, a portion of the cooling fluid flows to the second end  22  and exits out an outlet  70  to a rim cavity external to the turbine vane. 
     While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.