Patent Publication Number: US-7722326-B2

Title: Intensively cooled trailing edge of thin airfoils for turbine engines

Description:
FIELD OF THE INVENTION 
     This invention is directed generally to turbine blades, and more particularly to cooling systems in hollow turbine blades. 
     BACKGROUND 
     Typically, 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 blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades must be made of materials capable of withstanding such high temperatures. In addition, turbine blades often contain cooling systems for prolonging the life of the blades and reducing the likelihood of failure as a result of excessive temperatures. 
     Typically, turbine blades are formed from a root portion at one end and an elongated portion forming a blade that extends outwardly from a platform coupled to the root portion. A turbine blade ordinarily includes a tip opposite to the root section, a leading edge, and a trailing edge. The inner aspects of turbine blades typically contain an intricate maze of cooling channels forming a cooling system. The cooling channels in the blades receive air from the compressor of the turbine engine and pass the air through the blade. The cooling channels often include multiple flow paths that are designed to maintain all aspects of the turbine blade at a relatively uniform temperature. 
     The trailing edge of a turbine blade is difficult to cool because the trailing edge is often too thin to effectively cool using known embodiments. Because the trailing edge of a blade is difficult to cool and is often exposed to both high temperatures and high loads, the trailing edge may suffer from creep or oxidation during operation. The detrimental effects may be most pronounced in the radially outward portion of the blade proximate to the blade tip because the elongated airfoil is thinner at the tip. The problem is generally most severe in the rear stages of a turbine where the entire elongated airfoil is generally thinner than the elongated airfoils of the front stages. Thus, a need exists for a turbine blade cooling system that effectively cools the trailing edge of a rear stage turbine blade. 
     SUMMARY OF THE INVENTION 
     The present invention is directed to a turbine blade cooling system designed to cool the trailing edge of a turbine blade usable in rear stages of a turbine engine. The cooling system may be configured to cool aspects of the trailing edge despite the relative thin thickness of the turbine blade proximate to the trailing edge. In particular, the cooling system may exhaust cooling fluids through the tip rather than through the trailing edge, thereby not further weakening the region of the airfoil proximate to the trailing edge. 
     The turbine blade may include a leading edge cooling cavity and a trailing edge cooling cavity separated by an impingement rib with impingement orifices therein. The trailing edge cooling cavity may be in fluid communication with the exterior of the blade through at least one exhaust orifice in the tip of the blade. The trailing edge cooling cavity may be designed such that cooling fluid passing from the leading edge cooling cavity to the trailing edge cooling cavity impinges on a trailing edge cooling cavity surface proximate to the trailing edge. The trailing edge cooling cavity may also be designed so that a cooling fluid is drawn from the leading edge cooling cavity and into the trailing edge cooling cavity before exiting through the exhaust orifices in the blade tip. 
     The turbine blade may include a generally elongated blade having a leading edge, a trailing edge, and a tip at a first end. A platform may be located generally orthogonal to the generally elongated blade and proximate an end of the generally elongated blade opposite the tip. The blade may include a leading edge cooling cavity disposed generally spanwise within the generally elongated blade and may have a portion located proximate the leading edge. A trailing edge cooling cavity may be disposed generally spanwise within the generally elongated blade and may have a portion located proximate the trailing edge. The cross-sectional area of the trailing edge cooling cavity taken generally orthogonal to a radial axis of the generally elongated blade may generally increase moving from a radially inward end of the trailing edge cooling cavity toward a radially outward end of the trailing edge cooling cavity. The blade tip may include an exhaust orifice having a first opening in fluid communication with the trailing edge cooling cavity and a second opening located in an outer surface of the generally elongated blade. The blade may include an impingement rib separating the leading edge cooling cavity from the trailing edge cooling cavity and extending generally spanwise along the generally elongated blade. The impingement rib may include an impingement orifice positioned with the first opening of the impingement orifice in fluid communication with the leading edge cooling cavity and the second opening of the impingement orifice in fluid communication with the trailing edge cooling cavity. 
     In one embodiment, the impingement rib may include a plurality of impingement orifices. The plurality of impingement orifices may be asymmetrically distributed along the length of the impingement rib. The density of the impingement orifices may decrease moving from the end of the generally elongated blade proximate the platform toward the tip. 
     The cross-sectional area of the impingement orifices may decrease moving from the end of the generally elongated blade proximate the platform toward the tip. The cross-sectional area of the impingement orifices may decrease non-linearly. 
     The turbine blade may include a plurality of exhaust orifices in the blade tip. The total cross-sectional area of the impingement orifice openings may be less than, equal to, or greater than a total cross-sectional area of the exhaust orifice openings. If there is more than one exhaust orifice, the exhaust orifices may be distributed asymmetrically along the length of the blade tip. 
     The cross-sectional area of the leading edge cooling cavity taken generally orthogonal to the radial axis of the generally elongated blade may decrease moving from the radially inward end of the leading edge cooling cavity toward the radially outward end of the leading edge cooling cavity. The cross-sectional area of the leading edge cooling cavity may decrease non-linearly. 
     An advantage of this invention is that the cooling system enables the trailing edge region of a rear stage turbine blade to be adequately cooled without further weakening the region. 
     Another advantage of this invention is that the cooling system may provide impingement cooling to the trailing edge of the turbine blade. 
     Yet another advantage of the invention is that the trailing edge cooling cavity may be designed so that the impingement effect is not distorted by the cross-flow of cooling fluid. 
     Another advantage of the invention is that the cooling system provides improved convective cooling of the trailing edge by increasing the flow of cooling fluid in the trailing edge cooling cavity proximate to the trailing edge of the blade. 
     These and other embodiments are described in more detail below. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Other objects, features and advantages of the present invention will become apparent upon reading the following detailed description, while referring to the attached drawings, in which: 
         FIG. 1  is a perspective view of the a turbine blade containing a trailing edge cooling system of the present invention. 
         FIG. 2  is a cross-sectional view of the turbine blade of  FIG. 1 , taken along section line  2 - 2 , that shows a turbine airfoil having a leading edge cooling cavity, a trailing edge cooling cavity, an impingement rib, impingement orifices and exhaust orifices. 
         FIG. 3  is a cross-sectional view of the turbine blade of  FIG. 2 , taken along section line  3 - 3 , that shows a turbine airfoil having a trailing edge cooling cavity. 
         FIG. 4  is a cross-sectional view of the turbine blade of  FIG. 2 , taken along section line  4 - 4 , that shows a turbine airfoil having a trailing edge cooling cavity with a cross-sectional area larger than a cross-sectional area of the trailing edge cooling cavity shown in  FIG. 3 . 
         FIG. 5  is a cross-sectional view of the turbine blade of  FIG. 2 , taken along section line  5 - 5 , that shows an impingement rib having a plurality of impingement orifices asymmetrically distributed therein. 
         FIG. 6  is a cross-sectional view of the turbine blade of  FIG. 2 , taken along section line  6 - 6 , that shows an impingement rib having a plurality of impingement orifices with decreasing cross-sectional areas moving from one end to the other. 
         FIG. 7  is an end view of the turbine blade of  FIG. 1  that depicts the blade tip having an plurality of exhaust orifices asymmetrically distributed therein. 
         FIG. 8  is an end view of the turbine blade of  FIG. 1  that depicts the blade tip having an plurality of oval-shaped exhaust orifices asymmetrically distributed therein. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     As shown in  FIGS. 1-8 , this invention is directed to a cooling system  12  usable in a turbine blade  10  that is configured to be used in rear stages of a turbine of a turbine engine. The cooling system  12  may be configured to cool aspects of the trailing edge  18  despite the relatively thin thickness of the turbine blade  10  proximate to the trailing edge  18 . In particular, the cooling system  12  may exhaust cooling fluids through the tip  20  rather than through the trailing edge  18 , thereby not further weakening the region of the airfoil  10  proximate to the trailing edge  18 . 
     In one embodiment, the turbine blade  10  may include a generally elongated blade  14  having a leading edge  16 , a trailing edge  18 , a tip  20 , and a platform  22  that is positioned generally orthogonal to the generally elongated blade  14  and located at an end of the generally elongated blade  14  opposite the tip  20 . The trailing edge  18  may be a nonperforated trailing edge  18  that lacks any exhaust orifices, as shown in  FIG. 2 . The turbine airfoil may also include a root  24  positioned proximate to the platform  22 . A leading edge cooling cavity  26  may extend generally spanwise within the generally elongated blade  14  with a portion located proximate to the leading edge  16 . A trailing edge cooling cavity  28  may be disposed generally spanwise within the generally elongated blade  14  and may have a portion located proximate to the trailing edge  18 . As shown in  FIGS. 3 &amp; 4 , the cross-sectional area of the trailing edge cooling cavity  28  taken generally orthogonal to a radial axis  30  of the generally elongated blade  14  may generally increase moving from the radially inward end of the trailing edge cooling cavity  28 , as shown in  FIG. 3 , to the radially outward end of the trailing edge cooling cavity  28 , as shown in  FIG. 4 . 
     Although not shown, the cross-sectional area of the trailing edge cooling cavity  28  taken generally orthogonal to a radial axis  30  of the generally elongated blade  14  may remain constant or even decrease over a portion of the trailing edge cooling cavity  28  moving from the radially inward end of the trailing edge cooling cavity  28 . As used herein, “generally increases” indicates that along at least 50%, preferably along at least 75%, more preferably along at least 85%, of the length of the trailing edge cooling cavity  28 , the cross-sectional area increases relative an the immediately adjacent portion of the trailing edge cooling cavity  28 . 
     In an embodiment of the present invention, the cross-section of the trailing edge cooling cavity  28  may be constant or even decrease over a portion of the trailing edge cooling cavity  28  proximate the blade tip  20 . This may be used to optimize cooling near the blade tip  20  using a Venturi effect by increase the velocity of cooling fluid near the tip of the generally elongated blade  14 . 
     As shown in  FIG. 2 , the generally elongated blade may include an exhaust orifice  32  in the blade tip  20 , positioned such that the first opening  34  of the exhaust orifice  32  is in fluid communication with the trailing edge cooling cavity  28  and the second opening  36  of the exhaust orifice  32  is located in an outer surface  38  of the blade tip  20 . An impingement rib  40  may extend generally spanwise within the generally elongated blade  14  and separate the leading edge cooling cavity  26  from the trailing edge cooling cavity  28 . An impingement orifice  42  may pass through the impingement rib  40 . The impingement orifice  42  may be positioned so that the impingement orifice  42  has a first opening  44  in fluid communication with the leading edge cooling cavity  26  and a second opening  46  in fluid communication with the trailing edge cooling cavity  28 . The cross-sectional area of the impingement orifice  42  may be larger than the cross-section of the exhaust orifice  32 . 
     In one embodiment, the turbine airfoil  10  may include a plurality of impingement orifices  42 . As shown in  FIGS. 5 &amp; 6 , the impingement orifices  42  may be asymmetrically distributed along the length of the impingement rib  40 . As shown in  FIG. 5 , the density of the impingement orifices  42  may decrease moving from the end of the impingement rib  40  proximate the platform  22  toward the blade tip  20 . The cross-sectional area of the impingement orifices  42  may decrease moving from the end of the impingement rib  40  proximate the platform  22  toward the blade tip  20 , as shown in  FIG. 6 . The cross-sectional area of the impingement orifices  42  may decrease non-linearly, as shown in  FIGS. 5 &amp; 6 . The impingement orifices  42  may be any appropriate shape including, but not limited to, circular, oval, triangular, rectangular, and others. 
     The turbine airfoil  10  may include a plurality of exhaust orifices  32  in the blade tip  20 , as shown in  FIGS. 7 &amp; 8 . The total cross-sectional area of the plurality of impingement orifices  42  may be less than the total cross-sectional area of the plurality of exhaust orifices  32 . As shown in  FIGS. 7 &amp; 8 , the plurality of exhaust orifices  32  may be distributed asymmetrically along the length of the blade tip  20 . The exhaust orifices  32  may be any appropriate shape including, but not limited to, circular, oval, triangular, rectangular, and others. 
     As shown in  FIGS. 3 &amp; 4 , the leading edge cooling cavity  26  may be designed such that the cross-sectional area of the leading edge cooling cavity  26  taken generally orthogonal to the radial axis  30  of the generally elongated blade  14  decreases moving from the radially inward end of the leading edge cooling cavity  26 , as shown in  FIG. 3 , toward the radially outward end of the leading edge cooling cavity  26 , as shown in  FIG. 4 . The cross-sectional area of the leading edge cooling cavity  26  may decrease in a non-linear manner. In addition, cross-sectional area of the leading edge cooling cavity  26 , may remain constant or even decrease moving from the radially inward end of the leading edge cooling cavity  26  toward the radially outward end of the leading edge cooling cavity  26 . 
     In order to cool the trailing edge  18  of cooled rear stage turbine blades  10 , a leading edge cooling cavity  26  may be in fluid communication with the trailing edge cooling cavity  28 . Cooling fluid may be fed into the leading edge cooling cavity  26 , or any other channel adjacent to the trailing edge cooling cavity  28 , by a compressor (not shown). Cooling fluid may flow from the leading edge cooling cavity  26  through the impingement orifices  42  and impinge upon the wall of the trailing edge cooling cavity  28  that forms the trailing edge  18 . Additional cooling fluid may enter the trailing edge cooling cavity  28  from any other channel adjacent to the trailing edge cooling cavity  28 . The trailing edge cooling system  12  is designed such that cooling fluid entering the trailing edge cooling cavity  28  travels radially outward toward the tip  20  of the generally elongated blade  14  and exits through the exhaust orifices  32 . 
     Although not shown, there may be more than two cooling cavities within the generally elongated blade  14 . As used herein, the trailing edge cooling cavity  28  is the cooling cavity most proximate the trailing edge  18 . As used herein, the leading edge cooling cavity  26  is adjacent to the trailing edge cooling cavity  28  and in fluid communication with the trailing edge cooling cavity  28  by at least one impingement orifice  42 . The leading edge cooling cavity  26  will be more proximate the leading edge  16  than the trailing edge cooling cavity  28 , however, the leading edge cooling cavity need not be the cooling cavity most proximate the leading edge  16 . 
     Impingement cooling, particularly when combined with convection cooling, is recognized as being superior to convection cooling alone. The present invention provides high velocity impingement cooling proximate to the trailing edge  18  without the need for channels exiting through the trailing edge  18 . This approach may be superior to approaches using channels that exit through the trailing edge  18  because the use of channels in the trailing edge  18  weakens the trailing edge  18 , which is vulnerable to creep due to high loads and insufficient cooling even without exhaust chambers extending through the trailing edge  18 . The trailing edge cooling cavity  28  may be free of channels that exhaust fluid through the trailing edge  18 . 
     The cross-sectional area of the trailing edge cooling cavity  28  taken generally orthogonal to the radial axis  30  of the generally elongated blade  14  may increase from the end of the generally elongate blade  14  proximate the platform  22  toward the blade tip  20 . Using this approach, the turbine blade  10  trailing edge cooling cavity  28  may be designed to ensure the impinging jets of cooling fluid do not get distorted by the flow of cooling fluid generally parallel to the radial axis  30 , i.e. the radial flow. In particular, the cross-sectional area of the trailing edge cooling cavity  28  may increase to maintain the radial velocity of cooling fluid in the trailing edge cooling cavity  28  relatively constant from the end  48  of the trailing edge cooling cavity  28  proximate the platform  22  to the end  50  of the trailing edge cooling cavity  28  proximate the blade tip  20 . 
     There are several parameters that may be used to maintain the non-impinging cooling fluid in the trailing edge cooling cavity  28  at a relatively constant radial velocity. In order to maintain the proper pressure differential between the leading edge cooling cavity  26  and the trailing edge cooling cavity  28 , the cooling fluid in the trailing edge cooling cavity  28  must exit through the exhaust orifices  32 . As cooling fluid in the leading edge cooling cavity  28  passes into the trailing edge cooling cavity  28 , an equal mass of cooling fluid must exit through the exhaust orifices  32 . Thus, one way to maintain cooling fluid in the trailing edge cooling cavity  28  at a relatively constant radial velocity is to have the cross-sectional area of the trailing edge cooling cavity  28  increase in relation to the number and size of the impingement orifices  42 . Maintaining the cooling fluid in the trailing edge cooling cavity  28  at a relatively constant radial velocity improves the impingement effect created by the impingement orifices  42  by reducing distortion and diffusion of the jets of cooling fluid impinging on the wall of the trailing edge cooling cavity  28  proximate to the trailing edge  18 . 
     Based on the foregoing, it will be recognized that a turbine blade  10  designed may utilize many parameters to properly implement the trailing edge cooling system  12  of the present invention. The trailing edge cooling system  12  may be designed to have a pressure differential between the leading edge cooling cavity  26  and the trailing edge cooling cavity  28  such that the cooling fluid passes through the impingement orifices  42  with a velocity sufficient for impingement cooling of the wall of the trailing edge cooling cavity  28  proximate to the trailing edge  18 . Whether the velocity of the cooling fluid is sufficient for impingement cooling is, in part, a function of the distance between the second opening  44  of the impingement orifice  42  and the wall of the trailing edge cooling cavity  28  proximate to the trailing edge  18 . Accordingly, the design of a trailing edge cooling system  12  may reflect a proper balance between the velocity of the impinging cooling fluid, the radial velocity of non-impinging cooling air in the trailing edge cooling cavity  28 , and the distance between the second opening  46  of the impingement orifice  42  and the wall of the trailing edge cooling cavity  28  proximate the trailing edge  18 . 
     The foregoing is provided for purposes of illustrating, explaining, and describing embodiments of this invention. Modifications and adaptations to these embodiments will be apparent to those skilled in the art and may be made without departing from the scope or spirit of this invention.