Abstract:
A turbine airfoil apparatus includes: an airfoil including a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and a trailing edge; an endwall projecting laterally outwardly from the airfoil at one spanwise end thereof, the endwall having an outer surface facing the airfoil and an opposing inner surface; a plenum defined within the endwall between the inner and outer surfaces wherein the plenum is forked in plan view, with at least two branches, each branch having a throat disposed at its upstream end; and at least one film cooling hole passing through the outer surface and communicating with the plenum.

Description:
BACKGROUND OF THE INVENTION 
       [0001]    The present invention relates to gas turbine engines and, more particularly, to methods and apparatus for cooling endwalls of turbine airfoils. 
         [0002]    In a gas turbine engine, hot gas exits a combustor and is utilized by a turbine for conversion to mechanical energy. This mechanical energy drives an upstream high pressure compressor. The turbine comprises a plurality of rows of blades which are carried by a turbine rotor, alternating with rows of stationary nozzles. The turbine blades and nozzles are subjected to a flow of the corrosive, high-temperature combustion gases. These “hot section” components are typically cooled by a flow of relatively low-temperature coolant, such as air extracted (bled) from the compressor. 
         [0003]    As turbine inlet temperatures in modern gas turbine engines continue to rise, the endwalls of the hot section components (i.e. turbine blade platforms and nozzle bands) become more difficult to cool with traditional techniques. In addition, advanced aerodynamic features such as endwall contouring put extra pressure on maintaining acceptable material temperatures. 
         [0004]    The current state of the art is to drill film holes through the endwalls, to be fed by cooling air beneath the component. As a result, holes can only be placed in certain regions where they can be completely drilled to the other side or where the gas path pressure is low enough since the cooling air pressure feeding these holes is much lower than the airfoil cooling air. 
         [0005]    Some designs use hollow platforms that feed compressor bleed air to film cooling holes, but these designs are generally not adaptable to providing different cooling hole patterns based on varying operating conditions. 
         [0006]    Accordingly, there is a need for a turbine airfoil platform with improved cooling. 
       BRIEF DESCRIPTION OF THE INVENTION 
       [0007]    This need is addressed by the present invention, which provides a turbine airfoil having a cooling circuit cast therein. The cooling circuit can include various patterns of cooling holes. 
         [0008]    According to one aspect of the invention, a turbine airfoil apparatus includes: an airfoil including a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge; an endwall that projects laterally outwardly from the airfoil at one spanwise end thereof, the endwall having an outer surface facing the airfoil and an opposing inner surface; a plenum defined within the endwall between the inner and outer surfaces wherein the plenum is forked in plan view, with at least two branches, each branch having a throat disposed at its upstream end; and at least one film cooling hole passing through the outer surface and communicating with the plenum. 
         [0009]    According to another aspect of the invention, a is provided method of making a cooling hole pattern in a turbine airfoil apparatus that includes: an airfoil including a concave pressure sidewall and a convex suction sidewall joined together at a leading edge and at a trailing edge; an endwall that projects laterally outwardly from the airfoil at one spanwise end thereof, the endwall having an outer surface facing the airfoil and an opposing inner surface; and a plenum defined within the endwall between the inner and outer surfaces wherein the plenum is forked in plan view, with at least two branches, each branch having a throat disposed at its upstream end; the method comprising machining through the outer surface so as to define at least one film cooling hole communicating with the plenum. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0010]    The invention may be best understood by reference to the following description taken in conjunction with the accompanying drawing figures in which: 
           [0011]      FIG. 1  is a schematic perspective view of a turbine blade constructed in accordance with an aspect of the present invention; 
           [0012]      FIG. 2  is a view taken along lines  2 - 2  of  FIG. 1 ; 
           [0013]      FIG. 3  is a partially cut-away view of the turbine blade shown in  FIG. 2 ; 
           [0014]      FIG. 4  is a schematic perspective view of a turbine nozzle constructed in accordance with an aspect of the present invention; and 
           [0015]      FIG. 5  is a view taken along lines  5 - 5  of  FIG. 4 . 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0016]    Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views,  FIG. 1  illustrates an exemplary turbine blade  10 . The turbine blade  10  includes a conventional dovetail  12 , which may have any suitable form including tangs that engage complementary tangs of a dovetail slot in a rotor disk (not shown) for radially retaining the blade  10  to a disk as it rotates during operation. A blade shank  14  extends radially upwardly from the dovetail  12  and terminates in a platform  16  that projects laterally outwardly from and surrounds the shank  14 . The platform  16  may be considered a species of “endwall.” A hollow airfoil  18  extends radially outwardly from the platform  16  and into the hot gas stream. The airfoil  18  has a concave pressure sidewall  20  and a convex suction sidewall  22  joined together at a leading edge  24  and at a trailing edge  26 . The airfoil  18  extends from a root  28  to a tip  30 , and may take any configuration suitable for extracting energy from the hot gas stream and causing rotation of the rotor disk. The pressure sidewall  20  and the suction sidewall  22  extend radially outward beyond a tip cap  32  to define a structure generally referred to as a “squealer tip.” 
         [0017]    The blade  10  may be formed as a one-piece casting of a suitable superalloy, such as a nickel-based superalloy, which has acceptable strength at the elevated temperatures of operation in a gas turbine engine. At least a portion of the airfoil  18  may be coated with a protective coating of a known type, such as an environmentally resistant coating, or a thermal barrier coating, or both. 
         [0018]    The interior of the airfoil  12  is hollow and may include any one of a number of known cooling configurations including, for example, parallel radial or serpentine flow channels with various structures such as turbulators formed therein for improving cooling air effectiveness. The spent cooling air from the airfoil interior may be discharged through film cooling holes  34  and trailing edge discharge holes  36 . The cooling air is fed to the airfoil  18  through one or more feed channels  38  extending through the dovetail  12  and shank  14  into the airfoil  18 . 
         [0019]    The platform  16  includes an inner surface  40  and an outer surface  42 . A plenum  44  (see  FIGS. 2 and 3 ) is formed unitarily within the platform  16 . The periphery of the plenum  44  is defined and bounded by the inner and outer surfaces  40  and  42 , and by internal walls spanning the gap between the inner and outer walls  40  and  42 . The plenum  44  is formed as a part of the blade  10  using a known casting process. 
         [0020]    The plenum  44  includes, in sequence in a generally axial direction from front to rear, a first region  1 , a second region  2 , and a third region  3 . The cross-sectional area of the plenum  44  generally increases from front to rear. A fourth region  4  is disposed in flow communication with the first region  1 . A fifth region  5  is disposed in flow communication with the fourth region  4  and is disposed axially forward of the third region  3 . The overall shape of the plenum may be described as “forked” or “branched” in plan view, with the second and third regions  2  and  3  defining one branch and the fourth and fifth regions  4  and  5  defining a second branch. As will be explained in more detail below, each branch of the plenum  44  includes a throat- or nozzle-type structure at its upstream end. 
         [0021]    During engine operation, cooling air enters the dovetail  12  through the feed channel  38 . The first region  1  of the plenum  44  is fed cooling air by the feed channel  38 . Cooling air then flows from the first region  1  into the connected second region  2 . The second region  2  is the main region where convective cooling of the platform  16  takes place. The second region  2  has a relatively constricted flow area, seen as a reduced width or lateral dimension in  FIGS. 2 and 3 . This functions as a throat or nozzle to increase flow velocity and thereby enhance the heat transfer to the external surface of the platform  16 . The location (i.e. its position in the axial and tangential directions) of the second region  2  may be selected to correspond with the location on the platform  16  expected to experience the highest temperatures during engine operation. This may be determined by analysis or by testing. After being used for convective cooling in the second region  2 , the cooling air flows to the third region  3 . The third region  3  may be provided with internal heat transfer enhancement features such as ribs, fins, pins, or the like. In the illustrated example it includes a plurality of spaced-apart turbulence promoters or “turbulators”  46 . The cooling air exits the third region  3  through a plurality of film cooling holes  48  (best seen in  FIG. 2 ). The number, size, and location of the film cooling holes  48  is selected to discharge a protective film of cooling air over a portion of the platform  16 . As used herein, the term “film cooling hole” refers to a hole which is sized to discharge a film of cooling air over a surface, so as to protect the surface from high-temperature flowpath gases. While the exact dimensions will vary with the specific design, those skilled in the art will recognize a distinction between a “film cooling hole” and other types of holes, such as “impingement cooling holes” and “purge holes”. 
         [0022]    The film cooling holes  48  may be formed by known methods such as conventional drilling, laser drilling, or electrical discharge machining (ECM). These methods are referred to generically herein as “machining.” 
         [0023]    The flow path for cooling air from the first region  1  to the third region  3  extends in a direction generally parallel to a line between the leading edge  24  to the trailing edge  26 . 
         [0024]    The first region  1  also communicates with the fourth region  4 . Like the second region  2 , the fourth region  4  has a relatively constricted flow area, seen as a reduced width or lateral dimension in  FIGS. 2 and 3 . This functions as a throat or nozzle to increase flow velocity and thereby enhance the heat transfer to the external surface of the platform  16 . After being used for convective cooling in the fourth region  4 , the cooling air flows to the fifth region  5 . The fifth region  5  is generally rectangular in plan view and is positioned axially forward of the third region  3 . In operation, some cooling air from the first region  1  enters the fifth region  5 . One or more purge holes  50  may be provided in the fifth region  5 , exhausting into the secondary flowpath inboard of the platform  16  (through inner surface  40 ). The purge hole  50  permits a small amount of flow to exit the fifth region  5 , thereby preventing flow stagnation and build-up of debris in the fifth region  5 . The presence of the fourth region  4  reduces the weight of the blade  10 . Furthermore, the fourth region  4  provides a means by which the cooling configuration of the blade  10  can be revised and/or upgraded without changes to the basic casting. For example, the purge hole  50  could be eliminated by plugging it (e.g. using brazing or welding techniques), and one or more of film cooling holes  52  (see  FIG. 2 ) may be drilled through the surface of the platform  16  connecting to the fourth region  4 . 
         [0025]    The principles described above may be applied to other types of airfoil structures as well. For example,  FIGS. 4 and 5  illustrate an exemplary turbine nozzle  110 . The turbine nozzle  110  includes a pair of hollow airfoils  118  extending in a radial direction between an arcuate inner band  116  and an arcuate outer band  117 . Like the platform  16  described above, the inner and outer bands  116  and  117  may each be considered a species of “endwall.” Each airfoil  118  has a concave pressure sidewall  120  and an opposed convex suction sidewall  122  joined together at a leading edge  124  and at a trailing edge  126 . The airfoils  118  may take any configuration suitable for directly a hot gas stream to a downstream row of rotating turbine blades (not shown). The turbine nozzle  110  may be formed as a one-piece casting of a suitable superalloy, such as a nickel-based superalloy, which has acceptable strength at the elevated temperatures of operation in a gas turbine engine. At least a portion of the turbine nozzle  110  may be coated with a protective coating of a known type, such as an environmentally resistant coating, or a thermal barrier coating, or both. 
         [0026]    The interior of the airfoils  118  are hollow and may include any one of a number of known cooling configurations including, for example, parallel radial or serpentine flow channels with various structures such as turbulators formed therein for improving cooling air effectiveness. The spent cooling air from the airfoil interior may be discharged through film cooling holes  134  and trailing edge discharge openings  136 . The cooling air is fed to the airfoil  118  through one or more feed channels  38  extending through the inner band  116  into the airfoil  118 . 
         [0027]    The inner band  116  includes an inner surface  140  and an outer surface  142 . A plenum  144  (see  FIGS. 2 and 3 ) is formed unitarily within the inner band  116  (optionally, the outer band  117  could include a plenum). The periphery of the plenum  144  is defined and bounded by the inner and outer surfaces  140  and  142 , and by internal walls spanning the gap between the inner and outer surfaces  140  and  142 . The plenum  144  is formed as a part of the turbine nozzle  110  using a known casting process. 
         [0028]    The plenum  144  is similar in construction to the plenum  44  described above. It includes a first region  101 , a second region  102 , a third region  103 , a fourth region  104 , and a fifth region  105 . is The overall shape of the plenum  144  may be described as “forked” or “branched” in plan view, with the second and third regions  102  and  103  defining one branch and the fourth and fifth regions  104  and  105  defining a second branch. Each branch of the plenum  144  includes a throat- or nozzle-type structure at its upstream end. More specifically, the second region  102  and the fourth region  104  each has a relatively constricted flow area, seen as a reduced width or lateral dimension. This functions as a throat or nozzle to increase flow velocity and thereby enhance the heat transfer to the outer surface  142  of the inner band  116 . 
         [0029]    Cooling air exits the third region  103  through a plurality of film cooling holes  148 . The number, size, and location of the film cooling holes  148  is selected to discharge a protective film of cooling air over a portion of the inner band  116 . One or more purge holes  150  may be provided in the fifth region  105 , exhausting into the secondary flowpath inboard of the inner band  116 . The purge hole  150  permits a small amount of flow to exit the fifth region  105 , thereby preventing flow stagnation and build-up of debris in the fifth region  105 . 
         [0030]    Furthermore, the fifth region  105  provides a means by which the cooling configuration of the nozzle  110  can be revised and/or upgraded without changes to the basic casting. For example, the purge hole  150  could be eliminated by plugging it (e.g. using brazing or welding techniques), and one or more of film cooling holes  152  may be drilled through the surface of the inner band  116 , connecting to the fifth region  105 . 
         [0031]    The cooling configuration described above eliminates the cooling restrictions in prior art hot section gas components, namely the location, orientation, and quantity of film cooling holes. With those restrictions removed, holes can be placed anywhere on the endwall, since a majority of it is now hollow and contains higher coolant pressure to ensure positive cooling flow. This design provides lower temperature air and increased flexibility in cooling design. 
         [0032]    This design also provides the possibility of altering a component&#39;s cooling design without having to change the casting. For example, the same basic casting used to manufacture the turbine blade  10  described above could be machined with different patterns of film cooling holes communicating with the plenum  44 , depending on the specific end use, design intent, and analytical techniques available at the time the blade is designed and manufactured. 
         [0033]    The foregoing has described a turbine airfoil for a gas turbine engine. While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be made without departing from the spirit and scope of the invention. Accordingly, the foregoing description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of illustration only and not for the purpose of limitation.