Abstract:
A main body is provided for a gas turbine engine comprising an outer structure, a first internal partition and a second internal partition. The outer structure and the first internal partition may define an entrance leg of a cooling circuit for receiving a cooling fluid. The second internal partition may include a metering slot. The outer structure, the first internal partition and the second internal partition may define an intermediate leg of the cooling circuit. The intermediate leg may communicate with the entrance leg. The second internal partition and the outer structure may define an exit leg of the cooling circuit. The metering slot meters cooling fluid as it passes from the intermediate leg into the exit leg.

Description:
FIELD OF THE INVENTION 
       [0001]    This invention is directed generally to turbine blades and, more particularly, to a turbine blade having cooling cavities for conducting a cooling fluid to cool a trailing edge of the blade. 
       BACKGROUND OF THE INVENTION 
       [0002]    A conventional gas turbine engine includes a compressor, a combustor and a turbine. The compressor compresses ambient air which is supplied to the combustor where the compressed air is combined with a fuel and ignites the mixture, creating combustion products defining a working gas. The working gas is supplied to the turbine where the gas passes through a plurality of paired rows of stationary vanes and rotating blades. The rotating blades are coupled to a shaft and disc assembly. As the working gas expands through the turbine, the working gas causes the blades, and therefore the shaft and disc assembly, to rotate. 
         [0003]    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. 
         [0004]    Typically, turbine blades comprise a root, a platform and an airfoil that extends outwardly from the platform. The airfoil is ordinarily composed of a tip, leading edge and a trailing edge. Most blades typically contain internal cooling channels forming a cooling system. The cooling channels in the blades may 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 the turbine blade at a relatively uniform temperature. However, centrifugal forces and air flow at boundary layers often prevent some areas of the turbine blade from being adequately cooled, which results in the formation of localized hot spots. Localized hot spots, depending on their location, can reduce the useful life of a turbine blade and can damage a turbine blade to an extent necessitating replacement of the blade. 
         [0005]    Operation of a turbine engine results in high stresses being generated in numerous areas of a turbine blade. One particular area of high stress is found in the blade&#39;s trailing edge, which is a portion of the blade forming a relatively thin edge that is generally orthogonal to the flow of gases past the blade and is on the downstream side of the blade. Because the trailing edge is relatively thin and an area prone to development of high stresses during operation, the trailing edge is highly susceptible to formation of cracks. These cracks may propagate and cause failure of the blade, which may, in some situations, cause catastrophic damage to a turbine engine. 
         [0006]    A conventional cooling system in a turbine blade assembly may discharge a substantial portion of the cooling air through a trailing edge of the blade. Typically, the cooling system contains an intricate maze of cooling flow paths in the trailing edge. There exist numerous configurations of the cooling flow paths that attempt to maximize the convection occurring in a trailing edge of a blade. While many of these conventional systems have operated successfully, a need still exists to provide increased cooling capability in the trailing edge portions of turbine blades. 
       SUMMARY OF THE INVENTION 
       [0007]    In accordance with one aspect of the invention, a turbine blade is provided comprising an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root. A blade tip surface is located at an end of the airfoil distal from the root, and the airfoil outer wall includes pressure and suction side surfaces joined together at chordally spaced apart leading and trailing edges of the airfoil. The airfoil defines an airfoil cavity forming a cooling system in the blade. At least a first rib is positioned in the airfoil cavity to form at least a first generally elongated cooling cavity along at least a portion of the span-wise direction in an area adjacent the trailing edge of the airfoil, the first rib including an upstream side and a downstream side. The first cooling cavity comprises a cavity pressure sidewall and a cavity suction sidewall extending from the downstream side of the first rib. The first rib includes at least one orifice extending through the first rib from the upstream side to the downstream side, and the cavity pressure and suction sidewalls define convergent cavity sidewalls relative to the pressure and suction side surfaces of the outer wall. 
         [0008]    In accordance with another aspect of the invention, a turbine blade is provided comprising an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root. A blade tip surface is located at an end of the airfoil distal from the root, and the airfoil outer wall includes pressure and suction side surfaces joined together at chordally spaced apart leading and trailing edges of the airfoil. The airfoil defines an airfoil cavity forming a cooling system in the blade. A first rib is positioned in the airfoil cavity to form a first generally elongated cooling cavity along at least a portion of the span-wise direction in an area adjacent the trailing edge of the airfoil, the first rib including an upstream side and a downstream side. The first cooling cavity comprises a cavity pressure sidewall and a cavity suction sidewall extending from the downstream side of the first rib, the first rib including a plurality of orifices extending through the first rib from the upstream side to the downstream side thereof. A second rib is positioned in the airfoil cavity to form a second generally elongated cooling cavity adjacent to the first cooling cavity, the second rib including an upstream side and a downstream side. The second cooling cavity comprises a cavity pressure sidewall and a cavity suction sidewall extending from the downstream side of the second rib, the second rib including a plurality of orifices extending through the second rib from the upstream side to the downstream side thereof. The cavity pressure and suction sidewalls in each of the first and second cooling cavities define convergent cavity sidewalls relative to the pressure and suction side surfaces of the outer wall. 
         [0009]    In accordance with a further aspect of the invention, a turbine blade is provided comprising an airfoil including an airfoil outer wall extending in a span-wise direction radially outwardly from a blade root. A blade tip surface is located at an end of the airfoil distal from the root, and the airfoil outer wall includes pressure and suction side surfaces joined together at chordally spaced apart leading and trailing edges of the airfoil. The airfoil defining an airfoil cavity forming a cooling system in the blade. A first rib positioned in the airfoil cavity to form a first generally elongated cooling cavity along at least a portion of the span-wise direction in an area adjacent the trailing edge of the airfoil, the first rib including an upstream side and a downstream side. The first cooling cavity comprising a cavity pressure sidewall and a cavity suction sidewall extending from the downstream side of the first rib, the first rib including a plurality of orifices extending through the first rib from the upstream side to the downstream side thereof. A second rib positioned in the airfoil cavity to form a second generally elongated cooling cavity adjacent to the first cooling cavity, the second rib including an upstream side and a downstream side. The second cooling cavity comprising a cavity pressure sidewall and a cavity suction sidewall extending from the downstream side of the second rib, the second rib including a plurality of orifices extending through the second rib from the upstream side to the downstream side thereof. A third rib positioned in the airfoil cavity to form a third generally elongated cooling cavity adjacent to the second cooling cavity, the third rib including an upstream side and a downstream side. The third cooling cavity comprising a cavity pressure sidewall and a cavity suction sidewall extending from the downstream side of the third rib, the third rib including a plurality of orifices extending through the third rib from the upstream side to the downstream side thereof. Each of the orifices in the third rib is substantially centered on a line extending along a centerline of a corresponding orifice in each of the first and second ribs. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]    While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein: 
           [0011]      FIG. 1  is a perspective view of a turbine blade incorporating the present invention; 
           [0012]      FIG. 2  is a cross-sectional view of the turbine blade shown in  FIG. 1  taken along line  2 - 2 ; 
           [0013]      FIG. 3  is an enlarged detail view of the trailing edge of the turbine blade shown in  FIG. 2 ; 
           [0014]      FIG. 4  is cross-sectional view of the turbine blade shown in  FIG. 1  taken along line  4 - 4 ; and 
           [0015]      FIG. 5  is an enlarged detail view of the trailing edge of the turbine blade shown in  FIG. 4 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0016]    In the following detailed description of the preferred embodiment, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, a specific preferred embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention. 
         [0017]    Referring to  FIG. 1 , an exemplary turbine blade  10  for a gas turbine engine is illustrated. The blade  10  includes an airfoil  12  and a root  14  which is used to conventionally secure the blade  10  to a rotor disk of the engine for supporting the blade  10  in the working medium flow path of the turbine where working medium gases exert motive forces on the surfaces thereof. The airfoil  12  has an outer wall  16  comprising a generally concave pressure sidewall  18  and a generally convex suction sidewall  20 . The pressure and suction sidewalls  18 ,  20  are joined together along an upstream leading edge  22  and a downstream trailing edge  24 . The leading and trailing edges  22 ,  24  are spaced axially or chordally from each other. The airfoil  12  extends radially along a longitudinal or radial direction of the blade  10 , defined by a span of the airfoil  12 , from a radially inner airfoil platform  26  to a radially outer blade tip surface  28 . 
         [0018]    Referring to  FIGS. 2 and 4 , the airfoil  12  defines one or more cavities  30  positioned between the pressure sidewall  18  and the suction sidewall  20 . The cavity  30  may include one or more cooling paths  32  ( FIG. 2 ) for directing a cooling fluid, such as cooling air, through the airfoil  12  and out various orifices or openings in the outer wall  16  of the airfoil  12 . For example, leading edge orifices or openings  34  may be provided in the leading edge  22  of the airfoil  12 , and additional surface film cooling orifices or openings  36  may be provided in the pressure and suction sidewalls  18 ,  20 . In addition, the tip surface  28  may also be provided with cooling openings  37 , as required to reduce temperatures across the tip surface  28 . Further, the trailing edge  24  is preferably also provided with trailing edge cooling orifices or openings  38  spaced along the trailing edge  24  in a span-wise direction, as will be described further below with regard to the cooling configuration for the trailing edge area of the airfoil  12 . 
         [0019]    The cavity  30  may be arranged in various configurations. For example, as illustrated in  FIG. 2 , cavity  30  may form cooling chambers that extend through the root  14  and airfoil  12 . In particular, the cavity  30  may extend from a location adjacent the tip surface  28  to one or more cooling fluid inlet openings  40   a ,  40   b ,  40   c ,  40   d  at an end of the root  14 . Alternatively, the cavity  30  may be formed only in portions of the airfoil  12 . The openings  40   a ,  40   b ,  40   c ,  40   d  may be configured to receive the cooling fluid, such as air from the compressor. Cavity  30  may include a rib  42  dividing the cavity  30  into a first elongated cooling chamber  44  positioned proximate the leading edge  22 , and a second elongated cooling chamber  46  positioned proximate the trailing edge  24 . In addition, one or more plates  48  may be provided to control or direct flow of the cooling fluid through the cavity  30 , such as by closing off one or more of the inlet openings  40   a ,  40   b ,  40   c ,  40   d , and shown herein as closing off the inlet opening  40   b.    
         [0020]    The first elongated cooling chamber  44  may include any number of cooling paths. For example, and not by way of limitation, the first elongated cooling chamber  44  may include a divider  50  forming a leading edge cooling chamber  52  proximate to the leading edge  22 . The divider  50  may include one or more orifices  54  and, by way of example, may include a plurality of orifices  54  that may or may not be equally spaced relative to each other along the divider  50 . In addition, one or more of the leading edge orifices  34  extend from the leading edge cooling chamber  52  to the outer surface of the leading edge  22 , and may be arranged in the leading edge  22  to form a shower head to expel cooling fluid from the first elongated cooling chamber  44 . 
         [0021]    The second elongated cooling chamber  46 , which may also be referred to as a body cavity of the airfoil  12 , may include any number of cooling paths. For example, and not by way of limitation, the second elongated cooling chamber  46  may include one or more dividers  56  forming a serpentine cooling path. The sidewalls of the cavity  30  may further be provided with trip strips  58  along the interior surfaces  60 ,  62  of the pressure and suction sidewalls  18 ,  20 , respectively, to increase turbulence of the flow of cooling air along the interior surfaces  60 ,  62  (see also  FIG. 4 ), and thereby improve heat transfer at the boundary layer between the cooling air flow and the interior surfaces  60 ,  62 . The configurations described above for the first and second elongated cooling paths  44 ,  46  may be arranged as described above and shown in  FIG. 2 , or may have other configurations appropriate to dissipate heat from the airfoil  12  during use. 
         [0022]    Referring to  FIGS. 2 and 4 , the cavity  30  may additionally include one or more impingement ribs  64  dividing cavity  30  and forming one or more elongated trailing edge cooling cavities  66  adjacent the second elongated cooling chamber  46 . The one or more impingement ribs  64  and trailing edge cooling cavities  66  may extend along only a portion of the distance between the platform  26  and the tip surface  28  or, alternatively, may extend substantially the entire distance between the platform  26  and the tip surface  28 . In a preferred non-limiting embodiment illustrated herein, the impingement ribs  64  comprise a first rib  64   a , a second rib  64   b  and a third rib  64   c  forming a first cooling cavity  66   a , a second cooling cavity  66   b  and a third cooling cavity  66   c , respectively. It should be understood that the designations of “first”, “second” and “third” are provided for convenience in describing the invention, and are not intended to be construed as limiting as to the particular location and/or number of impingement ribs  64  and cooling cavities  66 . 
         [0023]    Referring further to  FIGS. 3 and 5 , each of the ribs  64   a ,  64   b ,  64   c  includes one or more orifices  68  extending from an upstream side  70  to a downstream side  72  of each of the ribs  64   a ,  64   b ,  64   c . The orifices  68  in each rib  64   a ,  64   b ,  64   c  are arranged in spaced relation to each other and may be located in uniform or equidistance spaced relation to each other. However, it should be understood that the present invention is not limited to any particular spacing between orifices  68 , and that the spacing between the orifices  68  along any of the impingement ribs  64  may vary. Further, although the ribs  64  are illustrated as having orifices  68  along substantially the entire span-wise length thereof, the orifices  68  may located at only selected span-wise locations along the impingement ribs  64 , as needed for the particular cooling requirements of the airfoil  12 . 
         [0024]    A pair of cooling cavity sidewalls comprising a cavity pressure sidewall  74  and a cavity suction sidewall  76  extends in a downstream direction from the downstream side  72  of the impingement ribs  64 . The cavity pressure and suction sidewalls  74 ,  76  of the first and second cavities  66   a ,  66   b  terminate at the upstream sides  70  of the second and third ribs  64   b ,  64   c , respectively, and the cavity pressure and suction sidewalls  74 ,  76  of the third cavity  66   c  terminate at an upstream side  78  of a trailing section  80  defining the trailing edge  24 . The orifices  68  exit the impingement ribs  64  at the middle of the downstream sides  72 , generally midway between the cavity pressure and suction sidewalls  74 ,  76 . 
         [0025]    As seen in  FIGS. 4 and 5 , the pairs of cavity pressure and suction sidewalls  74 ,  76  extend in the downstream direction in converging relation to each other, such that the cavities  66   a ,  66   b ,  66   c  each define a generally triangular or teardrop shape where the downstream side  72  of each rib  64   a ,  64   b ,  64   c  forms the base of the triangular shape. It may be seen with reference to the first cavity  64   a  in  FIG. 5  that the cavity pressure sidewall  74  angles inwardly at an acute angle θ away from a line  83  parallel to an outer surface  82  of the pressure sidewall  18 , and the cavity suction sidewall  76  angles inwardly at an acute angle φ away from a line  85  parallel to an outer surface  84  of the suction sidewall  20 , such that the thickness of the side walls  18 ,  20  increases along the cavity  66   a  in the direction of cooling fluid flow. The angle θ may be equal to the angle φ, or the angles θ and φ may comprise different acute angles. The converging cavity sidewalls  74 ,  76  increase the impingement angle of the cooling air jet passing through the orifices  68  relative to the sidewalls  74 ,  76  to increase the cooling effect on the pressure and suction sidewalls  18 ,  20  in the area of the trailing edge  24 . Each of the second and third cooling cavities  66   b ,  66   c  may be formed with angled sidewalls  74 ,  76 , similar to the angled sidewalls  74 ,  76  described for the first cooling cavity  66   a , angling inwardly from the respective pressure and suction sidewall surfaces  82 ,  84 . The convergent angles θ and φ are preferably in the range of approximately 10 to 30 degrees. 
         [0026]    Further, it may be noted that the outer surfaces  82 ,  84  of the pressure and suction sidewalls  18 ,  20  are preferably formed as substantially straight planar surfaces, extending the in the span-wise direction, in the area of the trailing edge  24 . Specifically, the airfoil  12  may be formed with at least the trailing edge  24  formed as a substantially straight edge. For example, the airfoil  12  incorporating the cooling configuration of the present invention may be formed in accordance with the external airfoil profile disclosed in co-pending U.S. Application Serial No. (attorney docket no. 2006P23679US), which application is incorporated herein by reference. 
         [0027]    The orifices  68  and trailing edge openings  38  are preferably formed as drilled holes, in contrast to orifices or openings formed by typical casting processes. The drilled holes permit a smaller orifice  68  and opening  38  to be formed than may be provided by casting. For example, the diameter of the drilled orifices  68  and openings  38  is preferably in the range of 0.8 mm to 1.0 mm, whereas due to the fragile nature of the ceramic core required for the casting process, it is typically necessary to form cast holes with a diameter on the order of 1.5 mm to 2.0 mm to avoid breakage of the delicate ceramic core material during manufacture of the airfoil. 
         [0028]    As illustrated in  FIGS. 3 and 5 , the orifices  68  in each of the successive ribs  64   a ,  64   b ,  64   c  and respective openings  38  are aligned or centered on a common centerline  86 . Accordingly, each series of orifices  68  in the impingement ribs  64   a ,  64   b ,  64   c  and the associated trailing edge opening  38  aligned along a common centerline  86  may be formed by passage of a drill, during a drilling operation, into a specified location at the trailing edge  24  of the airfoil  12 . The provision of drilled holes permits control of the flow rate through the trailing edge cavities  66  without the previous constraints associated casting geometry requirements, allowing the present configuration to achieve a lower cooling fluid flow rate as the cooling fluid travels toward the trailing edge openings  38 , and permitting optimization of the cooling fluid flow rate by allowing variation of the drilled hole size. Further, the drilled holes increase the design flexibility in that the particular span-wise locations, as well as number, of the orifices  68  and openings  38  may be determined and/or changed to obtain a desired temperature profile for the airfoil  12 . 
         [0029]    During operation of the turbine, cooling fluid, such as cooling air, passes into the second elongated cooling chamber  46  through the cooling fluid inlet openings  40   c  and  40   d , and passes through the orifices  68  in the first rib  64   a  and is expanded to impinge on the convergent walls  74 ,  76  in the first cooling chamber  66   a . The cooling fluid is then contracted through the orifices  68  in the second rib  64   b  and is expanded to impinge on the convergent walls  74 ,  76  in the second cooling chamber  66   b . The cooling fluid is then contracted through the orifices  68  in the third rib  64   c  and is expanded to impinge on the convergent walls  74 ,  76  in the third cooling chamber  66   c . Finally, the cooling fluid is contracted through the trailing edge openings  38  and discharged from the airfoil  12  at the trailing edge  24 . 
         [0030]    From the above description, it may be seen that the multiple impingement cavity design provided at the trailing edge  24  increases the cooling effectiveness in the area of the trailing edge  24 . Also, in contrast to known designs incorporating cavity sidewalls that are parallel to the sides of the airfoil, the present invention increases the convective heat transfer within the trailing edge cavities  66  by providing converging cavity sidewalls  74 ,  76  that are angled inwardly relative to the adjacent surfaces  82 ,  84  of the airfoil outer wall  16 , such that the angle of impingement of air passing through each orifice  68  is increased. As a result of multiple impingements onto the successive convergent walls  74 ,  76  in the cavities  66 , a higher rate of heat transfer is provided in the trailing edge area of the airfoil  12 . 
         [0031]    While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.